Airfoils at Low Speeds

AIRFOILS AT LOW SPEEDSAirfoils at Low Speeds Copyright © 1989 by Selig, Donovan, and Fraser All rights reserved H. A. Stokely, publisher 1504 North Horseshoe Circle Virginia Beach, Virginia 23451 USA Airfoils at Low Speeds FOREWORD AND ACKNOWLEDGEMENTS The history of this experimental program on low-speed airfoils is extensive. In August 1986, work toward testing model sailplane airfoils in a wind tunnel at Princeton University began on an ambitious scale. The initial plan was to test 30 airfoils: 15 existing airfoils and 15 new airfoils to be designed concurrently with the tests. As news of the project caught the attention of radio control (RC) model soaring enthusiasts, the project grew far beyond the original goals and expectations, thanks to their generosity. When the experimental apparatus was finally dismantled in January 1989, almost two and a half years later, over 60 models were tested and over 130 airfoil polars were generated. It is our hope that the results of this work will be valuable to modelers and researchers for many years to come. A word is in order to explain the role each of us played in this effort. The initial impetus for the project, its organization and day-to-day management, as well as the wind tunnel testing and data reduction were done by Selig and Donovan. They also designed all of the new airfoils except for the DF -series by Fraser. The custom measurement apparatus was built jointly by Selig and Donovan at Princet.on University and by Fraser at Fraser-Volpe Corporation. The digitizing of the models was done at Fraser-Volpe Corporation by Fraser, who also wrote the computer programs for reducing this part of the data. All three of us shared in the writing and editing of this book. We would also like to mention that everything from the data collection to the writing of this book was done by computer. There is not a single number anywhere in any part of the data thiJ.t was generated, computed, reduced, copied, averaged, printed, graphed, or manipulated by hand. Aside from the speed and convenience of this approach, the principal advantage is the complete elimination of several types of errors that may otherwise occur. All of the airfoil polar data is available on 3~ or inch IBM compatible diskette from Fraser. Sincere thanks go to Prof. Smits of the Princeton University Gas Dynamics Laboratory for his enduring support while this extracurricular project began to grow and consume seemingly endless hours of time away from the first two authors' regular thesis research. The gracious support of Prof. Lam and Prof. Curtiss of Princeton University, and the helpful discussions of Prof. Maughmer of The Pennsylvania State University are appreciated. Thanks also go to Lou Pizzarello who provided us with an air conditioner and new air intake filters for the tunnel. We are indebted to Ray Olsen for his many contributions at times when we needed them most. 5i v vi Airfoils at Low Speeds For monetary contributions which made possible the purchase of important elements essential to this work we thank Rolf Girsberger, H.A. Stokely, Jerry Jackson, Armin Saxer, Charles Griswald, C. Haverlan, H.J. Rogers, Preben Norholm, Brian Smith, Thomas Yamokoski, and Trey Wood. The expertise of many skilled model builders made the lengthy set-up stage all worthwhile. In this respect we very deeply appreciate the work of Bob Champine, Ron Wagner, Stan Watson, Mark Allen, Michael Bame, Tony Beck, Woody Blanchard, Charles Fox, Peter Illick, Harley Michaelis, Forrest Miller, Ted Off, Mike Reed, Tyson Sawyer, Chuck Anderson, Norman Anderson, Jerry Arana, Bruce Baker, Ken Bates, David Batey Jr., Rich Border, John Boren, Mike Chiddick, Doug Dorton, Roger Egginton, Dale Folkening, Harlan Halsey, John Hohensee, Dave Jones, Stan Koch, Terry Luckenbach, Carl Mohs, Lee Murray, Mark Nankivil, R.J. Ostrander, Jef Raskin, Les Rogers, Joe Ruminski, and Karl Widiner. Prof. Mark Drela of M.I.T. is gratefully acknowledged for making his ISES computer code available to aid in the analysis of the new airfoil designs. Finally, MKS Instruments, Inc. and Scientific Solutions, Inc. are acknowledged for their valuable contributions of instrumentation. Airfoils at Low Speeds TABLE OF CONTENTS Foreword and Acknowledgements . List of Polars and Lift Plots List of Symbols and Abbreviations 1 Introduction . . . . . . . . 2 Experimental Facility and Measurement Technique 2.1 Flow Quality . . . . 2.2 Wind Tunnel Models 2.2.1 Digitized Profiles 2.2.2 Digitizing Procedure 2.2.3 Digitizer Res1,1lts 2.3 Force Measurement Technique and Instrumentation 2.4 Comparison with Other Facilities 3 Low Reynolds Number Terminology 3.1 Laminar Separation Bubbles 3.2 Trips and Bubble Ramps . . . 3.3 Airfoil Hysteresis . . . . . . 4 Project Design Methods and Goals 5 Comments on Airfoils 5 .1 Airfoil Discussions 5.2 Stall Behavior 5.3 Trips and Surface Roughness 5.4 Trailing Edge Thickness 5.5 Surface Waviness and Contour Accuracy 6 References . . . . 7 Airfoil Coordinates 7.1 Nominal 7.2 Actual . . . . 8 Predicted Moment Data 9 Airfoil Thickness and Camber 10 Digitizer Plots . . . . . 11 Airfoil Comparison Plots 12 Airfoil Polars and Lift Plots 13 Tabulated Polar Data 14 Addresses . . . . . . . . . v viii xii 1 5 5 6 7 7 9 11 16 41 41 42 42 45 51 53 88 89 90 90 101 103 103 129 147 149 153 187 193 359 397 vii viii Airfoils at Low Speeds LIST OF POLARS AND LIFT PLOTS Figure 12.1 AQUILA-PT 12.2 AQUILA-PT lift data 12.3 CLARK-Y-PT ·. . . 12.4 CLARK-Y-PT lift data 12.5 DAE51-PT 12.6 DAE51-PT thickened trailing edge 12.7 DAE51-PT lift data 12.8 DF101-PT 12.9 DF102-PT 12.10 DF103-PT 12.11 E193-PT . 12.12 E193MOD-PT 12.13 E205A-PT . . 12.14 E205B-PT . . 12.15 E205B-PT lift data 12.16 E214A-PT . . . 12.17 E214B-PT . . . 12.18 E214C-PT 3° flap 12.19 E214C-PT 0° flap 12.20 E214C-PT -3° flap 12.21 E214C-PT -6° flap 12.22 E214C-PT u.s.t. xjc = 20%,hjc = .17%,w/c = 1.0% 12.23 E214C-PT lift data 12.24 E374A-PT . . . . . . . . . . . . . . 12.25 E374B-PT . . . . . . . . . . . . . . 12.26 E374B-PT u.s. bumps xjc =50%, type A 12.27 E374B-PT u.s.t. xjc = 20%,hjc = .17%,w/c = 1.0% 12.28 E374B-PT u.s. wavy clay, xjc = O% to 15%,h/c = .20% 12.29 E374B-PT thickened trailing edge 12.30 E374B-PT lift data 12.31 E387 A-PT . . . . . . . . . . 12.32 E387 A-PT repeated . . . . . . 12.33 E387 A-PT u.s.t. xjc = 20%, hjc = 0.17%, wjc = 1.0% 12.34 E387 A-PT high turbulence 12.35 E387 A-PT lift data 12.36 E387B-PT . . . . . 12.37 Flat Plate-PT 12.38 Flat Plate-PT lift data 12.39 FX60-l00-PT . . . . 193 194 195 196 197 198 199 200 201 202 203 204 205 206 207 208 209 210 211 212 213 214 215 216 217 218 219 220 221 222 223 224 225 226 227 228 229 230 231 Airfoils at Low Speeds 12.40 FX63-137 A-PT 12.41 FX63-137B-PT 12.42 HQ2/9A-PT 12.43 HQ2/9A-PT u.s.t. xjc = 20%,hjc = .17%,wjc = 1.0% 12.44 HQ2/9A-PT u.s.t. xjc = 40%, h/c = .17%,wjc = 1.0% 12.45 HQ2/9A-PT l.s.t. xjc = 50%,h/c = .17%,wjc = 1.0% 12.46 HQ2/9A-PT lift data . . . . . . . . . . . . . . 12.47 HQ2/9B-PT . . . . . . . . . . . . . . . . . . 12.48 HQ2/9B-PT u.s.t. xjc = 50%,hjc = .17%,wjc = 1.0% 12.49 HQ2/9B-PT u.s. blowing xjc =50%, type B 12.50 HQ2/9B-PT trips, Rn = 200,000 12.51 HQ2/9B-PT lift data 12.52 J5012-PT . . . . . . . . . . 12.53 MB253515-PT . . . . . . . . 12.54 MB253515-PT u.s.t. xjc = 20%,hjc = .17%,w/c = 1.0% 12.55 MB253515-PT lift data . . . . . . . . . . 12.56 M0&-13-128-PT . . . . . . . . . . . . . . . . . . 12.57 M0&-13-128-PT u.s. bumps xjc = 31%, type A . . . . 12.58 M0&-13-128-PT misc. u.s. trips, xjc = 31%,Rn = 200,000 12.59 M0&-13-128-PT lift data 12.60 NACA 0009-PT . . . . 12.61 NACA 0009-PT lift data 12.62 NACA 2.5411-PT . . . 12.63 NACA 2.5411-PT lift data 12.64 NACA 64A010-PT 12.65 NACA 64A010-PT lift data 12.66 NACA 6409-PT . . . . 12.67 NACA 6409-PT lift data . 12.68 RG15-PT . . . . . . . 12.69 RG15-PT u.s.t. xjc = 20%,hjc = .17%,wjc = 1.0% 12.70 RG15-PT u.s.t. x/c = 40%,hjc = .17%,wjc = 1.0% 12.71 RG15-PT u.s.t. x/c = 60%,hjc = .17%,wjc = 1.0% 12.72 RG15-PT u.s.t. xjc = 70%,h/c = .17%,wjc = 1.0% 12.73 RG15-PT lift data 12.74 52048-PT . . . . . 12.75 52048-PT with trips . 12.76 52048-PT misc. trips 12.77 52048-PT lift data 12.78 52055-PT 12.79 52091A-PT . . . 12.80 52091B-PT . . . 12.81 52091B-PT Gurney Flap type A 232 233 234 235 236 237 238 239 240 241 242 243 244 245 246 247-249 250 251 252 253 254 255 256 257 258 259 260 261 262 263 264 265 266 267 268 269 270 271 272 273 274 275 ix x Airfoils at Low Speeds 12.82 82091B-PT Gurney Flap type B 12.83 82091B-PT Gurney Flap type C 12.84 82091B-PT lift data 12.85 83010-PT 12.86 83010-PT lift data 12.87 83014-PT 12.88 83014-PT lift data 12.89 83016-PT 12.90 83021A-PT . . . 12.91 83021A-PT lift data 12.92 83021B-PT 12.93 84061A-PT . . . . 12.94 84061B-PT . . . . 12.95 84061B-PT u.s.t. xfc = 45%,hfc = .17%,w/c = 1.0% 12.96 84061B-PT u.s.t. xfc = 45%, Rn = 150,000 . . . . 12.97 84061B-PT u.s.t. xfc = 45%, Rn = 150,000 and 300,000 12.98 84061B-PT lift data 12.99 84062-PT 12.100 84180-PT . . . . 12.101 84233-PT . . . . 12.102 84233-PT u.s.t. xfc = 20%,hfc = .17%,wfc = 1.0% 12.103 84233-PT lift data . 12.104 8D2030-PT . . . . 12.105 8D2030-PT lift data 12.106 8D2083-PT . . . . 12.107 8D5060-PT •' . 12.108 8D5060-PT lift data 12.109 8D6060-PT . . . . 12.110 8D6060-PT u.s.t. xfc = 20%,h/c = .17%,w/c = 1.0% 12.111 SD6060-PT u.s.t. xjc = 40%,hjc = .17%,w/c = 1.0% 12.112 SD6060-PT lift data . . . . . . . . . . . . . . 12.113 8D6080-PT . . . . . . . . . . . . . . . . . . 12.114 8D6080-PT u.s.t. xfc = 10%,hjc = .17%,wjc = 1.0% 12.115 8D6080-PT u.s.t. xfc = 20%,hjc = .17%,wjc = 1.0% 12.116 8D6080-PT u.s.t. xfc = 30%,hjc = .17%,w/c = 1.0% 12.117 8D6080-PT thickened trailing edge 12.118 8D7003-PT . . . . . . . . . . . . . . . . . . 12.119 8D7003-PT repeated . . . . . . . . . . . . . . 12.120 8D7003-PT u.s.t. xfc = 60%,hfc = .17%,wfc = 1.0% 12.121 8D7003-PT u.s.t. xfc = 70%,hjc = .17%,wfc = 1.0% 12.122 8D7003-PT u.s. bumps xfc =50%, type A 12.123 SD7003-PT u.s. bumps xfc = 60%, type A . . . . 276 277 278 279 280 281 282 283 284 285 286 287 288 289 290 291 292 293 294 295 296 297 298 299 300 301 302 303 304 305 306 307 308 309 310 311 312 313 314 315 316 317 .hjc = . . bumps xjc = 70%. . xjc = 45%.Airfoils at Low Speeds 12. .144 SD7080-PT . Rn = 300.130 SD7032C-PT 0° flap 12. 12.0% 12.hjc = .152 SD8000-PT .127 SD7032B-PT 12. . type A 12.0% 12. . .s. .0% 12. . .hjc = . . . .146 SD7084-PT . .139 SD7043-PT u. . . . .17%.t.160 SPICA-PT 12. . . .138 SD7043-PT .hjc = .17%. . .128 SD7032C-PT 6° flap 12. . . . . . . .135 SD7032D-PT u. . .147 SD7084-PT lift data 12. . . . .t. .0% 12. 12.wjc = 1.17%.s. . . .000 12. .141 SD7062-PT . xjc = 15%.17%.150 SD7090-PT trips.t.137 SD7037-PT u. . .126 SD7032A-PT 12. 12. . .156 SD8000-PT lift data 12. . . 12. 12. xjc = 20%. xjc = 30%. xj c = 15%. . 12.131 SD7032CcPT -3° flap 12. .t. . . . .142 SD7062-PT u. . .17%.154 SD8000-PT u. .136 SD7037-PT .158 SD8020-PT lift data 12. .s. .s. . .w/c = 1. .0% 12. 12.17%. .w/c = 1. . .w/c = 1. xjc = 20%. 318 319 320 321 322 323 324 325 326 327 328 329 330 331 332 333 334 335 336 337 338 339 340 341 342 343 344 345 346 347 348 349 350 351 352 353 354 355 356 357 358 xi . .155 SD8000-PT u.0% 12. .157 SD8020-PT . .159 SD8040-PT 12.0% 12. h/ c = .163 WB135/35-PT lift data 12. .w/c = 1. .132 SD7032C-PT -6° flap 12.hjc = . .164 WB140/35/FB-PT .148 SD7090-PT .000 12. .129 SD7032C-PT 3° flap 12.153 SD8000-PT u.hjc = . . . . .s.125 SD7003-PT lift data 12. . .s. . w / c = 1.161 SPICA-PT lift data 12. .0% 12. . 12. . . .wjc = 1. . 12. .162 WB135/35-PT . 12. . 12.08%.t.s. . . .149 SD7090-PT loose/tight covering.w/c = 1.s. . . 12. . .151 SD7090-PT lift data . . xjc = 40%. . . .145 SD7080-PT lift data 12.t.140 SD7043-PT lift data .134 SD7032D-PT .s.17%. .t.133 SD7032C-PT lift data 12.17%. xjc = 70%.hjc = . . . . . . Rn = 300.124 SD7003-PT u.143 SD7062-PT u.t. . . . . . /4. pV00 cjp streamwise velocity in wake root-mean-square of the streamwise fluctuating velocity freestream velocity inviscid local velocity on airfoil surface trip width distance along airfoil chord. d h I L/D Paoo i::>.s.s. l.s. l. u. or horizontal distance vertical distance angle of attack fluid viscosity fluid density upper surface upper surface trip lower surface lower surface trip Princeton Tests .t. PT airfoil chord lift coefficient. dj~pV:foc pitching moment about the quarter-chord point drag per unit span trip height lift per unit span sailplane lift-to-drag ratio freestream dynamic pressure total pressure difference between freestream and wake Reynolds number.Po Rn u w X y a 1-' p u.c maximum lift coefficient drag coefficient.t.xii Airfoils at Low Speeds List of Symbols and Abbreviations c c1 elm a% Ca Cmc.s. I/~ pV:fc. In order to establish baseline data." Only then does the flow reattach to the airfoil. if the laminar separation point is sufficiently far aft or if the Reynolds number is very low. is generally well behaved. Instead the laminar boundary layer separates.000. 1 . say. And sometimes. and turbulent reattachment enclose a region of recirculating flow aptly called the "laminar separation bubble. it physically detaches from the airfoil surface. the flow entirely fails to return to the airfoil surface. Efforts towards drag reduction. This laminar separation. than 1-3 million-is well known. The boundary layer is laminar from the leading edge to a point typically near mid-chord where it makes a transition to turbulent flow. The flow then becomes unstable while separated. solving all the myriad problems that occur in a multi-year project of this SIZe. In either case large energy losses are associated with this process. as well as some of the older NACA sections and their close derivatives. They ranged from very simple. Unlike full-size airplanes. largely concentrate on reducing the size and extent of the bubble. At these low Reynolds numbers the flow is fundamentally different and more complicated than at high Reynolds numbers.000 and 500." It is this extended transition process that is the principal reason for the degradation in performance at low Reynolds numbers. as well as the flow behind it. These were selected primarily by the modeling community and are representative of what is presently used. this involved numerous preliminary steps from preparing and instrumenting the tunnel to arranging for the models to be built. model sailplanes typically operate at chord Reynolds numbers between 50. therefore. flat-bottom types. Aside from providing the baseline.Chapter 1: Introduction Chapter 1 Introduction The primary goal of this work was to design a new group of high-performance airfoils for radio controlled model sailplanes. The transition process is neither abrupt nor does it usually take place while the boundary layer is attached to the airfoil. and makes the transition to turbulent flow in "mid-air. establishing a design procedure and. transition to turbulence. of course. This transition. often called the low Reynolds number regime. a number of existing airfoil designs were tested first. that is. As can be imagined. testing these airfoils allowed us to compare our data with other facilities where the same sections had also been tested. The flow behavior over an airfoil at high Reynolds numbers-greater. to very modern FAI-contest airfoils. This book discusses a major experimental program that was carried out to do just that. Yet despite the high level of interest in this area. . computational efforts have not been entirely reliable in predicting this complex flow. and these are indicated by an asterisk (*). there have been numerous theoretical and experimental investigations which have resulted in three major conferences in the past four years 1 •2 •3 and a special AGARD publication 4 • Because of the difficulty in mathematically modeling the bubble. As far as we know. a great variety of airfoils was tested spanning virtually the entire range of usefulness to RC sailplanes. low-turbulence. In most studies and experiments.2 Airfoils at Low Speeds In an effort to understand the flow phenomena at low Reynolds numbers in general (which includes RC sailplanes). Nonetheless. we were able to design new airfoils.5411 NACA 64A010 NACA 6409 RG15 S2048 S2055 S2091* S3010 S3014 S3016 S3021* S4061* S4062 S4180 S4233 SD2030 SD2083 SD5060 SD6060 SD6080 SD7003 SD7032* SD7037 SD7043 SD7062 SD7080 SD7084 SD7090 SD8000 SD8020 SD8040 SPICA WB135/35 WB140/35/FB In order to ensure the enthusiasm of the modeling community which built all of the test sections. have models of them built. and test them in the tunnel. thereby "closing" the design loop. some airfoils previously unknown to us offered insights into the airfoil design process. As a result of this policy. several airfoils were duplicated in order to examine the effects of model variability. the primary emphasis has been on understanding the fundamental mechanisms that drive the bubble. this last step has not been done before on this scale for model aircraft applications.and SD-airfoils are the new designs resulting from this work. few systematic attempts have been made to apply the growing body of knowledge to the problem of airfoil design. 3 x 4 ft smoke tunnel at Princeton University. AQUILA CLARK-Y DAE51 DF101 DF102 DF103 E193 E193MOD E205* E214* E374* E387* Flat Plate FX60-100 FX63-137 HQ2/9 J5012 M06-13-128 MB253515 NACA 0009 NACA 2. three things happened. Second. And third. The following is a list of the 54 different airfoils. and to do it specifically for RC sailplane airfoils. All the models were tested in the low-speed. The DF. considerable progress has been made. we tested any airfoil that a builder wanted to supply. First. (2) the results of the tests on over 60 wind tunnel models. To help the non-specialist.Chapter 1: Introduction This book has two major parts: (1) the documentation of the facility and the quality of the data. and other supporting data are given in Chapters 7 through 13. The second major part begins with Chapter 5. tabulated data. Airfoil polars and lift plots. This part is intended primarily for those active in low Reynolds number airfoil research and may be skipped without loss of continuity by those more interested in the data. a more comprehensive report will come at a later time. the addresses are listed in Chapter 14. Chapter 4 briefly covers some of the ideas behind the new airfoil designs. The first part is covered in Chapter 2 and extensively documents the experimental methods we used in this work. 3 . which discusses the airfoil polars at length. Hopefully. for those who may wish to contact the authors or obtain additional copies of this book. coordinates. And finally. the terminology we use is explained in Chapter 3. 4 Airfoils at Low Speeds . The flow straighteners are 3 in square and 12 in long. The -3 dB point of the response curve was 33 kHz for chord Reynolds numbers of lOOk. At all conditions the wire was operated at an overheat of 0. It consists of an inlet and stilling chamber 9 ft high by 12 ft wide containing screens and flow straighteners. and 300k. 2. 2. In each case the lowpass frequency of the filter was set to somewhat less than the Nyquist sampling frequency to eliminate aliasing errors. By sampling over a range of frequencies. In this work.8. The tunnel speed in the test section is variable from 5 to 70 ftjs. the DC component (the mean) of the anemometer signal was subtracted ("bucked off") using an operational amplifier of an analog computer. All spectra presented here were found using a sampling frequency of 100 Hz with the hot wire located 3 in below the center of the tunnel. high resolution of the spectra was obtained. these frequencies are well above the energy-containing frequencies of the turbulence. no high-pass filter was used. The remaining signal was then amplified to fill the ±10 volt range of the 14-bit analog-to-digital converter and sampled at frequencies from 10 Hz to 10 kHz. the anemometer signal is high-pass filtered. As will be shown shortly. A sampling frequency of 100 Hz resolved the high-frequency end of the spectrum and extended down to sufficiently low frequencies.Chapter 2: Experimental Facility and Measurement Technique Chapter 2 Experimental Facility and Measurement Technique The tests were performed in the Princeton University 3 x 4 ft smoke tunnel. The frequency response was optimized using the standard test in which a square wave in voltage is injected at the Wheatstone bridge to simulate an impulse in velocity.1 Flow Quality Constant temperature hot-wire anemometry 5 (using Dantec model 55M01) was used to determine turbulence levels in the freestream. Instead. Downstream of the test section the flow is turned 90° and exits through a 50 HP fan. A common problem when measuring turbulence levels in low-speed facilities is determining the lowest frequency of interest. and 25 kHz for a Reynolds number of 60k. 200k. This location was 5 . Usually. A sketch of the tunnel is shown in Fig.1 and also on the back cover. however. This procedure reduces the apparent RMS turbulence level by removing low-frequency fluctuations which may be important to boundary layer transition. This section is followed by a 9:1 contraction leading to the test section which is 3 ft high by 4 ft wide. .064 200k 0. While large forces are desirable. to fiberglass-covered foam. this contribution to the turbulence would be lost.563 0. These surveys indicate a 2% variation in the velocity which was deemed sufficiently uniform.2 Wind Tunnel Models In selecting the model size to obtain the desired Reynolds number. For this work. i . several tradeoffs were considered. 9216 points were taken and then broken into ensembles of 1024 to calculate spectra. To achieve a given test Reynolds number. The model span was 33 in. the area under the curve is directly proportional to (u~m. For these reasons a 12 in chord was selected as a compromise between the two competing effects.v. The unfiltered turbulence levels at various Reynolds numbers are given below. Note that these levels correspond to a very low cut-off frequency of 0.170 0. All models were fully-sheeted except one. the measured forces increase with decreasing chord. As can be clearly seen. Construction techniques ranged from all-balsa with ribs.050 lOOk 0. spars and open bays. 2. 2. Perhaps fluctuations at frequencies this low have quasi-steady effects.358 0. Less than 4% variation was found in the static pressure and there was no measurable total pressure variation. In this way. in any event it is currently unclear what cut-off frequency should be used so both numbers are presented. For each run. a model shop was not used. Spectra at several different Reynolds numbers are shown in Figs. RMS Turbulence Intensities Rn 2:: 0.017 300k 0.J 2 .-the square of the turbulence intensity. models with small chords are difficult to build accurately. experienced model sailplane enthusiasts were solicited to build the models.01 Hz due to the sampling interval. Consequently. Power spectral density multiplied by frequency is plotted against the logarithm of frequency.6 Airfoils at Low Speeds representative of the turbulence characteristics throughout the central region of the test section.008 Mean-pressure surveys to determine the uniformity of the freestream were taken in the test section throughout a plane perpendicular to the flow. rather.01 Hz 2:: 1Hz 60k 0. the majority of the energy is found at frequencies below 1 Hz. If the signal were high-passed at 1 Hz.188 0. which had open-bay construction (see NACA 6409). Turbulence levels are also indicated below for the case of a cut-off frequency of 1 Hz.2 (a-d). construction tolerances were on the order of that found on model sailplanes. 2. every model was profiled using a digitizing Coordinate Measuring Machine (CMM) to obtain the actual airfoil shape. 2. many different kinds of deviations from standard shapes. angles. setting up this coordinate system. The computer then calculates the measuring diameter using basic geometry. and so on. It will refer all measurements either to the table itself or to an arbitrary coordinate system based on the part or its fixtures.Chapter 2: Experimental Facility and Measurement Technique 2. and "knows" what to do with the measurement once it is taken. This capability was used here. The procedure was as follows: Since measurements 7 .1 Digitized Profiles As a check for model accuracy and for later airfoil performance computations. A comparison was then made with the desired airfoil shape to determine the accuracy of the model. (This deflection is automatically accounted for by the CMM software.0003 in typical for all three axes. The machine is routinely calibrated to standards traceable to the N a tiona! Bureau of Standards.) Outputs from the three sensors are routed to a computer which runs commercial software allowing one to reduce the raw information from the three axes to determine diameters. Mounted on this gantry is a second gantry which in turn holds a vertical column to which the probe is attached. Profiling was performed at Fraser-Volpe Corporation in Warminster Pennsylvania using a Helme! Checkmaster CMM with full computer software for measurement processing. or reference frame as it is called. including fixturing and post-measurement data reduction. Coordinates are read from the sensors when the probe is deflected in any direction by approximately 0. 2. differential locations. The machine itself consists of a marble slab which is finished to a high surface flatness. The computer will then prompt the operator for what point to measure next. rotations. is a major part of any measurement.2. the computer can be "taught" a program of steps which represent a specific series of measurements.0003 in from its rest position. allowing the entire process of measuring a model to be completed in about 25 minutes.0005 in absolute and 0.2 Digitizing Procedure The complete procedure from selecting the airfoil model to the end of the data reduction was standardized. The probe consists of a hard plastic ball mounted to the end of a steel shaft which is screwed into a precision motion detector.3. On the slab is a gantry which can traverse the length of the table. This arrangement allows the probe to be positioned anywhere within a large volume beginning at the surface of the slab. lengths. A drawing of the CMM is shown in Fig. The equivalent measuring diameter of a probe is found (the probe is said to be "qualified") by touching it to a sphere of precisely known diameter. In addition to simply making a measurement. Indeed. This machine and software made it possible to determine the location of a point in space within 0. such as a flap joint or a long crack in the sheeting. however. over the central parts of the airfoil it was as great as ~ in. The leading edge of the airfoil was aligned with one axis of the CMM table so the measured chord direction was perpendicular to the span. the airfoil was examined to be sure that this location was representative of the total section. near the leading and trailing edges the spacing was small. This was detected by moving the probe vertically past it with the vernier lead screw at progressively closer settings until it touched. The final point touched in the upper surface sequence was the leading edge. blocks mounted to the fixtures were used to establish the reference coordinate system for the upper surface. larger distortions which covered most of the span. All the sections were of a constant chord so this point is not particularly important. the model and fixtures were turned over as a unit and the leading edge was once again aligned. any inadvertent movement during the data collection was detected by comparing the leading and trailing edge points as measured from both sides. After the upper surface was done. see Fig. Consequently. Because of the possibility of distorting the model. Small. These blocks could be touched by the probe with the fixture and model in either position (up or down). were included.ing it with the probe from directly behind. 2. so the actual position of the airfoil and fixtures on the machine was unimportant. it was smoothed. H they were at the same points with respect to the reference frame established by the blocks on the fixtures. Between 20 and 30 points were then touched on the upper surface. . the first point here duplicating the last point on the upper surface.) As mentioned above. machined reference blocks were permanently attached to the fixtures. H the covering had wrinkled. (Because of the relatively non-rigid construction of the models compared to metal parts. There will be more on this later. the stands could not be rigidly held down to the CMM table. and so the same chord would be measured on both sides of the model. Measuring continued at the leading edge. the fixtlfring was free-standing on the machine's table. The airfoil was held in two identical fixtures that supported the model in a level attitude on three points.4. local distortions were excluded from measuring. Consequently. These fixtures were designed so that when they were turned over the model was still supported on three points. when the section was turned over these blocks made it possible to maintain a single reference frame for both sides. The chordwise location of the trailing edge was then determined by touch. All points on the airfoil were referred by the software to the blocks. then no motion that could affect the measurement had occurred.8 Airfoils at Low Speeds were generally taken at only one spanwise location. but this method made it possible to check that nothing had moved during the measurements. This point was used to determine the actual chord of the section. Because of the impossibility of measuring both sides of the airfoil in a single position. The spacing of points was more or less proportional to the local curvature. a second program was used to reduce the CMM output to the actual coordinates of the airfoil. the two airfoils are fitted in a least squares sense. Camber error as well as thickening is frequently seen at the trailing edge. This difference is shown under the trailing edge of the upper plot. and to normalize the airfoil to coordinates between 0 and 1. 9 .6. since a chordwise error translates to a much smaller vertical error except at the leading edge. A header was added to identify the airfoil and show the actual chord. If both lines sweep up or down together then the camber is in error. However. or error between the two airfoils on a much expanded scale. The results of the measurements were collected on a Leading Edge personal computer. Of the 67 duplicated leading edge pairs.) In Chapters 10 and 11. All the output from the CMM was saved on a disk file.001 in.002. the plot would be two straight lines lying on the horizontal axis.0057 in. each of the actual sections is plotted against the nominal coordinates at half scale (6 vs 12 in).2.5. The lower plot shows the difference. 59 within 0. regardless of its position with respect to the axis. The fitting uses two variables-relative vertical location of the entire section and relative rotation of the entire sectionto produce the lowest possible RMS difference without distorting the airfoil.005 and all within 0.3 Digitizer Results (Note. 52 within 0. normally the test section. normally the prototype. The upper surface difference is the solid line and the lower surface the broken one.Chapter 2: Experimental Facility and Measurement Technique and the final point duplicating the first point on the upper surface (the trailing edge). Consequently. which was also used to reduce the data. On a 12 in chord this is trivial. If the solid line is above (or below) the broken one. respectively. 2. to rotate the actual chord so it was parallel to the reference axis. 46 were within 0. The CMM data are the locations of the center of the ball at the end of the probe. 2. the solid line is always the first airfoil. In a few cases different prototypes are compared. and the broken line the second. the section is too thick (or thin) at that point. If the two sections were perfectly matched. these accuracies imply a general thickness measurement error of under 0. 2. 65 within 0. A displacement above or below the axis means the test section surface lies above or below the nominal. The legend inside the airfoils shows what is being compared. because of the nature of the leading edge. This data was also saved as a file. Typical output of the CMM software is shown in Fig.003.001 in. Part of the difference within any pair is due to the impossibility of finding the exact point that was measured from the other orientation. Before they are plotted. not the surface itself. A typical output file from this program is shown in Fig. "-PT" (Princeton Tests) is appended to the airfoil name to distinguish the digitized coordinates from the nominal coordinates. the agreement was within 0. Because of the type of construction.the leading edge and the beginning of the sheeting. the vertical locations of the exact edges. Foam core sections usually had sharper trailing edges. rib and spar section built over a weekend for under $10 had one of the most accurate profiles measured. The most difficult point to measure is the vertical position of the leading edge. Several observations can be made about methods of construction based upon the models digitized in this study. the accuracy of some models costing many times this amount was only average.0007 in. a balsa-sheeted. The remaining four profiles were taken at 3 in intervals centered on the span and were intended to discover how much spanwise variation a good airfoil model might show.) This is because the slope becomes infinite and a very small change in the chordwise position of the probe produces inordinately large changes in the measured thickness. sheeted models tended to have a problem with the blend between . The SD7080 pair was done early in the profiling as a general check for repeatability.49. with the preponderance being up. the increased thickness at the leading edge carried back through a large part of the airfoil. As a check on the digitizing procedure.10 Airfoils at Low Speeds The short. once at the beginning of the profiling. but the spanwise stations were not the same. For example. have somewhat reduced accuracy.006% of chord. it is warped either up or down. On the other hand. except the joint at the leading edge was too wide. Built-up. the most common error is a poorly contoured trailing edge.46-10. 0. Even so. five times at the end (a time span of about 75 days). As can be seen from the plots. Consequently. because many points were digitized near the leading and trailing edges and because the contribution of each point to the overall accuracy number was weighted in proportion to the distance between it and the adjacent points. as shown by the position of the tics. However. Neither the cost nor the type of construction was a good indicator of the accuracy. two of the models were digitized more than once: the SD7080-PT and the SD7003-PT. 10. In addition. with built-up sections the errors were more local. the effect of the end points on the overall error is very small. the SD7003-PT and SD7003-PT (R) were taken at the same station and are a good indication of the overall repeatability of the measurement setup-about 0. (It is quite possible for a model to have more than one leading edge. The trailing edge also tended to be thick. This was a problem that was not present in models that used a single piece-usually wood-leading edge. inward-facing tics show the positions of the leading and trailing edges. As can be seen in Figs. The SD7003-PT was digitized six times. One model had excellent contours for the separately molded upper and lower surfaces. The open-bay models have no single profile-over the ribs it can be accurate.003 in. Trailing edges are a problem for all types of construction. Two profiles. it is very small indeed. the most forward and most rearward points themselves were not included in the error calculation. Since the . but inevitably there is sag between the ribs. but any errors in contour were more prolonged. 12). Rather than computing the drag based on just one vertical survey. and the drag was found indirectly using the momentum method 6 . This has significance when comparing polars. because small differences in performance on similar sections could be a result of the inaccuracy of the model or random variations in the test results rather than an indication of the superiority of one prototype section over another. These airfoils are. but the point is that one must be careful in claiming performance for the prototype based on the model's performance. 11 .3 Force Measurement Technique and Instrumentation Lift was measured directly using an electro-mechanical force balance. One model. where the model is smoother than the nominal. of course.) Since the point of these variants was to explore the effects of changes on the forward upper surface.Chapter 2: Experimental Facility and Measurement Technique sensitivity of performance to trailing edge location is high. The minor differences along the rest of the airfoil are due to the fact that the sample chords were not all at the same spanwise stations. and finally with a flap (SD7032C-PT). there simply was no representative section. In cases where the model is inaccurate. not to a nominal airfoil. quite similar. and because the fitting routine tends to distribute the deliberate "error" over the entire airfoil so as to keep the RMS error down. This is particularly noticeable on the upper surface of the WB135/35 between 1% and 3% chord. and 82048 are an example of this. One model. One section. was actually a better fit to the E205 than to its true nominal coordinates (see Figs. For a few airfoils (SPICA. and several plots compare these prototype sections. the S4180-PT. and therefore the spline routine that compares the airfoils can have residuals of the order of the errors. WB135/35.5 and 11. then with Monokote covering (SD7032B-PT). 10. a SD7032. (The DF101-PT is compared to the nominal. (This was the only model with such a major contour discrepancy.) Some of the nominal airfoils differ less between themselves than the models do with the ideal coordinates. was first tested in the tunnel with no covering over the sanded balsa sheeting (version A: SD7032A-PT). The HQZ/9. In these cases small errors in fit are not meaningful because hand-generated coordinates are not smooth in the mathematical sense. the relevant prototype is the DF101-PT. area. the performance applies to the model airfoil and not necessarily to the nominal airfoil. RG15. the E193-PT. The plots show what and how much was added or removed in that. Only the flapped version was digitized. had a very thin trailing edge which was so warped that it was meaningless to measure it at all. the wake was surveyed and the drag computed at four spanwise locations and then averaged. clearly there is a general problem here. and WB140/35/FB) the coordinates were supplied by the builders. The DF102-PT and DF103-PT are compared to the DF101-PT. 2. precision ball bearings were used throughout. The torque required to do this is directly related to the lift. the most difficult being to meet the requirement of low system friction.002 lb at the lowest Reynolds number.25 chord lengths downstream of the trailing edge to find . however.5 mm (0. To compute drag using the momentum method. Nine-point calibrations of the force balance were performed frequently to minimize the effects of drift. In practice several problems occur. Any angular displacement from a reference zero is sensed by an AC potentiometer. The term full scale here refers to the maximum force experienced over a given run at constant Rn. As built. As with a standard beam balance. This corresponds to ±0. drag obtained by mechanical means includes three-dimensional effects due to the side walls.002 lb. whichever is larger. the angle of attack of the two tips must be kept equal to that of the central portion and the gaps must be minimized. the dead weight of the airfoil and the support structure are counterbalanced with weights. Althaus 7 investigated the effect of a gap on the drag at low Rn and found that with a 0. The remaining forces (the lift and residual imbalance) are balanced by the torque from a brushless DC torque motor mounted on the beam axis. Spherical bearings were used to minimize moments transmitted to each linear bearing.0135 of C1 at Rn = 60k and ±0. A force transducer coupled to the carriage through a pushrod sensed the lift (actually half the total model lift was transmitted to the transducer). To achieve this.0055 at 300k. In addition. Two linear ball bearings spaced 8 in apart provided essentially frictionless movement for a carriage which held the airfoil and angle of attack control hardware. One side pivoted. the drag was increased 12% at an angle of attack of 9°. These effects can be reduced by using a three-piece model with only the central panel connected to the force balance.3%) gap and 250 mm (156%) span. The force transducer in this study was a servo balance rather than a standard strain gauge or capacitance type transducer. and the error signal is used to drive the torque motor until the error disappears. The airfoil model was mounted horizontally in the tunnel between two ~ in clear plastic (Plexiglas) end plates (omitted for clarity) to isolate the model ends from the tunnel side-wall boundary layers and the support hardware. the system was capable of measuring 7 lb (half of the lift) with stiction and hysteresis limited to 0. a pitot tube was surveyed through the wake 1. The residual friction (as well as some magnetic hysteresis) was further reduced by adding a small amount of electrical dither to the torque motor. The drag was measured using the momentum deficit method because the mechanical one is both difficult and expensive.12 Airfoils at Low Speeds A sketch of the apparatus used to measure lift is shown in Fig.25% of full scale or ± 0. and the other was free to move vertically on a precision ground shaft. 7. The current needed to generate the torque is a very linear analog of the torque. and therefore of the lift. 2. The overall system had an accuracy of ±0. Spanwise non-uniformity in the wake is well known 7 •8 • Indeed. The drag calculation requires only the streamwise component of the velocity. the drag force per unit span can be found as: I 00 d= P u(V00 - u)dy (1) -oo where the integral is performed perpendicular to the freestream. It was factory calibrated against a standard traceable to the National Bureau of Standards. and they were averaged to provide a better measure of the airfoil performance.Po)(. Surveys on several airfoils indicated that static pressures in the wake were nearly equal to the freestream static pressure. As mentioned previously. the measured total pressure is essentially independent of flow angle. and b. and u is the x-component of velocity at the downstream location. indicating that it was sufficiently far from the trailing edge so that transverse velocity components were negligible. The freestream velocity is V00 .JP:. This method of determining the drag is valid only if the wake survey is made in a region where the static pressure is equal to that in the freestream. A Baratron model 220B unit made by MKS Instruments. Drag values were found to remain constant as the survey location was moved upstream and downstream of the 1. (Using a single pitot tube and moving it through the wake provided better spatial resolution of the wake than using a rake with multiple. four spanwise stations spaced uniformly over the central 1 ft of the airfoil were used. requiring a sensitive transducer. transverse velocity components at the surv:ey location can decrease the measured drag.:- Vh>O.15% of reading. For pitot tube misalignments of Jess than 10°.b.25 chord location. thus. y is in the direction normal to the freestream. was used for this purpose with a full-scale range of 1 mm Hg and an accuracy of 0.P0 is the difference between the total pressure in the freestream and the total pressure in the wake. Inc. Equation (1) can be rewritten to give: I 00 d= 2 (VPdoo.) Based on the application of the two-dimensional momentum and continuity equations to a control volume about the airfoil 6 .b.Chapter 2: Experimental Facility and Measurement Technique the deficit.Po)dy (2) -oo where Pdoo is the freestream dynamic pressure measured with a pi tot tube which was 15 in upstream of the airfoil and 8 in below the centerline. fixed pitot tube locations. 13 . Drag was calculated using the difference between the total pressure upstream of the airfoil and that in the wake. This pressure difference is small and difficult to measure. downstream of the airfoil. the drag variation can be more than 50% at the lower Reynolds numbers. The important features and accuracies of this positioner are: Spanwise motion: 24 in Vertical motion: 14 in Resolution: less than 0. both directions Each axis was instrumented with a precision DC potentiometer and was driven by a small.slot were sealed to prevent air leakage into the tunnel. This decision to ignore speed was soon regretted when it became apparent how long each run in the tunnel required.) Using this two-axis positioner. 2. Each survey consisted of between 20 and 80 pressure measurements (depending on the wake thickness) with points nominally spaced 0. The arm that carried the pi tot tube projected into the tunnel through a slot cut in the floor. and the motors drove the carriage and potentiometers through a linkless steel and plastic chain.001 in. A 1 mm Hg full-scale unit measured the difference in total pressure between the wake and freestream as previously mentioned. The open-loop gain was quite high. the entire carriage was mounted on a large aluminum "U" channel which was mounted to the bottom of the tunnel floor. The carriage which held the pitot tube ran on precision bushings around centerless ground and polished rods. Because accuracy was the design goal.08 in apart.020 in vertical: 0.002 in Setability: 0.8). and both the arm and. (Changes were made later that resulted in some improvement in the speed. the surveys were made through the wake at four spanwise locations. . For stability. The third transducer had a 10 mm Hg full scale and was used to measure the dynamic pressure at the upstream pitot tube. Three pressure transducers (MKS model 220B) were used in this study. there was no attempt to make the positioning particularly fast.005 in. Each motor and potentiometer together with associated electronics formed a position servo loop. both directions Readout accuracy: spanwise: 0. however. since it was read directly from the feedback potentiometers. which effectively yielded a time-averaged drag value for each spanwise station. DC motor. Analog inputs to the positioner were provided by a computer with two digital-to-analog converters. geared. A typical survey through the wake took two minutes. the accuracy of the reading was independent of the gain. Another 1 mm Hg unit measured the difference between the test section stagnation pressure and atmospheric pressure to allow an accurate calculation of the density in the test section.14 Airfoils at Low Speeds A two-axis traversing mechanism provided position control for the downstream pitot tube (see Fig. a velocity ratio based on the Reynolds number was determined before every run. Drag was measured only for increasing angles of attack. 7 hours. Since the upstream pitot-static probe did not sense the dynamic pressure at the airfoil. Throughout this work. Using the continuity equation. Inc. First. A velocity was found with the fixed upstream pitot-static probe and with the downstream pitot-static probe placed near the tunnel centerline with the airfoil installed but generating no lift. 14-bit analog-to-digital converter interfaced to an IBM PC. and the location of the wake was found. so hysteresis was not examined. the tunnel velocity drifted slightly. which usually took about 1.02°. In all cases. the amount of run time would have doubled to 3. a calibration was performed to correct its reading. This was done for two reasons. After manually setting the tunnel speed to achieve the desired Reynolds number. the velocity at the airfoil was greater than that upstream where the freestream dynamic pressure was measured. Error estimates indicate that the accuracy of the measured Cz is ±1% and that of the Cd is ±4%.Chapter 2: Experimental Facility and Measurement Technique Due to the tunnel blockage from the lift apparatus installed in the test section. During a run. hysteresis is a sign of gross laminar separation-a high-drag condition. Next. respectively. the velocity difference between the upstream location and the airfoil was less than 6%. This investigation was directed towards 15 . The ratio was determined at several speeds in the neighborhood of the actual run speed. the angle of attack was increased and the process repeated. Thus. the four wake surveys were performed. The PC controlled the wake pitot tube position and the airfoil angle of attack. this process continued into stall. All transducer voltages were recorded using a Scientific Solutions.6. When they were complete.4 hours on average.Po were normalized by the instantaneous value of the freestream dynamic pressure. The accuracy in determining a was ±0. Because the effective area of the apparatus is clearly a function of Reynolds number. The angle of attack of the airfoil was controlled using a gear motor with a worm drive and a sector gear and was sensed using an angular transformer like that used in the force balance. Wind-tunnel corrections 6 were applied to values of Cz and Ca and were approximately 4% and 2%. Second. a polar at a given Reynolds number consisted of between 15 and 20 angles of attack from -3° to 15°. the measured lift and the . A linear interpolation based on Reynolds number was then used during the run to determine the velocity at the airfoil based on the dynamic pressure of the upstream pitot-static probe. the data collection was completely automated and proceeded as follows: The first angle of attack was set. it can be seen that the velocity ratio between the velocity at the airfoil arid the velocity upstream of the blockage is simply the ratio of effective areas. depending on atmospheric conditions. To ensure an accurate determination of the lift and drag coefficients. slow fluctuations in tunnel speed affected only the Reynolds number and not the determination of the aerodynamic coefficients. Usually. and Notre Dame 13 •14 • Comparisons of the available drag polars are shown in Figs. 2. 2. The majority of available data with which to compare the Princeton data is from the Model Wind Tunnel at Stuttgare• 9 •10 .4 hours that would have been required to obtain a complete drag polar at 1° increments in a. In addition to taking lift and drag data simultaneously. and 200k. Rn :::. Approximately 140 data points were taken. again because there was no interest in high-drag conditions and also due to time constraints. if the measured drag coefficient exceeded approximately 0. Results of the S2091 are compared with Stuttgart (1986) in Fig. the Stuttgart data indicates lower drag throughout much of the lift range. and comments are provided in several cases. In this case.10 compare drag polars obtained in the Princeton tunnel (using the E205B-PT model) with those in the Delft tunnel 8 and in the Model Wind Tunnel at Stuttgart 7 for the E205 at Reynolds numbers of 60k. At 200k all three facilities agree to within 10% over the central region of the lift range. Stuttgart tests of the S3021 in 1986 are compared to the Princeton data in Fig.19. See Fig.16 Airfoils at Low Speeds the characteristics of low-drag airfoils in application to RC sailplanes. . Fig. hence. Consequently. Furthermore.4 Comparison with Other Facilities Measurements in other facilities can provide a basis of comparison for the lift and drag obtained in this project. 2. Delft 8 •12 . 300k.050 the run was stopped. however. it is difficult to make a broad range of· systematic comparisons because relatively few of the airfoils tested in this project have been tested in other facilities at the same Reynolds numbers. lOOk. A listing is given at the end of the discussion. In this mode of operation. little effort was directed at designing and testing airfoils in the Reynolds number range 60k :::. the Princeton drag values are generally lower. Increasing and decreasing angles of attack are denoted by solid-circle and open-square symbols. the angle of attack was increased up to a pre-set value and then decreased. at 60k the agreement is worse. which was relatively slow.18. in many cases a second run was made in which just lift was measured.20 for Reynolds numbers of lOOk and 200k. Overall agreement is reasonable. however.9 through 2. and this process usually required 5 minutes-much less than the 3. A comparison with Stuttgart data from 1980 on the NACA 0009 is shown in Fig. The agreement between Delft and Princeton data at lOOk is also quite good. 12. comparisons were also made to data obtained from NASA Langley 11 . 2.20. This lift data is included along with the polar data in Chapter 12. Until recently the primary application for airfoils operating at Reynolds numbers considered here was for model aircraft. However. 2.2 for example. high-drag conditions were of little interest. 2. Hysteresis loops present in the lift behavior were then sometimes observed. allowing the angle of attack to be incremented relatively rapidly. respectively. Unfortunately. Chapter 2: Experimental Facility and Measurement Technique At Reynolds numbers of lOOk and below.4%.17. the agreement is poor. 1980) 9 2.15 FX63-137B-PT vs FX63-137 (Althaus. Nevertheless. 2. 1980) 9 2. Delft. 2. the general agreement between the data from the NASA Langley.14 FX60-100-PT vs FX60-100 (Althaus.13 E387 A-PT vs E387 comparisons 9 • 11 •12 2. Fig. Delft 12 . (3) methods of measurement. 2. and Stuttgart 9 • Note that the E387 model used in the Princeton test (E387 A-PT) is decambeted (see Fig. 1980) 9 2. 1985) 14 2. Fig. Fig.9 CLARK-Y-PT vs CLARK-Y (Althaus. Fig. but we have documented those of the present project to allow for future comparisons.13 shows a comparison of the E387 data from the present work to that from the NASA Langley 11 .16 FX63-137B-PT vs FX63-137 (Bastedo and Mueller.11 E214B-PT vs E214 (Althaus. the Notre Dame data indicates a significantly higher drag than either the Stuttgart or Princeton values. Fig. At this time it is difficult to determine how much of the disagreement is due to each of these areas. Fig.12 E374B-PT vs E374 (Althaus. and (4) model accuracy. 1986f 2. In the case of the FX63-137. the agreement is quite good. 1980) 9 2.18 NACA 0009-PT vs NACA '0009 (Althaus. By 200k.10 E205B-PT vs E205 comparisons 7 •8 2. Fig. 1986f 17 . Fig.20 S3021A-PT vs S3021 (Althaus. Fig.17 M06-13-128-PT vs M06-13-128 (Pohlen and Mueller. The discrepancies found in these comparisons are primarily due to differences in (1) flow quality. and Princeton is good. 10.16 and 2. Fig.13) which is reflected by a shift in the polar to lower lift values. Fig. Comparisons were also made with data from Notre Dame as shown in Figs. Fig. The camber error is approximately 0. with Stuttgart generally finding higher drag. 1983) 13 2. 1985) 10 2. 1986f 2. (2) accuracy of measurements. Fig.19 S2091B-PT vs S2091 (Althaus. ???w>tp.- with Variable Speed Windows for Tunnel Air Intake pwmry!Zd.· b' ' ' ' ' I T g' Working Section with Gloss Windows Fig.Turbulence 3' X 4' Smoke Tunnel ~ ~ \\~ Screens g' \ 14' -----i T ~.Turbulence..~ 0 I Low.Motor 'I I I.... Double-Con traction Cane 9:1\ .d~ ~ Lob Wall J 16 Bladed Fan \ \ \ Princeton University Low-Speed.75 ~ t-< Top View '-- ~ ..go "' ~ f} . 2. ::... 1750 RPM .'v:-! t ' ' ' jU Flow Straightener p 23' 7.1 Diagram of the Princeton University low speed wind tunnel (not to scale). _. l .._'- '- __L Turning Vanes 6 75 T l 12' 4' s::. Low.f-' 00 Outside Exhaust 50 HP DC. -----------------------------------.~---------------------------------..Chapter 2: Experimental Facility and Measurement Technique • . * 200k * q ~"' (/) Fig.g 'e>q.2 Fluctuating velocity energy spectra in the freestream 19 . 2. 300k • ~"' ~q. ---------------------------------. lOOk • 'oC!.20 Airfoils at Low Speeds . 2. C! ~ '$2 ~.. .2 Fluctuating velocity energy spectra (concluded). Fig. '*"' 60k C!.---------------------------------. ·-· ' Fig.Chapter 2: Experimental Facility and Measurement Technique r .3 Coordinate Measuring Machine (CMM) used for digitizing._.4 Holding fixture for digitizing the test sections.e cH '"'~' eerE n Fig. 2. 2. ' 21 . 22 Airfoils at Low Speeds 37 31/1 XZ POINT X= 13.00453 0.24506 0.53849 0.00651 -.04595 0.39368 0.04901 0.00123 o.54502 0.13491 0.00000 -.69655 0.82217 0.07922 0.7423 Lower L.83603 0.00748 -.99746 0.5 Typical CMM digitized data.97800 0.00721 -.oo238 0.14185 0.00659 0.86567 0.00000 0.90040 0.05262 0.04561 0.94486 0.2182 Upper L.03199 0.02491 0.06080 0.79285 0.01583 0.00435 0.02142 0.00753 0.70417 0. 2.05817 0.00714 -.9B829 o.38540 0. 1.07190 0.08634 0.00884 0.00000 0.03337 0.62461 0.00000 -.6939 Z= -1.00169 0.07805 0.00665 -.E. .00000 Chord= 11.01576 0.00546 -.6 Typical normalized coordinates from the digitized data.01274 0.00106 0.00657 -.80749 0.83894 0.94190 0.6770 Fig.7193 Z= -1.07430 0.23528 0.99740 1.3126.00375 -.990 0.6941 Z= -0.00767 -.02951 0.7199 Z= -0.08876 0. 13 711 XZ POINT X= 13.00083 0.00237 -.00464 -. 2.97400 0. AQUILA profiler data.1266 38 32/1 X Z POINT X= 13.E.49234 0.00259 -.00553 -. 15 9/1 XZ POINT X= 13.02575 0.02617 0.03831 0.1 0.08685 0.00587 -.01570 0.88509 0.00000 Fig.76590 0. Y traversing mechanism.8 Sketch of the X. . (Plexiglas end plates are not shown for clarity. 2. 2.Chapter 2: Experimental Facility and Measurement Technique 23 / Fig.) Fig.7 Test rig indicating model orientation and lift measurement method. t"' ...::. .9 Comparison polars: CLARK-Y-PT vs Stuttgart 9 . ' . .. T. 0..01 0. 2. . ················ . . T.04 Fig. .000 Model W. at Stuttgart 200. > . ::::::::::: : ::: ::::::::::rr: :.... at Stuttgart 0 0 ~~~~~~ ~~~~~~~~~~~~-~-~-~~ .. . ' . . :.000 Princeton 200. =· .:.. :..03 0. 1:::.... .. at Stuttgart 100.00 0. . .:::.000 Model W...000 Model W.000 Princeton 100.000 Princeton 60. I 0.24 Airfoils at Low Speeds CLARK-Y-PT vs CLARK-Y (Althaus. ' .. . ' ... : "() "(' "'t "!"'(''"' "'!"'t"·:. .. . . . EjJ'· !"' .::·: :~:: :::i:::i:::i::: ::: ~ . . T. 1980) o o A liX EB -v 60.05 ..02 0.. '' ' ······-···· . 00 = 200.Chapter 2: Experimental Facility and Measurement Technique E205 data comparison for Rn 0 EJ v 0.04 0.10 Comparison polars: E205B-PT vs Delft 12 and Stuttgart 13 25 . 2.05 Fig.02 0.01 0.000 Princeton Low-Turbulence Tunnel at Delft Model Wind Tunnel at Stuttgart 0.03 0. 02 0.04 0.000 Princeton Low-Turbulence Tunnel at Delft Model Wind Tunnel at Stuttgart 0.03 Fig. 0.00 = 100.26 Airfoils at Low Speeds E205 data comparison for Rn 0 8 'V 0. 2.05 .10 Comparison polars (continued).01 0. 0.03 Fig.01 0.000 Princeton Low-Turbulence Tunnel at Delft Model Wind Tunnel at Stuttgart 0.00 = 60. 2.05 27 .Chapter 2: Experimental Facility and Measurement Technique E205 data comparison for Rn o EJ v 0.02 0.10 Comparison polars (concluded).04 0. 05 . T.000 Model W. 2.28 Airfoils at Low Speeds E214B-PT vs E214 (Althaus. T.000 Princeton 200.01 0.04 Fig.00 0.000 Princeton 60.000 Princeton 100.11 Comparison polars: E214B-PT vs Stuttgart.000 Model W. at Stuttgart 100. at Stuttgart 0 I 0.000 Model W. T.03 0. 0. at Stuttgart 200.02 0. 1986) 0 EJ 6 llll III "' 60. ).... ···:···:··· ---: . .... .. T.) . ..' ' ::: ::: :::~:::~:::. 0. . . ...... . ).) ...).. at Stuttgart 100.03 Fig. .. T. .000 Princeton 100...···:···:···:···:··· ···:···:···:···:··· . ..... . ·:···:···:··.' '' .01 ' ' . .. .' 0.). .Chapter 2: Experimental Facility and Measurement Technique E374B-PT vs E374 (Althaus. .000 Princeton 60.12 Comparison polars: E374B-PT vs Stuttgart.' '' . .. ....)...... .). . .. 0..00 '' . .)... ' ' . .000 Model W.. . T.000 Model W. . . 2.000 Princeton 200..02 ' ' . 1985) 60. :::r::~::T:T:: :::~:::~:::~:::~::: :::~:::f:::~:::f::: ::: ::: ::: ::: ::: ... at Stuttgart o EJ A llX 8l -v L{) ~~~~~~~~~~~!::-~ . ' 0 ' ' ·-· . ). )... . .000 Model W... ' ' . at Stuttgart 200. .) ..). .' '' 0. ). . . 29 . .. .. .. . --~ ... .: ··~ : : : : ::rrr~... ···:.... Lr 0 Princeton & NASA Langley LTPT .-~ ~- I 0.. .... 0. . .... 2....04 Fig. ..30 Airfoils at Low Speeds E-387 data comparison for Rn ... . . . . ...··:. ... -~ ..:.. ' ~ ~-· . . .. .. -~--. .. ..J_::-'-.. :. . ::rrr: : : ···:···:<···:···---: ··:··· ···:··· ···:···:···:··· . .... ·: .. .. ··: . --:... :· .. . . : : : : . Delft University.::rr~ ::~::rrF· ::ELl:::::: :::::::l::r:-:::_f:Lcc :::ci:::i:::c :::::::[::: ::::::: :::c~:::~::: ::: "' o ... . : ··: :···: . .. ~··:···:· ·:··-:--· ···:···:· -:---:--· ···: ··:··· ···:··· ···:···:···:··· ....000 .. ... ~···:···:···:···:··· . ' ..05 .. . : ·.'·_::..... :· . .:r: ::r. . :::t:::tf... :···:···:·· :·. : .01 0.. ' .. ....yrr:c ···:···:···:···:···~··:···:··· 0 0 : : ::::::::: "' 0 :: : : : : : : : : • • ::rrrr: ::::.... ~ -~-.. : .00 0. . ' ···: ···:···:···:···:··· ···:···:··· ···:··· ···:···:···:··· ... ' .. .. .. . and Stuttgart..02 0.. ···:···:· ·:···:··· ···:···:···:···:··· ··:···:--... .·: .. ··:······· .... . ··t·· ..... ..:· ··:···: ..... ···:···..:1-:J~Pi£ :3~:..... ···:···:···'··· .:.::::: ::t::!::t:[::: :::L:!::: :::c ::LLL: : : . . . = 300. -~ ...' . L. : . ·: ..:-.13 Comparison polars: E387A-PT vs NASA LTPT.... : . . ..03 0... ..•.·· . .. ~ .. .. ..:. .. -... ..... ... ·:.... . .. :·. . . .. 2... :· . ..13 Comparison polars: (continued)..Chapter 2: Experimental Facility and Measurement Technique E-387 data comparison for Rn =200.01 0... ..... .. ............. :···:··· ... ·: ... . ·: .. ... . ....02 0.. :... . .05 31 ... I 0. 0. . ·:. . ·: . ··················· .... . ... ..03 Fig. : ..... ·:. . .. ··:· . . ··: .. ···:···:···:··· ..... ..... ·: ···:... . . . : . ·: ..04 0. .. .. . ·: ...... ·: ..000 Princeton NASA Langley LTPT Low-Turbulence Tunnel at Delft Model Wind Tunnel at Stuttgart 0 to.. . ..00 0. ·: ..... .. ··: ... a v u~ ~4-d ~~~~3:EU ········ "' 0 0 0 ········-···-·· 8'=+='P~~!Tb$H=I1_ ···:···:···:···:··...... ....... . .. . .....······ ' ' ..000 Princeton NASA Langley LTPT Low-Turbulence Tunnel at Delft Model Wind Tunnel at Stuttgart ..... : . ... '' ... ... ···.......: ... ... : ... .... .. . : ... ........ ... i. ... ... ... . ......... . : .... . . -·-···-. .. : •.. ..• : •••••• : •.. ..... .. . ..... .. : . :..' -·-· .. ...03 Fig.. ... .. .... . . ... . .. ... . ········ . . ...... . ~~----:~r=... ···:···:···:···:·······.........:... .. i. . i.... .::l::L:~::: ::: :::l::+::::::l::: : TT : : tJi·~~ ! [ T ••• 1: T ::::::L:t::::::: ::::::~: ~:::t:::t:::l::: :::t:::t:::t::: ::: :::t:::l:::l:::t::: :::::~~~~Zlc ::LL::::: ::: j::L::::L: ::::::::::::::::::: .... .:~~=~=~~=~·~·~:_L_=~~~=~~=~=-=--~~~=~~~ 0.... ... ... ................ : .. : . . .. : ... ... . .... : .. : . . • • • • • • • . : . ~---~..···:···:··· .... .~...... .. ....... ... ..•• : •...c. ......t~--:~~~ ·--~---~·-·: ... . .....00 0. . .01 0........ -·················· ... . ... .... 0. .. .. • • ..13 Comparison polars: (continued).. . . .. .. ..32 Airfoils at Low Speeds E-387 data comparison for Rn 0 = 100.... .. .. . . ........ ..... 2.. .. . : ......... .. . ' . .......j·::··:~·>j·· ~ 0 0 o l{l .. . ... .. . ··········· . LJ~ LJ~j[~iJ=cj:~:J:~~~-~~-~-~:~~~:~:~:~:=e:.... .... --':-+-t:.. :. .. 9 L... ..... .. &--:-:-'... ... ..... ..05 - .. : : ::::::::::: ::::::::::::::::::: .... . ... ... ....::·::::::::::· ::::::L::::?·~~~S... . ········· ....... . ...... ---~---:---~---:··· .:....04 0.. .02 0.... -·· ..:.. : . Chapter 2: Experimental Facility and Measurement Technique E-387 data comparison for Rn 0 t, 8 "' ... --- = 60,000 Princeton NASA Langley LTPT Low-Turbulence Tunnel at Delft Model Wind Tunnel at Stuttgart . . . . . . . . . . . . ---~---~--- ---~---~---~---~--- ·--~---~---~---~--- ---~---~---~---~--- ···:···:··· ... --· tY i:·J--[-·J·J···l:-~--L:·-~-~--l·:-~--I:-~--L:··J·:~--~--~--I;-~-~t;"""~~~~--~-~~~:~:-~:~~;:~:~-~~d-~:::: ~-···:···:··<K·;..,- '"""";,.······ . : :···:···:··· s . . · ·:· ·:· · · ·:· ·:· ·:· ·:· · .,., 0 0 0 ~~~~~~~~~~~~~~~~~~~ .,., 9 >;::,:.. ... ?.~:::+- ; ... :...: ... :...: ... ... ;... ~ ... ;... ;...... ;...;.. : ~ -~·-· .. .' --~--- .. : ... : ... : .. : .... ..: ... : ... : ... : ...... ... ; .. ;... ;... ;.....;... ;... ;... ~--- ::::::::::::::::::: ::::::::::::::::::: ::::::::::::::::::: ::::::::::::::::::: ::::::: ::: ::::::: L_~~~~~~~~~~~~~~~~~~~~__J 0.00 0.01 0.02 0.03 Fig. 2.13 Comparison polars: (concluded). 0.04 0.05 33 34 Airfoils at Low Speeds FX60-100-PT vs FX60-100 (Althaus, 1980) 0 EJ 6 lll EB "' 0.00 60,000 Princeton 100,000 Princeton 200,000 Princeton 60,000 Model W. T. at Stuttgart 100,000 Model W. T. at Stuttgart 200,000 Model W. T. at Stuttgart 0.01 0.02 0.03 0.04 Fig. 2.14 Comparison polars: FX60-100-PT vs Stuttgart. 0.05 Chapter 2: Experimental Facility and Measurement Technique FX63-137B-PT vs FX63-137 (Althaus, 1980) 100,000 t:. 200,000 m 100,000 '9 200,000 EJ 0 Princeton Princeton Model W. T. at Stuttgart Model W. T. at StuttQart .........,··:· .···: ··:··¥:·· ···: . ,..... ,.B'1f(~::::···;·· ,.. ···:···:···;···:··· ::::::r: ::::::vrl ::::r.·.,.--~ ::~::.,. +~::rr: :::f::::::r::::: :: ::: .~~0. ···III IYtL iI~Jic••• III· : : : : /1 : : : : : L; : : : : : : : . :···:··:··:··· .. :;~t: ···:··:···::6:; ···:···:···:··· ... ···:···:···:···:·· v 0 0 ~ I :.rrrr: .+T'p~~:.!:+JE_ ~rr: :_. ::rrrr: .. .. .. ' .. . ' ' .. . . .. .. .. .. .. .. . . : [:::[:::[:::~: :: :[: :[: : : : : : ::::::::::::::::::: : : : : : :[: : : : :::~ . r:FF -m ' ' ::t:t::::::::: ::::f::t:: ::t::: :::::::::)::r: ::r:::) : : : ::r:::::t::r: 0.00 0.01 0.02 0.03 0.04 Fig. 2.15 Comparison polars: FX63-137-PT vs Stuttgart. 0.05 35 36 Airfoils at Low Speeds FX63-137B-PT vs FX63-137 (Mueller, 1985) o 100,000 Princeton A 200,000 Princeton m 100,000 Notre Dame ""~ 200,000 Notre Dame A A.--A l() ~ ...... ···i···i· ....:,... ;.:...::... :.),.~1~0' . ::.,.:. ···t··· ···i ··t··· ···t···t···t···t··· ; .. , ...... ···~···~··· ... ... 0 ~ ~~ 0 0 l() ::: ::: ~-· ... .;..., . ... ··~··· ···~··· ···~ ··~· .. ···~···:···:···:··· ::lT:: ::r::::1::::: :::[:.-.~·.-r:::: ::r: :::[:::[::: ::::::L~~~/'1· . . . . . . . .. . . . . ~. N .. llif lif~J lJ(~~-4?:ft'J[[ic? I ill ; ::: ' • rNt ii'!~rt~• ::; : . • .: :. .i ; . '.. :::: . . • • • • ··················· . . . ... ··················· .. .. .' . .. .. ··················· .. .. .' ' ··················· . . . .. ··············· .. .. ··················· . .. ' ··················· .' .. ··············· ' . .. ... ··················· .. .. .. .. .' . .. .. ··················· .. . . • • • • . . . • • • • • ' • 0 -- ::t:::t:::t:::j::: :::f:::j:::t:::f::: :::) ::t:::;-::j::: ::[:::::::::: ::: :::::::::::t:::;::: 0 . . . . . . . . . . . . . I 0.00 0.01 0.02 0.03 0.04 0.05 Fig. 2.16 Comparison polars: FX63-137-PT vs Notre Dame. Chapter 2: Experimental Facility and Measurement Technique M06-13-128-PT vs M06-13-128 300,000 Princeton 300,000 Notre Dame (Mueller, 1983) 0 0.00 0.01 0.02 0.03 0.04 0.05 Fig. 2.17 Comparison polars: M06-13-128-PT vs Notre Dame. 37 38 Airfoils at Low Speeds NACA0009-PT vs NACA0009 (Althaus, 1980) o ~ 181 l!K . 60,000 150,000 60,000 150,000 Princeton Princeton Model W. T. at Stuttgart Model W. T. at Stuttgart . . . . . . .. .. . .. . . . . . . . . . . . . ... . . . .. :' ... : . . . . .. .: -.. .: - .. .: ... .:. . . . .. : ... : ... .: ... :. .. ..':.... .: ... .: ... .:. . . . .. .: ... .: ... .: ... .: ... ' .. -............................ . ' ' .. .. .. . . . -·... - ..................................... .. .. .. . .' .. .. .. . ...................... • . ~.. • • • • • • • • • • • • • 0 • • . ::!:::. ::::::. .;: ::::1:::1::;;~ ;h~:f- ·. ...~¥<::~: . ;_ -; . :. .,.. . -~· · . . . . ~·... . ,,,,,~~F-.,,,,,,,, ::::::::: :::::::: /:l··:'f .,.1.:~c : : : : : : : ~ 0 0 : : : : : : : : : : : : : : : : : : : : :::~::::: ::):1-:J:;::::r:: : : : : : : : : : : : : : : : : : : : : : ~:::::: ~::r : : 1: <:> : : : : ~ II~•, 1 ...:::::::::::::: ::::::::::::::::::: ::::::::::::::::::: o.oo o.o1 : : : : fiiiJIIIiil, o.o3 .of o.o4 Cct Fig. 2.18 Comparison polars: NACA 0009-PT vs Stuttgart. o.os Chapter 2: Experimental Facility and Measurement Technique S2091B-PT vs S2091 (Althaus, 1986) 60,000 Princeton 100,000 Princeton A 200,000 Princeton llll 60,000 Model W. T. at Stuttgart m 100,000 Model W. T. at Stuttgart v 200,000 Model W. T. at Stuttgart 0 o l{) Lra ~~~~JA+~~~~~~~~~~~~~~~~ l{) 0 I ~~~~~~·~·~~~~~~~~~-L~~~~ 0.00 0.01 0.02 0.03 0.04 Fig. 2.19 Comparison polars: S2091B-PT vs Stuttgart. 0.05 39 40 Airfoils at Low Speeds S3021A-PT vs S3021 (Althaus, 1986) EJ e. EE ""' 100,000 200,000 100,000 200,000 Princeton Princeton Model W. T. at Stuttgart Model W. T. at Stuttgart . .. .' ··················· ' .. .. ' . ... ··- ........... -·· ' . .. .. ......................................... .. .. .. .. 0.00 0.01 . . 0.02 0.03 ··- ..... . 0.04 Fig. 2.20 Comparison polars: S3021A-PT vs Stuttgart. 0.05 e. In addition. The effect on the shape of the polar is to produce a bulge in the mid-C1 range that is concave towards the C1 axis (see Fig. the SD7003. the shape of the polar clearly reveals the severity of the laminar separation bubble. To assist the general reader and to avoid confusion we here define those terms that are specific to low Reynolds number aerodynamics in order to supplement what can be found in textbooks 15 •16 . however. This type of separation and the subsequent formation of a laminar separation bubble are the principal reasons for the degradation in airfoil performance with decreasing Rn. For example.) If one compares the tripped and untripped cases (Figs. We also have gone into greater detail in the earlier airfoil discussions in Section 5. with increasing C1 the drag decreases again just before stall. even without trips. Here. At low Rn the situation is often markedly different. to a more favorable convex shape for the tripped case.1.1 Laminar Separation Bubbles As described in Chapter 1. the graphs. separation is a major factor and can contribute a large increment of drag not normally found at higher Rn's.19 and 12.13).Chapter 3: Low Reynolds Number Terminology Chapter 3 Low Reynolds Number Terminology The concepts of modern airfoil design and the attendant jargon are familiar mostly to specialists. For such airfoils. Interestingly.g. are more typical of those for higher Rn's. At high Rn (greater than 1 million) a graph of cl VS cd for most airfoils shows a rounded appearance with the convex side towards the C1 axis. 12. 41 . 12. the concept of a bubble ramp is derived from a transition ramp. See References 15 and 17 for examples of polars at higher Rn's. be minimized by proper design. 3. laminar separation takes place at low Rn due to the reluctance of the boundary layer to make a natural transition from laminar to turbulent flow on the airfoil surface. the E214 has a major problem with this at Rn's of 60k and lOOk. but they are not synonymous. For example. some of the terms commonly used even by aerodynamicists have modified or expanded meanings when applied to low Reynolds number airfoils.1 so that the concepts will be familiar when they are referred to more briefly in the later ones. Separation can. (This will be discussed in Section 5.22) the difference in shape illustrates the difference in the separation-the polar changes from concave for the untripped (separated) case. Because of these effects. 2 Trips and Bubble Ramps Several means can be used to destabilize the laminar boundary layer and promote an early transition. and may be more appropriately called a "bubble" ramp. as applied to airfoil aerodynamics. the turbulator becomes unnecessary and handicaj)S the performance by making transition happen too early. these are external ridges or bumps applied to the surface of the airfoil in a direction parallel to the span. Is it better to use only a bubble ramp or only a turbulator or a combination of both? Usually airfoil turbulators are simple two-dimensional strips. At low Rn. In fact. C d. or C m. the transition ramp is still useful. In this case the ramp would probably be better. for a Rn near 60k and at moderate lift coefficients. full-size sailplane airfoils. At higher Rn 's than those found on model aircraft. almost the entire upper surface of the airfoil is needed to ensure transition and subsequent reattachment. but some three-dimensional trips (such as zig-zag tape and bump tape) have proved highly successful in application to modern. turbulators may be better than bubble ramps. gradual pressure recovery. Another method of inducing transition is through the use of a turbulator or trip. Several detailed experiments. called a transition or instability ramp. but as the Rn increases. the bubble shortens naturally.42 Airfoils at Low Speeds 3. the question remains unclear. trips and ramps were used to begin to explore their respective benefits and operating regimes. more and more of the airfoil surface is required for the bubble ramp 18 •19 . The gradual pressure recovery of the ramp in this case shortens the length of the bubble by shortening the distance required for reattachment. Currently at issue in the design of low Reynolds number airfoils are the factors governing the use of bubble ramps and turbulators. In the middle of the Jow-Rn range. The purpose of the ramp in this case is to ensure that the boundary layer is fully turbulent and energetic before reaching the main pressure recovery. The most direct means of doing this is through the airfoil shape. a short. They protrude into the boundary layer in such a way as to energize it sufficiently to promote transition. have amply demonstrated that if an airfoil has high drag and hysteresis owing to a laminar separation bubble.3 Airfoil Hysteresis The term hysteresis.1• at a given angle of attack when this angle of attack is approached from a higher and then a lower value. 3. a turbulator often alleviates these adverse effects by shortening the length of the bubble. As the Rn decreases. Thus. This behavior may be seen . At very low Reynolds numbers. Typically. In our research. is sometimes used before the steeper main recovery. as summarized by Mueller 4 . Many free-flight model aircraft employ turbulators to improve performance. although it needs to be longer. means the difference in c. However. the detailed effects of hysteresis were not closely examined. 12. and then suddenly jump up to rejoin the original Cl vs a curve.Chapter 3: Low Reynolds Number Terminology in the in the Cl vs a plot. In general. Since this effort concentrated on those airfoils with low bubble drag. Hysteresis in the aerodynamic coefficients with both Reynolds number and angle of attack is common to many of the airfoils tested. Invariably. See for example Fig. as the angle of attack is increased from zero to stall. the Cl will not retrace its original curve. when a is decreased. 43 . airfoils with hysteresis in the middle of the Rn-envelope of the aircraft should be avoided. hysteresis is a sign of a large. In cases with hysteresis. the C l will usually reach a maximum value and then drop off at some particular a. rather it will stay at the "stall" C1 until a is somewhat below the previous stall value.2. laminar separation which in turn yields high bubble drag. 44 Airfoils at Low Speeds . JCi This relation emphasizes the fact that the Cd should be minimized for a value of C1 and the corresponding Rn.. Based upon the wind tunnel results. In particular. Thus. it predicts airfoil performance more accurately than the current version of the Eppler and Somers . Over the Reynolds number range considered in this investigation. the optimum airfoil design is clearly dependent upon the configuration and desired tasks of the aircraft for which it is designed. If the performance was poor. Before discussing airfoil design. With a minimum number of parameters. the agreement with the experiment at Reynolds numbers of 200k and greater is very good. The Eppler and Somers code formulates the design problem in a way that allows quick and easy manipulation of the airfoil shape. the agreement depends on the choice of the n value used in the en transition criterion. The ISES code solves the two-dimensional Euler equations coupled with a momentum integral boundary layer formulation using a global Newton method. the new airfoil was redesigned and the process repeated. because this code does not accurately predict the performance of airfoils in the Reynolds number range considered here.Chapter 4: Project Design Methods and Goals Chapter 4 Project Design Methods and Goals Three main tools were used to design new airfoils: the Eppler and Somers design code 20 . and then predict the performance at a Reynolds number of 200k using the ISES code. Upon reaching a suitable design through this iteration process. However. the ISES code written by Drela and Giles 21 •22 . a wind tunnel model was built and tested. The designs discussed below are based upon RC sailplane 45 .code. However. and the wind tunnel described previously. almost any desired velocity distribution can be obtained.and Somers code. it should be pointed out that for any aircraft in straight and level flight the relation between C1 and chord Reynolds number is given by: Rn oc 1 . The design approach was to generate an airfoil with the desired inviscid velocity distribution using the Eppler. wind tunnel results were the ultimate test of an airfoil. While the ISES code provided a relatively good estimate of the performance. the new airfoils were further refined and the process repeated. it was used mainly to obtain the inviscid velocity distributions and to give an estimate of the transition point behavior. Over the forward 40%.24-12. At low Reynolds numbers. To reduce the drag.1. however. but with the transition point farther forward. Before this point is discussed further. transition does not occur at a point but rather over some finite distance. and the majority of recovery takes place over the aft 50% with a relatively strong adverse pressure gradient.55 is shown in Fig.000. A comparison between the experimentally determined drag polars for the E374 and SD6060 is shown in Fig. Removing the kink shifted the transition point farther forward for C1 greater than 0. At lower Reynolds numbers.8. with relatively little importance placed on the performance at low speeds. . However.5. the separation bubble will be shorter and the drag will be lower. As a result of the kink in the velocity distribution at 40% chord.5 to 0. The inviscid velocity distribution about the E374 for a C1 of 0. The experimentally determined drag polars for this airfoil are shown in Figs. It is commonly used on aircraft intended for high speeds. respectively. The resulting transition point behavior and velocity distribution are shown by the dashed lines in Figs. the velocity changes little. A popular RC soaring. This longer region of smaller adverse pressure gradient is termed a bubble ramp.as shown by the dashed line in Fig. and then decreases from 0. Knowledge of the shape of the transition-point curve is helpful when designing with the Eppler and Somers code because it is similar to the distribution of design parameters which specify the airfoil (a' with 11) 20 • In the "redesign" of the E374.0 to 0. it is important to observe the behavior of the transition point on the upper surface with increasing c(. the drag increases dramatically as C 1 moves from 0. the general principles apply to any type of low-Reynolds number aircraft.) In this case the point refers to the location at which transition was predicted to occur by the Eppler and Somers code using a method based on the boundary layer shape factor for Rn of 200.3.2.1 and 4. 4. There has been a reduction in drag throughout the central portion of the polars for all Reynolds numbers because the bubble ramp has reduced the length of the separation bubble. 4.5.29. This behavior indicates the formation of a large laminar separation bubble on the upper surface. In this case.1. the pressure gradient should be reduced. the transition point moves rapidly forward with C1 as shown in Fig. (Some of this . (Of course. then the recovery region must start farther upstream.46 Airfoils at Low Speeds configurations. the kink in the velocity distribution was removed to define a new airfoil-the SD6060. 4. This airfoil works well at high speeds because of the small values of the drag coefficient at the higher Reynolds numbers throughout a range of low C1 values. if the same pressure differential is to be recovered. cross-country airfoil is the E374. 4. separation will occur earlier because of the steeper initial gradient.2. A "kink" in the upper-surface velocity distribution beginning at 40% separates it into two distinct regions. 12. 4. this pressure gradient results in a large laminar separation bubble. even at 60k. This airfoil has an upper-surface velocity distribution which is similar to the E374 in that it also contains a kink. the lift coefficient decreases so that for typical low Reynolds number configurations. As discussed earlier. What is not known at this -point is how far this design philosophy can be "pushed". it is possible to design entirely new. A further example illustrating the effectiveness of a bubble ramp in the uppersurface velocity distribution can be seen by comparing the E205 and the 83021 23 • The E205 is usually used as a "multi-task" airfoil because of its relatively good performance at both high and low lift. These examples illustrate that significant improvements can be made over existing designs by relatively minor changes in the velocity distributions (which. such as the 8D7003. The differences are similar to those noted between the E374 and 8D6060. that is. the optimum shape and location of the ramp remain to be determined. low Reynolds number airfoils that show little or no evidence of increased drag due to the bubble. which is consistent with the smoother forward movement of the transition point.9. as the speed increases.9% thick and the 8D6060 is 10.0 is more gradual in the case of the 8D6060. Figure 4. 47 . Even though improvements have already been demonstrated. the E374 is 10. The velocity distribution of the 83021 is essentially the same as that of the E205 except the kink has been replaced with a bubble ramp as in the 8D6060. for low Reynolds number aircraft. Employing airfoils with bubble ramps on model aircraft will provide further insight into the benefits of this type of design and will help guide further study.Chapter 4: Project Design Methods and Goals reduction in drag is due to a thinning of the airfoil. Other airfoils. the increase in drag as C 1 approaches 1.9. at 300k the lift coefficient would be considerably less than 0. at all Reynolds numbers the drag of the 83021 is lower than that of the E205 in the central region of the polars. the 83021 will perform better than the E205.4 shows a comparison between the drag polars of the E205 and 83021 at several Reynolds numbers.) In addition to the decrease in drag in the central region. at the highest Reynolds number (300k) the E205 has lower drag than the 83021 for C1 = 0.4% thick. of course. directly alter the airfoil shape). demonstrate that if sufficient attention is paid to the control of the bubble. However. Thus. 0 x/c Fig. 4.r> 0 --------------.----0 oi----r--~----~--~---r--~--~----r---~--~ 0.r. ' '. 4.0 0.48 Airfoils at Low Speeds .55). 0 ~.OL = 5° ( cl = 0.-------------------------------------------~ ""' 0 E374 SD6060 ~ 0 '\ -N "' uo ' '' ..t.0 0.1 Inviscid velocity distributions: Ctw. '' ' 0 0 0 "'0N I 0 "'0I 0.2 Upper surface transition point at Rn = 200.5 1.r> ~---------------------------. ..--=---- --- 0 ~ t 8 '' ' ' E374 SD6060 .5 x/c Fig.000 as predicted by the ~~ppler code.. :r--~----. J -. .. ..... .. .. ~ .. ···:···:··· ···:···:···:···:··· ···:···:···:···:··· ···:···:···:··· --.SD6060-PT = 300..... .000 = 100.' -~ ''!" ~---~- ' ---~---:--·: --~- ' ..01 0... .000 . . .. ·! . :.. . -~. .E374B-PT = 300... ~ ..05 Fig.. .....03 0..... -~ ...... '' ..02 0... .... ..... .. . .. '~ .000 . ·: ' " " ·-..'"'' '~- --~ ~ liiittJi~~ ~ 1::::::~ till · · · · 'f.. .. . ······· .. .... . .. . ... . ~ ..· · U"o · · · N ·.04 0.... . .. .. --~ ...000 .. .. ---:---:---:---:--......Chapter 4: Project Design Methods and Goals 49 E374B-PT and SD6060-PT o v ~ ~> Rn Rn Rn Rn = 100....00 0.... .... ' ' ..... -~ .. ... 4. .. -: .. . .· ~ •••i·•i•••l¥ •••l••v~Efl••• ••i•·•i•••l•••l••• • 1• 1• •1•·•1• • ~ ' ' ' t\:9 ' ~~ ' ' ' ' ' ' ' ' ' ' ' ' ' !Ill' J~~ FE IIII till 0. . ....'" ---:---:--. : .··: .... .. -·· ...-..... ' ....3 Comparison polars: E374B-PT and SD6060-PT at two Reynolds numbers. ···:···:···--:---:---:---: .... -~ . . . ' .. . ······ .. .... . ..... .. ... ~·· . ···:···:··· ········-··:·-·:··· ···:···:···:··-:··· ···:···:··· .000 = 300. ..000 = 100... .. . .02 .....:::!:::!:::. . .. ···~·· ~.03 0.. . ....::: :::::::.. . .. .. ' ...~ ~~~·-'·· _:_g .. : : Y'-':.. ···•·····--······-· .-4--....' ..... ..00 0. ' ... .$.. .... ~-~ . 4.. ···:··· .. ..... ..04 0. .. . ...01 0. . ... . . ··-·-·····-·-····· . . .ih::!:):::::: :t:::::t:::::: :::::::::: ::::::: :::::::L::L: ij(J::: ::::::!l~~::: .....:::::::::: :... .. tli" .. . '' .....000 = 300....000 E205B-PT S3021A-PT ' ..-~~=. . . ..... ..::: 0 0 ... .. . ···[ --j....' .05 Fig.. ' 0.. ... .. ... '' ··················· . = 100. ..:::!:::. . -·--········· ' ...-..... .. .. ·········--········ ' . ::t:::: L ::: ~::: ::::::. ..4 Comparison polars: E205B-PT and S3021A-PT at two Reynolds numbers. C! . 0...50 Airfoils at Low Speeds E205B-PT and S3021A-PT o Rn Rn Rn Rn v <:> e.. ......... ......... '' . On the other hand. Some miscellaneous notes follow: (1) For several of the airfoils camber-changing flaps are recommended for best performance.not originally intended for use on RC sailplanes. thermal-duration. particular aircraft. but have come to be used for this application by trial and error. This is one of the reasons why there are so many "favorite" airfoils. Following these comments. F3B. It should be noted in these examples that while one airfoil may be good for a particular configuration. Emphasis is placed on highlighting the important characteristics with respect to the other airfoils. Furthermore. To this ~nd. Many of the airfoils tested were . In some of the airfoil discussions we give examples of sailplane performance in order to compare one airfoil with another. the reader is left to make the final decision as to which airfoil is best suited for a particular sailplane. for example. "flaps" and "full-span flaps" and "camber-changing flaps" are used interchangeably. Even though most of the discussions are related to the RC sailplane. the SD-series airfoils were designed for different. a computer and an accurate performance prediction program are invaluable aids in the airfoil/aircraft integration process. it does not necessarily follow that it is good for all configurations. smaller sections discuss: (1) stall behavior (2) trips and surface roughness (3) trailing edge thickness (4) surface waviness and contour accuracy. the aerodynamic phenomena described are co=on to all airfoils operating at low Reynolds numbers. Rather than tailor a design to one. etc. In light of this. the new SD-airfoils (as well as the DF-series) were designed for the RC sailplane. multi-task. In these cases. Additional improvements in performance can probably be made by properly integrating the initial airfoil design process into the overall aircraft design 24 . (2) At the end of each airfoil discussion is a list of related airfoils given in order of decreasing similarity to aid the reader in the selection process. 51 . general classes of flying. the airfoils are by no means restricted to this use.Chapter 5: Comments on Airfoils Chapter 5 Comments on Airfoils In this section the performance of each airfoil is discussed in detail. s. (5) The airfoil maximum thickness and camber is listed at the end of each airfoil discussion for quick reference. bump. one may chose to build using the actual coordinates (with the PT-designation) listed in Chapter 7 for this purpose. or some other major modification. etc. u. etc. u.3. no specific details of the ramps are provided and airfoil velocity distributions are not given.2. as predicted by the Drela and Giles ISES code.8 for Rn of 200k for some of the airfoils. (7) In addition to testing the plain airfoils. For example see Fig. C.7 shows the geometry of the plain. is used to indicate the different models. (11) Although some of the discussions of the airfoils deal with the effects of bubble ramps. the effects of a number of different boundary layer trips were examined in a search for improved performance. For an estimate of the moment about the quarter-chord point. the actual coordinates should be used for any theoretical computations. should not be confused with the model "version" designations A. .s. C. (8) Airfoil moment data were not taken. When the trip type is not explicitly stated. and blowing trips. These abbreviations have the following meanings: upper surface upper surface trip lower surface lower surface trip xjc normalized trip position measured from the model leading edge to the trip leading edge h/ c normalized trip height w/c normalized trip width. B.52 Airfoils at Low Speeds (3) As previously mentioned in Section 2. the designation "-PT" (for Princeton Tests) is used after the names of the actual airfoils tested to distinguish them from the nominal or ideal airfoils.t. Figure 13. Depending on the accuracy of the wind tunnel model. C.s. the trip is the simple two-dimensional trip strip. Also. zig-zag.22. In some cases the versions may differ by the type of surface finish. (9) Airfoil polars and lift plots are gi:ven separately in Chapter 12. (6) If two or more models of the same airfoil were built. Note _that the trip "type" designations A. B. l. 12. t. the designation A. Chapter 8 lists the average moment coefficients over the range 0. these are not important with regards to building and actually using the airfoils. they are also listed in tabular form in Chapter 9.s. (4) The nominal Reynolds number is used in the labels of the airfoil polar plots while the actual Reynolds number is listed with the tabulated data. B.2 < C1 < 0. l. Therefore. (10) An abbreviated description of the trips is found in the figure titles. etc. since in some cases the deviation from the nominal is large. the addition of a flap. a computer program (available from Fraser) for evaluating and comparing sailplane performance. The popularity of the AQUILA waned. Airtronics introduced the SAGITTA with the Eppler 205 airfoil. 12. Also see: S2091. One can only speculate on the performance differences between Skip Miller's "modified" AQUILA and the stock version. 10. FX60-100 Digitizer plot: Fig. (13) The data used in generating the polar plots is tabulated in Chapter 13 and is keyed according to the airfoil name and figure number. and 1.1 Polar plot: Fig. Recognizing this deficiency. pounds. the polars presented here compare aircraft that are identical except for the airfoil and weight. No allowances are made for parasitic drags nor is any correction applied for a non-elliptical lift distribution. SD7032. However. The configuration is shown on the graph.Chapter 5: Comments on Airfoils (12) There are plots comparing the nominal and actual airfoils (termed here "digitizer plots") in Chapter 10. 5. and production was finally discontinued. (14) The aircraft polars in this chapter were made with SAILPLANE DESIGN.38% Ca. nevertheless.25.) Neither is significant when comparing similar aircraft. which had the desired high-speed qualities. and plots comparing different nominal and actual airfoils in Chapter 11.1 Airfoil Discussions AQUILA • AQUILA-PT (Fig. 1. the high camber was a handicap at high speed. seconds. it is hard to believe that in 1977 Skip Miller flew his "modified" Airtronics AQUILA sailplane to a first place finish at the first F3B RC Soaring World Championships. 12. 12. The three sloping lines are lines of constant sink speed: 1. All units are feet.5 ftjs.1) With the level of sophistication in·F3B today.2 Thickness: 9.0. (The former are impossible to accurately quantify and the latter would be very small in any case.05% 53 . the AQUILA airfoil gave excellent thermal performance. the stock AQUILA was a formidable RC sailplane from the mid 70's to early 80's.1 Lift plot: Fig. The profile drag of the wing is taken from the tunnel data and the induced drag is calculated using lifting line theory 15 for both the wing and horizontal stabilizer. In comparison to other popular airfoils of the time. Although SAILPLANE DESIGN allows the user to vary any parameter. The data is ordered as it appears in the figure title.lnber: 4. SD7037. The fuselage and empennage profile drags are calculated using equation 5.18 of Reference 15. 72% Camber: 3. Although only the third airfoil to be discussed. The effects of a laminar separation bubble are apparent only at a Rn of 60k through the mid-lift range. 0. 83021. reaching a minimum of 0.3) The famed CLARK-Y airfoil needs no introduction. However this bubble does not necessarily detract from the performance since the drag rise occurs mostly in the mid-lift range. 12.0 3. low-lift performance is superior. l is a mere 0. but the range of lift is significantly less. while maintaining . a flight that duplicated in reality the mythical flight of Daedalus and his son.3 Lift plot: Fig. the variety of performances and the trade-offs made to achieve them are becoming clear.2 2. Also see: DF101. was designed by Mark Drela using the I8E8 code developed by him and Giles 21 •22 • As was mentioned in Chapter 1.55% DAE51 • DAE51-PT (Fig. 12. a range not used by the vast majority of sailplanes at 60k. it is perhaps the most popular airfoil ever used on both full-scale and model aircraft. 12. Thus it was not designed for one operating point. No bubble "bursting" for Rn greater than 75k 4. The DAE51 was designed for the propeller of the DAEDALUS.5 ::. Icarus. where the drag coefficient reaches a maximum of 0. No doubt the CLARK-Y will stay popular for some time to come.2 Polar plot: Fig. human-powered aircraft 25 .9 for 60k.032 at a C1 of 0. 12. What is surprising is the small price for the improved performance-the maximum lift coefficient (Cl~ . 83010. c~~ .027 at a C1 of 0. Transition ramp optimized for Rn = 125k.5. The drag of the DAE51 at Rn of 300k is considerably lower than the AQUILA and CLARK-Y.1 less than the AQUILA. 10. The requirements were 26 : 1...4 Thickness: 11. As compared with the AQUILA. low-Rn regime of the RC sailplane. 8D5060 Digitizer plot: Fig. this same code was used to analyze the new 8D-series of airfoils. :::: 1. C1 ::. the drag reduces as the bubble shortens. the high-speed. along with several other DAE airfoils.54 Airfoils at Low Speeds CLARK-Y • CLARK-Y-PT (Fig.5) The DAE51. For application to RC sailplanes a broad lift range can sometimes be recovered by use of a full-span flaps. but rather for the range of anticipated conditions. 1. At the high-lift. Thickness less than 9% The achievement of these design goals was demonstrated by no less than a record 74-mile flight from Crete to 8antorini across the Mediterranean. 12. Over most of the low-drag range. The design requirements were: 1. fiat bottom aft of 30% 3.3 Polar plot: Figs.6) The DAE51 was tested with a substantially thickened trailing edge (see Fig. E193 Digitizer plot: Fig. The purpose of the variations was to test the effects of a controlled contour change in a region previously recognized as important to performance.2.Chapter 5: Comments on Airfoils the low-drag characteristics of the unfiapped airfoil. Also see: S4061. Similar results were observed by Althaus at Reynolds numbers 1-3 million 27 . See Figs.37% DAVID FRASER AIRFOILS (DF) This series of airfoils was designed for very large sailplanes where thin airfoils are not structurally possible. the same physical section was used for all three. minimum drag at cl = 0. SD6080. 11% thick 2. 12. the DF102 and DF103 are variations on the basic airfoil. the DF -series is thicker and has a different thickness distribution. these airfoils as a group provide a measure of the sensitivity of the airfoil performance to surface modifications. the DFlOl 's performance is quite good. the lowdrag range could probably be extended up to C 1 of 1.7 Camber: 3. suggesting that deviations either way from a good nominal design are as likely to hurt performance as to help it. • DAE51-PT thickened trailing edge (Fig. 12.1. The DF103 has its thickness decreased by approximately the same amount in the same place. The DF102 has an increased thickness (about 116 in max. As further examples. however.6 Lift plot: Fig. whether deliberate or inadvertent. a 4-5% increase in drag was found.8) As the polars show.. The starting seed was the SD5060.1 and 11. 12.5%) between 2% and 30% on the upper surface. 10. and modified a second time to become the DF103. local changes in the airfoil contour.5. To make the test more precise. E387. the initial design is the best. 0. With a 20% flap. 5. it was modified once and became the DF102. which would make it competitive against the unfiapped AQUILA and CLARK-Y. Besides demonstrating the effects of minor. 11.2. 12. and the same trends were observed. in 55 .1 for contour) to measure the possible loss in performance. Compared to other airfoils. The DFlOl is the initial design·.98% Thickness: 9. • DFlOl-PT (Fig. the E374 and SD6080 were also tested with thickened trailing edges. 00% Camber: 2. CLARK-Y. a condition that can cause the wing tips to stall before the inboard sections. CLARK-Y. In conclusion.2 Polar plot: Fig. something that could be expected from the CLARK-Y's much higher camber (3. but rather the low-drag range from C1 of 0.56 Airfoils at Low Speeds particular. But this is not the aerodynamicist's "free lunch. Also see: DF101. Also see: SD5060.5411 Airfoil comparision plot: Fig. 12. at Rn's of 60k and lOOk the stall occurs earlier than at the higher Rn 's. the DF101 is a better airfoil. In this sense. Also see: DF101. 12. The maximum lift is reduced.30% Thickness: 11. 11. 12. NACA 2. Instead. 12. DF103.30% • DF103-PT (Fig. it has a remarkable lift range for the 11% thickness. 11. 12.1 at the high end.11) The E193 has been used on a wide variety of designs from hand-launch gliders to large cross-country ships. the DF101 seems to be the best choice of the three. 83010. Furthermore. 11. 83010 Airfoil comparision plot: Fig.00% Camber: 2.9) Contrary to what might be expected. leading to "tip stall" problems." since overall the drag has increased for lift coefficients above 0.1.10 Camber: 2.5411 and better than the CLARK-Y. 12. removing thickness has produced the opposite effect. What makes the E193 a popular choice is not the performance in terms of drag. 11.1 Polar plot: Fig.10) As compared with the DF102. the addition of upper-surface thickness did not change the low-lift characteristics. DF102.9 Thickness: 11. 83021 Digitizer plot: Fig.5411 and not quite as well as the CLARK-Y.00% E193 and El93MOD • E193-PT (Fig. DF103.30% for the DF101).8 Thickness: 11. 10. but the stall behavior has improved over the DF101.5. Still.55% for the CLARK-Y vs 2.30% • DF102-PT (Fig. it may be considered as a multi-task airfoil. DF102.4 Airfoil comparision plot: Figs. At low speed it performs better than the NACA 2. At high speeds it is the equal of the NACA 2. With this range a sailplane will have good wind penetration and thermal performance under most .1 to 1.1. the lift range was extended by 0.2 Polar plot: Fig. There is an associated slight increase in drag. the better low-lift performance of the E205 offers improved wind penetration. 12.12 Polar plot: Fig. the E205 has become one of the most favored RC soaring airfoils.2 when flying at high-speed cruise. 12. 57 . Since this time. Also see: E193. Compare this to the AQUILA. S4233. the B version was the more accurate.12). owes much of its success to the E205.11 Thickness: 10. E205.12) Dale Folkening generated the E193MOD by scaling the ordinates of the E193 by a factor of 1. 10. effectively adding camber and thickness. 11. especially near the leading edge.25.12 Thickness: 11.6. As compared with the E193. 12. 12.5 Airfoil comparision plot: Fig.6 Polar plot: Fig.22% Camber: 3. The SAGITTA.'s than thermal soaring airfoils. Of course an airfoil with the same lift range but with lower drag will always be a better choice. This illustrates the general rule that good multi-task airfoils have the low drag region shifted to lower C. The result increases the lift (as indicated by the change in the zero-lift angle of attack) and the width of the polar. Not surprisingly the drag of the B version is everywhere lower than the A.15% E205 • E205A-PT (Fig. Of the two models constructed. which was designed to compete effectively in multi-task events as well as F3B competition. In terms of the overall accuracy of the E193-PT. 12.57% • E193MOD-PT (Fig.14) The popularity of the E205 grew tremendously with the introduction of the Airtronics SAGITTA standard class sailplane.25.13) • E205B-PT (Fig. E193MOD Digitizer plot: Fig. owing mostly to the added thickness.Chapter 5: Comments on Airfoils weather conditions.85% Camber: 4. but it is also higher there than the E205 is at 0. where not only is the minimum drag at any Rn reached at a C1 of about 0. E387 Digitizer plot: Fig.19. E387. S3021. From the polar it can be seen that the drag curves bunch at the lowest drag point near C1 of 0. Because the lift coefficient of a RC sailplane is about 0. it should be noted that the E193-PT matches the E205 better than the intended E193 (Fig 11. Also see: E205. the E205 is clearly better here than the AQUILA. 10. Version A had a little less aft camber than the true E214. the large aft camber leads to some trouble. RG15. • • • • E214C-PT E214C-PT E214C-PT E214C-PT 3° flap (Fig. it is clear that it not designed like the E193.01% Thickness: 10. 11.14 Lift plot: Fig. positive (downward) flap deflection. 12.) It should be noted that for the E214 this loss in lift at low Rn's does not necessarily hamper the performance of a sailplane using this airfoil. DF101. The large amount of aft camber (sometimes called aft loading) markedly distinguishes the E214 from the others. 12. SD5060.58 Airfoils at Low Speeds Also see: S3021. Since the operating Reynolds numbers of aRC sailplane decreases (along with the speed) as the lift coefficient increases. 10. Two models of the E214 were constructed.18) 0° flap (Fig. the mid-range lift coefficients at low Rn are not used. The additional camber of version B produces a slight increase in lift. E374. 12. E374. owing to the excellent low-drag. 12. however. Aside from these small differences the two models are excellent reproductions of the E214. high-lift capability.16) • E214B-PT (Fig.13. SD7084 Digitizer plot: Figs. (A more thorough discussion is given below in the discussion on tripping the E214.12 Polar plot: Figs. much like that produced by a small. To improve the high speed characteristics full-span flaps are often employed. SD7080. Low drag is.19) -3° flap (Fig. 12.15 Aircraft polar: Fig. While the bubble ramp is good for performance. E193.17) By inspection of the E214 contour.0. This loss in lift is due to a large. 10. Notice that at lift coefficients below 1. the lift at a given angle of attack d~creases as the Rn is reduced. The E214 is used primarily for thermal-duration flying. E205.7. desirable at high lift and low Rn.20) -6° flap (Fig. and version B a little more. 12.48% E214 • E214A-PT (Fig. 12. On a more subtle level. 12. and this the E214 has. or E387.9 Camber: 3. uppersurface separation over the aft end of the airfoil.21) . 12. the E214 is designed with a laminar separation bubble ramp-a mild pressure recovery on the aft upper surface. 5.8 Airfoil comparision plot: Fig. was found necessary in order to achieve the indicated drag reduction. Tests were performed for flap settings of 3°.17% of chord. By tripping the boundary layer. probably because the flap was set at a small. The height.s. a two-dimensional trip strip or turbulator placed on the upper surface at 20% chord leads to substantial improvements in the E214's performance. For some airfoils tripping the flow either had no effect at any Reynolds number or actually increased the drag.) As will be discussed later (see SD7090) the height of the trip is an important parameter. This decreases drag provided that the trip drag is smaller than the bubble drag. the -3° flap deflection leads to improved high-speed performance by shifting the polar downward. such as the MILEY. The flap setting was 0 accurate to within • It is instructive to start with the 0° case. In some areas the drag is reduced.t.19.Chapter 5: Comments on Airfoils A 22% flap with a lower-surface tape hinge was added to the B model (see Fig. ±! • E214C-PT u. there are no detrimental effects on performance. 0°. for the -6° case nothing is gained over the.004 in) step of the flap seal.17%. In short.2. did a trip reduce the drag at 300k. (It can be reasonably assumed that at 60k a similar performance gain takes place. Interestingly. There is little change in the drag. For the 3° setting. It appears that at 300k the longer length of turbulent flow behind the trip generates more drag than the untripped bubble. On the other hand. but the increase in maximum lift is hardly noticeable. In fact. 12. For the E214 we recommend that trips be used only for Rn's below 200k.0% (Fig. the lift curve is shifted upward by about 0.03% (0. the tripped E214 offers insight into this phenomenon. xjc = 20%. a small step at this location typically has little influence on the drag. At the location of the hinge line the upper-surface boundary layer is turbulent and therefore insensitive to the small. covering a range typically used in practice.hjc = .wjc = 1. 5.22) For Rn 's less than 200k. transition to turbulent flow takes place sooner than without tripping. -3° and -6°. A trip or turbulator is not a panacea. 200k is the break-even point for most of the airfoils tested.4 for flap configuration). its purpose is to promote transition which shortens the separation bubble. the boundary layer approaches the trailing 59 . Although the boundary layer on the lower surface may be laminar at the location of the flap hinge at high angles of attack. Note that the lift range is shifted downward slightly. negative angle rather than exactly zero. The advantage is seen by comparing the lOOk data with Fig. As a result. for the E214 positive flap deflections do not lead to improved thermalling performance.-3° deflection. while in others it is increased. Recalling that for the untripped E214 at a given angle of attack the lift decreases with decreasing Rn. 12. only in a few cases.0. Compared with the unflapped version. 0. type A (Fig. It is a widely held view that three. Two models of the E374 were produced for these tests. 12. Also see: SD7032.24) • E374B-PT (Fig.s. Below C1 of 0.03% Thickness: 11. Joe Wurts flew the E374 to a new cross country distance record of 141 miles. the A version has slightly more aft camber and is slightly thinner rearward of 50% chord than the B.22). E193. 12. the trip reduces not only the size of the separation bubble but also the extent of the turbulent trailing-edge separation. and E387. the flow can stay attached and maintain the lift (compare Figs. Recently.60 Airfoils at Low Speeds edge with more energy to negotiate the sharp pressure recovery over the aft 5% or so of the airfoil. there is no benefit.10 Polar plot: Figs. 12. 82091.9%) allows it to be used for the large spans found in cross-country sailplanes. • E374B-PT u. AQUILA Digitizer plot: Figs. 12.27) On the more accurate E374B model.9. 10. the E374 shows the effects of a laminar separation bubble through the mid-lift range at the lower Rn's. 10.23 Camber: 4. and because its thickness (10. however. two kinds of trips were tested: uppersurface. Only below C 1 of 0.dimensional boundary layer disturbances produced by bumps or zig-zag tape are more unstable than the initially two-dimensional disturbance produced by a continuous trip strip 28 • Because the trips were placed at different locations.35 does the lift at lOOk begin to depart from the other curves.hjc = . probably because of separation on the lower surface which effectively decambers the airfoil. With the trip. comparisons of their relative effectiveness cannot be made. bumps xjc =50%.0% (Fig. For the bumps. even at the lower C 1 's where it would be thought that the additional camber would be detrimental.10% E374 • E374A-PT (Fig. xjc = 20%. both models were very accurate. the A version is marginally the better of the two in terms of performance. for example.5.22 Lift plot: Fig.s.16-12. For reasons that cannot be satisfactorily explained. Most likely the leading edge of .19 and 12. the 150k case is most interesting. SD7043. What can clearly be seen from these results is that a trip strip at 20% leads to improved performanc~ for Rn less than 200k. As with the E205. 12.5 the bubble is effectively tripped.t.25) The E374 is often selected because its low camber (2. Overall. three-dimensional bumps (type A) and the more commonly tested twodimensional trip strip.2%) makes it more of a high-speed type than a floater.17%. In summary.wjc = 1. above 0. 12.26) • E374B-PT u. 12. All edges were carefully smoothed. 12. was on the order of 7% with a maximum height of 0. upper-surface waviness near the leading edge.hjc = 0. say.24% Thickness: 10.0% (Fig. 12. No attempt was made to untwist the model for the repeated runs or for the tripped and high-turbulence runs.12 Polar plot: Figs.s. xjc = 20%. • E387A-PT u. the results do provide clues to potential performance losses due to. 61 . Interestly. Also see: SD6060. 12.24-12. • E374B-PT u. hasty repairs near the leading edge. at low-lift coefficients the disagreement is larger than we found elsewhere.11. However. 12.30 Camber: 2.31) • E387 A-PT repeated (Fig. which makes it ineffective because it is in the recirculating region of the bubble. there is a small drag increment everywhere.32) was taken near the end. 12. In general the agreement is good. As with the DAE51. as Fig.2.h/c = .s.33 shows that the addition of an upper-surface trip at 20% lowers the drag over the entire range except at high lift and at 300k. xjc = 0% to 15%. 12.91% E387 • E387 A-PT (Fig. there are no large adverse effects. CLARK-Y Digitizer plot: Figs. • E374B-PT thickened trailing edge (Fig.5411. 12. The first data set (Fig.31) was taken very early in the schedule and the second (Fig. which was random. indicating once again that a thin trailing edge is better. The discrepancy may be related to a spanwise twist in the model.17%. The wavelength. It also served as one of two calibration models used to check for measurement repeatability over the course of the six-month testing schedule.Chapter 5: Comments on Airfoils the bubble is upstream of the trip. 12. SD7090.wjc = 1. E205.33) Figure 12. 10.2%.t. The shape is given in Fig.32) The E387 A-PT built by Bob Champine was used to compare our data with E387 data taken in other facilities. Although the waviness was greater than anything expected in typical construction. wavy clay. a generous amount of artist's clay was smeared over the forward 15%.28 shows. 5. NACA 2.20% (Fig. 10. the twist was removed by applying an opposite twist over a two day period. 12.29 Lift plot: Fig.29) The effect of a thickened trailing edge was also measured.28) To investigate the sensitivity in performance to gross. 12. For the first run. for example. While it may be possible in some cases to improve an airfoil by modifying its shape. Since the camber of this model is substantially less than that of the prototype. Comparing the data with the symmetrical SD8020. As expected. any modification will usually degrade its performance. as this model illustrates. 12. 12.80% Thickness: 9. No turbulence measurements were taken under these conditions as this was done merely to compare the data with that taken under the normal. . From both theory and many previous measurements it is known that for a fiat plate developing lift.37) This airfoil was the only one machined from solid metal. there was no elevator deflection. 12. This promotes a rapid transition to turbulent flow.13. the data served to check several aspects of the instrumentation.31-12. An inaccurate copy of an airfoil will probably share few traits with the parent section. 12. a screen was placed in the test section 3 ft upstream of the model leading edge. only angle of attack changes. • E387B-PT (Fig. Also see: E193. 12.34) To test the E387 under high turbulence conditions. 10.62 Airfoils at Low Speeds • E387 A-PT high turbulence (Fig. as it is here. The independence of drag with Rn and the relatively high value of the drag are commonly found when the flow is turbulent. SD6080. In other words.36 Lift plot: Fig.35 Camber: 3. large differences in performance can been expected and were measured. SD7037. S4061. the recovery pressure gradients on the suction side are both early and steep.34. the lift and drag coefficients are practically the same over the Rn range tested. The polar data is representative of a full-flying stabilator. low-turbulence test conditions. S2091 Digitizer plot: Figs. had a rounded leading edge.06% Flat Plate • Flat Plate-PT (Fig.36) The E387 version B data provides some measure of the need for accuracy. As shown in the polars. Since this model was completely symmetrical. 12. 10. shows that the drag of the fiat plate is considerably higher. The geometry of the fiat plate is typical of that used for slab. and tlie aft 3 in were tapered to a :{2 in trailing edge. the high turbulence leads to lower drag because the turbulence shortens the bubble in a way much like that produced by the upper-surface trip.14 Polar plot: Figs. sheet-balsa stabilizers. Consequently we recommend a more streamlined airfoil like the SD8020 for stabilizers. clearly if the airfoil is a good one to begin with. It was ~ in thick. search and rescue. long endurance. MB253515. 10. The FX63-137 was first tested without Monokote covering (version A) and then with covering (B). etc.39 Thickness: 9. 12.08% Camber: 0.16 Polar plot: Figs. Despite these qualities. As it came from the builder. Also see: SD7037. 12. Because of the strong interest in the FX63-137. the FX60-100 has not received wide acclaim.4). for example: long-duration reconnaissance.41 Camber: 5. 12. SD6080.55% FX63-137 • FX63-137 A-PT (Fig.40) • FX63-137B-PT (Fig. It will later be shown (see SD7032) that the differences between smooth and rough surfaces can have measurable effects on performance by influencing transition. 12. and the width of the low-drag portion of the polar is similar to other 10% sections. 12.37 Lift plot: Fig. SD7043. meteorology and mapping. it has been widely tested in other facilities. thin trailing edge presents more of a construction problem than is warranted by the good performance. surveillance. Or perhaps it was simply overlooked when the E214 became popular.15 Polar plot: Fig. Also see: SD7062. the model was sheeted with heavy grain obeche which was varnished smooth.38 Thickness: 2. however. unmanned aircraft 24 • Such aircraft are envisioned for missions where manned platforms are impractical.59% 63 . 12. 12. Possibly the unusually prolonged. 12.41) The high-lift. there are no signs of excessive bubble drag.00% FX60-100 • FX60-100-PT (Fig. relay of radio and television. the grain was not completely filled by the varnish.39) From a quick glance the FX60-100 looks like a good airfoil.40. With the Monokote covering the wood grain was effectively filled. 13. At Princeton the FX63-137 was used to make comparisons with the other tunnels (see Section 2. 10.97% Camber: 3.Chapter 5: Comments on Airfoils Also see: SD8020 Polar plot: Fig. 54233. E387 Digitizer plot: Fig. but this had little effect on the airfoil performance.6% thick Wortmann FX63-137 is often selected in feasibility studies of high altitude.94% Thickness: 13. SPICA Digitizer plot: Fig. who collaborate in the design of airfoils for full-scale competition sailplanes (for example.6.t. and 11. as they are with the NACA sections. HQ2/9 • • • • HQ2/9A-PT (Fig. these differences are of the same order as the differences between the nominal airfoils and the models of the HQ2/9. To illustrate. wing/body interference drag.44) HQ2/9A-PT l. And even though the airfoils were not . one arrives at the conclusion that the nominal airfoils were not really tested. the upper-surface velocity distribution has a gradual pressure recovery region which lowers drag. the Ventus. 12.4 compare the nominal HQ2/9 with the nominal RG15 and S2048. As for the actual airfoils tested. Another important point. 11. is that the performance differences between these models are probably not meaningful (at least as they apply to RC sailplane performance) despite the fact that some differences in the overall shape and appearance of the polars are apparent. fuselage drag. 12. Finally. it should be noted that the airfoil designation "HQ" is also used by Horstmann and Quast. xjc = 40%. xjc = 20%. one very important point needs to be made. Looking at the figures.64 Airfoils at Low Speeds HELMUT QUABECK AIRFOILS (HQ) The HQ-series of airfoils are apparently generated in a manner similar to the NACA four. xjc = 50%. Figs.wjc = 1. and ASH-25).0% (Fig.-the subtle differences between similar polars tend to be lost. Figs. In this method. ASW-22.17%.wjc = 1.43) HQ2/9A-PT u.5.0% (Fig. rather a group of very similar airfoils was tested instead.7 compare the HQ2/9B-PT with the HQ2/9A-PT. and S2048 are small.0% (Fig.hjc = . which has been widely published in the European model press. which immediately follows. 12. 11. RG15.42) HQ2/9A-PT u. Quabeck's design philosophy. When all the other components of drag are factored into the performance of the sailplane-induced drag.t.s. RG15-PT. the thickness distribution is wrapped around a camber line. concentrates largely on the effects of the camber distribution on the airfoil pitching moment 29 .and five-digit airfoils. In our opinion this feature is the principal key to the success of Quabeck's more popular sections. 12. 11. and 82048-PT. For some of them.hjc = . The differences between the nominal HQ2/9. Moreover. based on the lack of smoothness of the coordinates it appears that neither the thickness distribution nor the camber are analytical functions.45) Before discussing the lift-drag characteristics of the HQ2/9.hjc = . and S2048 actually tested.s.t.17%. Although it is not often mentioned. Although no details of Quabeck's method have been given. The airfoils themselves are characterized by a fair amount of aft loading.wjc = 1. etc.s. RG15.3 and 11.17%. Quabeck has been guided by his prolific building and expert flying skills. This is the approach taken by Quabeck. • HQ2/9B-PT u. while the Eppler airfoils (E205. this type of airfoil. The polar taken on the plain A model is shown in Fig. xfc = 50%. 12. • HQ2/9B-PT (Fig.hfc = . The first model. and only slightly hurts the performance at 300k.45 shows the effect of a trip at 50% chord on the lower surface. the drag has increased around C1 ·of 0. With a tapered wing. the upper surface pressure gradient is rather gradual. version A.t. If anything.17%. it is generally advisable to use trips on any portion of the wing that usually operates at Rn less than 150k-200k. was found to be inaccurate and was later modified to become version B. is intended for use with flaps in order to achieve a wide lift range.untripped HQ2/9B-PT has performance much like the (less accurate) HQ2/9A-PT.000 (Fig. 12. especially when compared with the Eppler airfoils. since the flaps will be used to recover the range. 12. therefore. If an airfoil is intended to be used with flaps then a certain amount of lift range (at any given position of the flap) can be traded for lower drag.45 show data taken with trips. 12.s. Rn = 200. blowing xfc =50%.0% (Fig. 12. Trips were placed on the upper surface at 20% and 40% chord.5. As depicted in Fig. 2. An upper surface trip at 50% improves performance below Rn of 200k. while Figs. which is popular in F3B type flying.49) • HQ2/9B-PT trips. 7. Two versions of the HQ2/9 were tested. Even though tripping the boundary layer with a trip strip improved the performance of the airfoil (Fig.50) The A/B model was hollow so that tests involving boundary layer blowing through a row of spanwise holes could be conducted. 5. For the HQ2/9. tripping by blowing degraded it.49.47) at low Rn. even with a flap.wfc = 1.47) • HQ2/9B-PT u. there is an improvement only for Rn's of 150k or less. 12. 12. E193 type) are designed for a wide speed range without flaps. A complete polar for the type B blowing is shown in Fig. For the 20% case shown in Fig. 65 .Chapter 5: Comments on Airfoils tested with flaps (though flaps are usually used) it is doubtful that any one of them would have a decided advantage.8 ram air from a single "total head tube" fixed to the lower surface of the model was used for the air supply. Perhaps the amount of blowing was too large.43-12.42. that is. First.48) The . which improves the management of the laminar separation bubble. since the tip operates at a lower Rn. Figure 12.s. 12. The blowing configuration is shown in Fig. The second reason was described in the overview of the HQ-series airfoils.44 the break-even point seems to be between 150k and 200k. There are two reasons for this: 1. 5. type B (Fig. 12. What stands out is that the drag is quite low. the trip location near the root should be further aft (in percent of chord) than it is at the tip. 12.43. For the 40% case shown in Fig. In this regard.66 Airfoils at Low Speeds Figure 12.52) The J5012 is a 12% thick. 10. the . This "deadband" is a common characteristic of symmetric airfoils. e. At Rn of 60k near zero angle of attack.50. Consequently a stabilator deflection around zero lift produces little response. S2055.g.18 Airfoil comparision plot: Figs. The relatively high drag of this 15% section leaves much to be desired. the aerodynamic characteristics are not quite symmetric about zero angle of attack because. 11. shows that all of the trips are worse than the untripped case at 200k. as well as some other symmetric airfoils (NACA .. the lift curve is nearly fiat. symmetric section designed by Jef Raskin. which compares the blowing data with the cases with and without trips.17.00% MB253515 • MB253515-PT (Fig.00% Camber: 0.0009. 12. with the wing operating very near the maximum lift coefficient.3-11.53) The MB253515 (designed by Michael Bame) was one of the most intriguing of all the airfoils tested.50 Lift plot: Figs. 12. the airfoil is a candidate for use on tail surfaces.45.47-12. however.51 Thickness: 8. Beside this application. NACA 64A010 Digitizer plot: Fig.46. 12. Since this model arrived late in the experiments. S3021 Digitizer plot: Figs. has an undesirable characteristic if it is to be used as a full-flying surface. The airfoil is intended for aerobatic slope soaring where inverted flying is as important as normal flight. SD8000. although it is not always present (compare the SD8020). the J5012.97% Camber: 1. 10. 12. NACA 64A010). Under most types of flying conditions the RC sailplane spends considerable time climbing in thermals. Also see: RG15.19 Polar plot: Fig. NACA 0009. 10. nevertheless. Unfortunately for most airfoils. 12. SD2083.19. 10.99% J5012 • J5012-PT (Fig. 12. a stabilator. Also see: SD8020. It may be that the attraction has more to do with the lift characteristics than the drag. as shown in Fig. just beyond C1~ . S2030. the airfoil profile itself is not symmetric.42-12. As an aside. it is favored by some.52 Thickness: 12.7 Polar plot: Figs. 12. extensive testing was not possible. S2048. 55.s. and possibly on the entire wing.11%. then becomes more typical above 3°. Usually the lift increases smoothly with angle of attack and finally breaks away. the lift decreases with Rn. Note also that as the Rn is decreased the lift characteristics change markedly. two important observations can be made.0% (Fig. which indicates the presence of a large bubble or large trailing edge separation or both. posing handling problems. In a thermal the turbulence is quite high and.hjc = .53.) In Fig. followed by a dip and a long plateau which is maintained down to as low as 30k. the lift characteristics of the MB253515 are shown for Rn from 30k to lOOk. First. the turbulent conditions can cause portions of the wing to stall intermittently. The second observation is the well-defined hysteresis loop in lift which is clear indication of laminar separation. Also see: S4233. a trip should be used at least on the wing tips. The highly desirable stall characteristics of the MB253515 may explain why it is favored by some flyers. suggesting that the airfoil could be improved with a trip. El93MOD. Referring to Fig. This characteristic of the MB253515 separates it from most low-Rn airfoils.54) Figure 12. SD7062. 12. When separation begins to take place. with a stall following shortly thereafter.ck margin gives the MB253515 section a docile feel in thermals and ultimately helps the thermaling efficiency.Chapter 5: Comments on Airfoils airfoil stalls. (See also Section 5.11 67 .t. below the angle of attack at Cz~·· (a ~ 10°). 10.wjc = 1. Some type of separation may also be deduced from the high drag below Cz~·· shown in Fig. 12. therefore. the lift increases rapidly between -2° and 0°. Note that the stall angle of attack is at least 18°-very far from the thermal operating point. • MB253515-PT u. depending on the expected speed range.55 for Rn = lOOk. not through low drag but through handling-by decreasing the work load of the pilot.2. xjc = 20%. 90k and 40k cases. 12. the airfoil may make up for this deficiency with good handling. WB140/35/FB Digitizer plot: Fig. WB135/35. flattens between 0° and 3°. The problem is further aggravated by the lower tip chord Rn 's because with most airfoils the stall angle of attack decreases with Rn.20 Airfoil comparision plot: Fig. for a sailplane operating close to its stall angle of attack. maximum lift ( Cz~·· = 1. For airplanes using this airfoil. Although the drag characteristics of the MB253515 are hardly dazzling. 12.54 shows that a trip at 20% chord improves the performance for Rn less than 200k. The net effect of the local stalling and tendency to tip stall makes efficient thermalling difficult. 11. Hence the relatively high drag can be attributed to a laminar separation bubble. Comparing the lOOk.0) is reached at 10°. This large angle of atta. 58) (A-trip) hjc = . designed for Rn's greater than 600k. The exact mechanism of these performance differences and how it relates to the shape of the trip is still not well understood. As for the bumps (D-trip). from Mississippi State University with a thesis titled "Analysis of the Design of Airfoil Sections for Low Reynolds Numbers. 12. there has been no consensus as to how the term low Rn's should be defined.wjc =. 12. type A (Fig. A simple way would be to state a cut-off Reynolds number number below which everything is considered low Rn's. the A-trip works the best." His thesis culminated with an example airfoil. type B (D-trip) bumps An upper-surface trip on the MILEY airfoil at 31% produces a dramatic drag reduction-up to 73% at Rn = 200k. and second. laminar and turbulent separation were assumed negligible and were ignored in the anaylsis.wjc =.17%. Miley 30 received his Ph. • M06-13-128-PT u. 12.68 Airfoils at Low Speeds Polar plot: Figs. they do not work as well as the others at high lift. with the zig-zag tape ( C-trip) being next.57) • M06-13-128-PT u. But low Rn's usually implies laminar separation and high drag. For all the cases shown. . 12. although elsewhere they are as effective as the A. As the title may suggest.53.56. with this definition the Rn would be different for each airfoil. there is some degree of ambiguity.and B-trips. The B-trip is apparently not high enough (half that of A-trip).55 Camber: 2. Consequently we do not recommend it for model sailplanes. the M06-13-128. the drag between the stall limits was the highest of any airfoil tested. At all Rn's the MILEY airfoil has a high-drag laminar separation bubble as indicated by the large bulge in the polar. 12. However. 12. as the drag is slightly greater. xjc = 31%.52% (B-trip) hjc = . As with any definition.96% M06-13-128 (MILEY) In 1972. Thus another definition is Rn' s where laminar separation is significant.43% Thickness: 14.s. In fact for a Rn of 200k. 12. and it may simply be best to state first what Rn's are being considered.D.52% (C-trip) zig-zag tape.Rn = 200.J.000 (Fig. In Miley's case.s. • M06-13-128-PT (Fig. trips.08%.54 Lift plot: Fig. the MILEY airfoil was tested below its design Rn of 600k. whether or not laminar separation is included in the particular problem under study. bumps xjc = 31%.56) As shown in Fig. S. there is a dead band in lift about 0°.56-12. 12.81% NACA 0009 • NACA 0009-PT (Fig.5411-PT (Fig. It is interesting to note.22 Polar plot: Fig.61 the deadband for the NACA 0009 is worse than the J5012. thin bubble forms on both the upper and lower surfaces. as with the J5012. FX63-137. a long. Also see: SD8020. J5012 Digitizer plot: Fig.61 Camber: 0.5411 is a four-digit NACA section with 2. Also.62) The NACA 2. the lift characteristics are slightly asymmetric. however.00% NACA 2.63.16% Thickness: 12. Again. As expected.60 Lift plot: Fig. Unless 69 .58 Lift plot: Fig.22 reveals. 12.5% maximum camber at 40% of chord. In fact. 12. However as shown in Fig. SD7062 Digitizer plot: Fig. As mentioned in the discussion of the MB253515.59 Camber: 5. as it provides warning that stall is imminent.00% Thickness: 9. a lift plateau with the associated high drag is desirable. NACA 64A010. The drag polar shows that the overall drag is low. 12.21 Polar plot: Figs. There is a smooth.60) The NACA four-digit symmetric airfoils are often used on tail surfaces. the drag of the 9% thick NACA 0009 is lower than the 12% symmetric J 5012. between approximately -2° and 2°. Flight near maximum lift with this section would be exceedingly difficult because there is virtually no angle of attack margin between maximum lift and stall. 12. Any slight asymmetries present (either in contour or surface finish) will alter the boundary layer behavior. most commonly the stabilizer. and this gives a good first impression. 12. but as shown in Figs.5411 • NACA 2. and 11% thickness. 10. continuous increase in lift up to a sharp and abrupt stall. As mentioned before. that the lift range of the thinner NACA 0009 is on a par with the J5012. as Fig. 10. thereby affecting the symmetry of lift. 12. the stall characteristics are undesirable. 12.Chapter 5: Comments on Airfoils Also see: SD7003. any substantial nonlinearity in the lift-curve slope is caused by some type of laminar or turbulent separation. the NACA 0009-PT is not perfectly symmetric which leads to the asymmetric lift and drag characteristics about 0°. For the NACA 0009 at low Rn's. 10. Figure 12.00% NACA 6409 • NACA 6409-PT (Fig. The shape of the polar.62 Lift plot: Fig. is caused by a long run of a favorable pressure gradient followed by a steep recovery region. 12.025 in (0. with the large decrease in drag just before the stall. This does not mean that the separation has vanished. SD5060 Digitizer plot: Fig. 10. 12.5411 and closely related sections (NACA 2412. The actual wind tunnel section was the only model that had open bay construction from leading to trailing edge. 80k and 60k.64) Compared to the two symmetric airfoils discussed so far (J5012. the same line of reasoning cannot be applied at low Rn's. J5012 Digitizer plot: Fig. NACA 0009).2%) worst case. and generally much . NACA 2415 9 ) are not recommended for slow.23 Polar plot: Fig. giving an open-bay cell aspect ratio of 1:4. While this design approach works for high Rn's. 12.65 shows the lift characteristics for Rn's of lOOk. 12.50% Thickness: 11. the intention was to achieve long runs of laminar flow for low drag.64 Lift plot: Fig. CLARK-Y.00% NACA 64A010 • NACA 64A010-PT (Fig. its performance at high speed.66) The NACA 6409 is considered more of a free-flight airfoil than one for RC soaring. and stall strips make the thermalling performance at least acceptable. The sagging of the covering was about 0. such as is needed in windy conditions. DFlOl.63 Camber: 2. 10. the separation on the upper and lower surfaces apparently have the combined effect of cancelling each other out. It is interesting that for Rn' s of lOOk and 80k the nonlinearities almost vanish.24 Polar plot: Fig. Also see: NACA 0009. Also see: E374.65 Camber: 0. 12. SD8020.70 Airfoils at Low Speeds measures such as stall strips on the leading edge are used to alleviate this problem. is very good.00% Thickness: 10. The ribs were ~ in (1% chord) thick and had a spacing of 3 in (25%). instead. However. the 10% thick NACA 64A010 is the worst for low-Rn applications. the NACA 2. In the design of this airfoil. At 60k the presence of a long bubble can be inferred from the nonlinearities around zero angle of attack. 12. thermal duration flying. producing more drag than that found for the lower Rn's (150k in this case).0% (Fig. 12. At 40% chord.00% Thickness: 9.17%. It is instructive to begin with the trip at 70% chord shown in Fig. In other words. much like the AQUILA airfoil. w lc = 1. xl c = 20%. 12. but. the premature stall is not found. can trips still make modest improvements in performance? Note that the question is specific to this airfoil and cannot be generalized. 12.25 Polar plot: Fig.17%. The location will depend on the average local chord Rn.71) RG15-PT u. wlc = 1. As observed before. w I c = 1. hie= .t. At this location. 82091. 12. 70) RG15-PT u. which shows lift data only.t.68) was relatively good. the airfoil stalls at 8° at 60k. the NACA 6409 is an excellent low speed. 71 .68) RG15-PT u. 12. xlc = 70%.0% (Fig. At 20% chord. xl c = 40%. 12.72. But the drag at lOOk has decreased significantly from the untripped case as the trip moves forward.00% RG15 • • • • • RG15-PT (Fig. a trip should be employed for Rn less than 150k-200k. data was taken only up to 200k.005 to 0. hi c = . xlc = 60%. 12.72) The RG 15 was tested extensively with trips because the performance of the untripped model (shown in Fig. Due to the lack of torsional rigidity.s. the large camber severely limits the high speed performance.s.015 in. At 60% chord. the higher-Rn polars overlap since the boundary layer is tripped too soon. it has virtually no effect since it is downstream of laminar separation.t. there is even more overlap at 300k. The lack of repeatability may have been caused by the gusty weather at the time of the lift run. 12. Also see: AQUILA.0% (Fig. hi c = . it is either inside the laminar separation bubble or immersed in the turbulent boundary layer. 12. 12. hie= .69) RG15-PT u. 12. the trip efficiently trips the boundary layer for Rn of 150k at the lower lift coefficients.66 Lift plot: Fig.67.0% (Fig. in Fig. As Fig. 10.66 shows. SD7043 Digitizer plot: Fig. floater airfoil. however. As for the performance.s.s.17%.t.67 Camber: 6.17%. w I c = 1. We asked ourselves: when one starts with a "good" airfoil (the RG15-PT). As is true with the HQ2I9. the trip causes the 150k and 200k curves to spread apart.Chapter 5: Comments on Airfoils less-0. more gradual pressure gradients). trips (Fig. 12. however. S2048. A common feature of most of the Selig airfoils is a long.26 Airfoil comparision plot: Figs.S. This feature usually gives the polars an appearance more closely resembling those measured at high Rn. Implicit in this approach is the recognition that low-Rn airfoil design is still in the early stages of development. 11.75) • S2048-PT misc. and how best to achieve this are topics of research. 12.72 Airfoils at Low Speeds Also see: HQ2/9. Wisconsin and has been used on the Synergy F3B sailplanes that have represented the U. then to determine from these comparisons what factors (in the velocity distribution and boundary layer development) were needed to produce good.e.3.68-12. SD8000. 12. 10. leading ultimately to a better understanding of low-Rn airfoils. at the World F3B Championships held in recent years (1987 and 1989). even today the details of how "gradual" and how "slowly". low-Rn airfoils. the upper-surface transition point moves forward slowly and continuously with increasing angle of attack. Selig uses four digits in the airfoil name. gradual pressure recovery region on the upper surface called a bubble ramp. The airfoil is a "redesigned" HQ2/9 with slightly longer bubble ramps on the upper and lower surfaces (i. 11. These are general design principles. 12. These factors were subsequently incorporated into the design of several new airfoils for RC sailplanes. The $-designation is also used by Somers. Because the performance improvement due to changing these ramps was expected to be . S2048 • S2048-PT (Fig.74) • S2048-PT with trips (Fig.73 Camber: 1. 76) The S2048 was originally presented at the 1985 MARCS Symposium held in Madison. SD2030 Digitizer plot: Fig. As a result of this ramp. Surprisingly.92% SELIG AIRFOILS (S) The original group of Selig airfoils 23 was designed with the aid of both the Eppler and Somers Computer Code for the Design and Analysis of Low-Speed Airfoils 20 and the experimental results of Althaus 9 • The approach taken was to compare the theoretical and experimental results of several airfoils. This is because the airfoils were designed to have a variety of characteristics in the hope that trends ~ould emerge. while Somers uses only three.72 Lift plot: Fig.76% Thickness: 8.6 Polar plot: Figs. 12. The performance of the Selig airfoils is often markedly different from one to the next. 12. The effects of trips on the upper and lower surfaces are shown in Fig.78 Thickness: 7.Chapter 5: Comments on Airfoils small in any case. Nonetheless the possibility of bettering existing F3B airfoil designs may still yield to investigation at another time. 12. 12. 76. And even if it could be measured. Of all the airfoils tested. Also see: 8D2030. HQ2/9. As with the RG15 and HQ2/9.4. 11. 78) The 82055 is a slightly thinned and decambered 52048. 75.76 Lift plot: Fig. trips nearer the leading edge can be expected to produce lower drag at lower Rn's. As with other F3B airfoils. 11.94% 82055 • 82055-PT (Fig. 10. 8D2030 Digitizer plot: Fig. flaps are necessary to fully realize the capabilities of the airfoil. 10. The results of this small perturbation in shape are marginally lower drag and a downward shift of the lift range. Although this is the lowest drag airfoil tested and should do very well on sailplanes where speed is paramount. more accurate models than the ones we tested would be required to detect and measure it. 52048. 12. the drag was increased except when the trip was far aft.74-12. and then it had no effect. 8D8000. Instead. actually realizing or maintaining that difference would be difficult.66% 82091 As mentioned in the discussion of the AQUILA airfoil. The 82091 was designed 73 . none worked. the 52055-PT had the lowest drag at 300k.77 Thickness: 8.28 Polar plot: Fig. 12.7 Polar plot: Figs. RG15. 82055. Although no attempt was made to systematically study the effects of trip location on this airfoil. Also see: HQ2/9. 8D8000 Digitizer plot: Fig.63% Camber: 1. RG15. Although several attempts were made to reduce the drag further through the use of trips. the 8% thickness offers a significant construction challenge. . results of miscellaneous trip locations are shown in Fig. it does seem that a slight improvement in high-speed performance can be obtained through the use of an upper-surface trip around 60-70% chord combined with a lower-surface trip near 60%.99% Camber: 1. its high-speed performance is severely compromised by its high-camber.27 Airfoil comparision plot: Figs. 12. the goal of extending the lowlift end of the polar beyond that of the AQUILA airfoil has been achieved. 12. The effect of the roughness of version A as compared with the smooth version B is to decrease the drag at 60k. Smaller improvements are found for lOOk. Details of how this was achieved through airfoil design may be found in Reference 23. but this comes mostly through the relaxation of the flat-bottom requirement. The first model. For these tests the tab was 0. (Note that there is a small drag penalty associated with the Gurney flap.002 in) brass shim stock with a length of 0. For the type A Gurney flap. Referring to the B version data (Fig. In addition. The model was later re-contoured and given the B designation.80). Undoubtedly this influenced transition in a manner similar to that of a trip strip.83) The S2091B-PT was tested with a so-called Gurney flap.74 Airfoils at Low Speeds primarily to be an improvement over the AQUILA by extending the polar to lower lift. while maintaining the AQUILA's low-speed. thin tab on the order of 1% chord. the leading edge was rough in certain areas from the fiberglass beneath the paint. high-lift characteristics. often used on racing car wings which develop a download to increase traction. but diminishing returns begin to appear for type C. the S2091 is an adv~nce in performance over the AQUILA. version A.6% and 1. In summary. 12. 5.81) • S2091B-PT Gurney Flap type B (Fig.) . Futhermore.82) • S2091B-PT Gurney Flap type C (Fig. ·a Gurney flap is a simple.1 over that of the AQUILA. Of course a flat lower surface does not necessarily mean the airfoil is deficient. 79) • S2091B-PT (Fig. 12. • S2091A-PT (Fig. 12. It can be reasonably expected that the efficiency of the small 0. 12. Only the B version was digitized for coordinates.5.6% for C. which is perpendicular to the lower side of the airfoil at the trailing edge. As can be seen in Fig.2% Gurney flaps in increasing lift is not unique to the S2091. 1. the whole polar is shifted upwards in lift coefficient by 0. while no improvement exists at 200k.1-an impressive result for such a simple modification~ Similar results are found for type B. as illustrated by the performance of the DF101 and the nearly flat-bottom S3021. it comes as some surprise that at Rn's above 60k the maximum lift coefficient is increased by 0.2% for B. • S2091B-PT Gurney Flap type A (Fig. 12.6% chord for type A.017% chord thick (0. and 2. was found to be too thin as verified by hand-held templates.80) Data on two models are shown. E205. Other examples include the flat-bottom 83014. 83014. 10. All these designs took advantage of the increased understanding of the laminar separation bubble.86 Thickness: 10. Although the 83016 has a slight edge over the 83010 and 83014 at high speed and the 83014 looks best at a Rn of 60k.83 Lift plot: Fig. and as a result the high-drag bulges in the middle of the low-Rn polars are virtually gone.89 Lift plot: Fig. 10.09% 75 . although as shown in Fig. and in most places the drag of the 83010 is lower than the E205.31.30 Airfoil comparision plot: Fig.Chapter 5: Comments on Airfoils Also see: AQUILA.87. and 83021.10% Camber: 3. 12. 10.84 Thickness: 10. 12. 12. CLARK-Y. It is interesting to note that in many respects the 83010 performance characteristics are very similar to the DF101. 12.91% 83010 • 83010-PT (Fig.82% 83014 and 83016 • 83014-PT (Fig. SD7037.32 Polar plot: Figs.87) • 83016-PT (Fig.89) The 83010. SD7032 Digitizer plot: Fig. 12. none of the three has a decided advantage over the others.79-12. 12.29 Polar plot: Figs. DF101.32% Camber: 2. Also see: 83021. 83016. Also see: 83021. 12.57% 83016 Thickness: 9. E205 Digitizer plot: Fig. 11. SD7080 Digitizer plot: Figs.8.85) The 83010 is the first example of a Selig airfoil that has a lift range much like the E193 and E205 airfoils.46% Camber: 2. 12.88 83014 Thickness: 9. 11. 10. Instead. the polar and its edges are more rounded. 83014 and 83016 are quite similar aerodynamically since the latter two were derived from the first with only minor modifications. 12.85 Lift plot: Fig. a comparison of the profiles actually tested reveals that the 83010 and DF101 shapes are very different. 12.8 Polar plot: Fig.52% Camber: 2. 83016. E214. 000 (Fig. Also see: E205.9. xj c = 45%.94) The version A model of the 84061 is inaccurate. • 84061B-PT u. 83014. founder of Off the Ground Models.t. 12.s. 12. but. 12.96% Thickness: 9. 12.92 Lift plot: Fig. these differences in the data come as no surpnse. Certainly some of the reduction is because the 83021 is thinner than the E205. 83010. The more accurate version B shows lift and drag characteristics much different than those of the A.14 respectively) and shows that the drag of the S3021A-PT is almost everywhere lower than that of the E205B-PT. SD7080. Rn = 150.97) .s. and a third in Unlimited. it is clear that most of the drag reduction comes from better management of the laminar separation bubble.33. 12. Rn = 150. Paul Carlson. flew his newly designed and kitted PRODIGY sailplane to a first place finish in the 2-Meter class.90. version A is the more accurate one. As was amply demonstrated in other examples. xjc = 45%.t. 12.9 Camber: 2. a second in Standard.4 compares the data from the 83021A-PT and the E205B-PT (Figs. 12. and the popularity of the 84061 has grown proportionally. 8D7084 Digitizer plots: Fig.0% (Fig. w j c = 1. At a recent AMA Nationals {the NATS).t. • S4061A-PT (Fig.000 (Fig.90) • 83021B-PT (Fig. 5.47% 84061 Although originally intended as a cross-country airfoil. 12.s.93) • S4061B-PT (Fig. 5.17%. To date over 3500 PRODIGY kits have been produced. Two models of the 83021 were built. the 84061 has since demonstrated it's versatility under a variety of conditions. 10. and the resulting reduction in drag is dramatic.000 and 300. 12.92) The 83021 has a bubble ramp longer than the transition ramp on the E205. The ISE8 code predicted a similar drag reduction when it was used to compare the nominal airfoils. 4. 12. from the shape of the airplane polar shown in Fig. Fig.90 and 12.34 Airfoil comparision plot: Fig.95) • S4061B-PT u. 10. hj c = .96) • S4061B-PT u.76 Airfoils at Low Speeds 83021 • 83021A-PT (Fig. xjc = 45%. Version B has a refiexed trailing edge which produces a steeper pressure recovery on the upper surface--a change which leads to a higher bubble drag. 12. 11.91 Aircraft polar: Fig.9 Polar plot: Figs. nothing conclusive can be said about the two-dimensional airfoil characteristics. this design change has not lead to improved performance. retired NACA/NASA test pilot. It is interesting to note that this problem is also predicted by the l5ES code which was not available when the 54062 was designed. Also see: SD6080.tes were not digitized because there is no representative section. E387. SD6080.wjc = 1. 12. Also see: S4061.17% reduces drag for Rn less than 150k. 12. 77 . 10.s.98 Camber: 3.90% Thickness: 9.35.93-12. the 54062 has higher drag owing to a larger laminar separation bubble.97 Lift plot: Fig. S4062. To date the 54233 has probably been most used by Bob Champine.100) The thin trailing edge of this model was wavy for the aft 25% chord along the entire span. 12. Because of the significant three-dimensionality.36% S4233 • S4233-PT (Fig. But as the data shows. 77% Camber: 4.95. 5D7037. DAE51 Digitizer plot: Fig.14% Thickness: 9.97 show that a two-dimensional trip produces the same result as zig-zag tape (type A). DAE51. The model coordin<~.53% S4180 • 54180-PT (Fig. His enthusiasm for this airfoil is reflected in the accuracy of the wind tunnel model which he built. 10.36 Polar plot: Figs. The major modification is a longer run of laminar flow designed into the S4062. 12. 10. 12. Also see: S4061 Polar plot: Fig.99 Camber: 4. xjc = 20%.37 Polar plot: Fig. who recently achieved his second L5F level 5 flying his "stretched" GEMINI with the 54233.t. Rather.60% S4062 • S4062-PT (Fig. E214 Digitizer plot: Figs. 12. E193.99) The S4062 was intended to be an improvement over the 54061.101) • 54233-PT u.0% (Fig.102) The 54233 was designed to be an improvement over the MB253515.96 and 12. a trip with height of 0.Chapter 5: Comments on Airfoils As shown in Fig. 12. hjc = .17%. 12. 12.100 Thickness: 9. Figures 12. 102 Lift plot: Fig. Also see: 82055. 10.103 Thickness: 13.78 Airfoils at Low Speeds Note that the 13. Also see: SD2083. Since the lift range is narrow.85%) for good. 83021 Digitizer plot: Fig. so one would expect a trip to improve matters.38 Polar plot: Figs. As with the MB253515. El93MOD. SD2030 • SD2030-PT (Fig. 12. 12.105 Thickness: 8. 12. SD7062 Digitizer plot: Fig. high-speed performance.64% Camber: 3.56% Camber: 2. camber-changing flaps are recommended.85% .25% SD2083 • SD2083-PT (Fig.101. 12. 82055. Fig.96% Camber: 2. and there is too much camber (2.106 Thickness: 8. 12. 10. The airfoil suffers from a fairly large separation bubble. 10.104 Lift plot: Fig. SD8000.102 shows that the low-Rn drag can indeed be reduced using a trip. HQ2/9 Digitizer plot: Fig. WB135/35. the 84233 shows a bulge in the drag at a Rn of lOOk. but the performance with the trip is slightly worse at Rn of 300k. Also see: MB253515.40 Polar plot: Fig. RG15.104) The SD2030 was designed for operation at speeds higher than those typically found on most RC sailplanes. 12. SD2030. For this reason it is suitable for high-wing-loading F3B and F3E type aircraft.26% SELIG AND DONOVAN AIRFOILS (SD) A discussion of the design philosophy used behind the SD designs is given in Chapter 4.106) The SD2083 was an early and not very successful attempt at an F3B design. 12.39 Polar plot: Fig. 12.6% thick 84233 has a wider lift range and lower drag than the 15% thick MB253515. 12. Chapter 5: Comments on Airfoils SD5060 • 8D5060-PT (Fig. especially at the high Rn's. which are important for cross-country flying.s. wjc = 1.41 Polar plot: Fig. 12. • 8D6060-PT u. one can see the advantages of the 8D6060.112 Thickness: 10.t. In fact the airfoil stalls abruptly.45% Camber: 2. although at 200k and 300k the drag is increased. 12. 8D7090 Digitizer plot: Fig.h/c = . CLARK-Y Digitizer plot: Fig. 10.110). 12. It is not surprising to find that the 8D6060 and E374 are quite similar when both are tripped at 20% (compare Figs. however.17%. E205. Almost everywhere it has lower drag than the E374.109}.5411.wjc = 1. are not as benign as the 83021.84% 79 . 10. NACA 2. Moving the trip forward to 20% further reduces the drag at lOOk.107) For operation at Rn's from lOOk down (e.g. 12. 12.109-12. 12. h/c = .s. The E374 clearly shows large drag at low Rn's due to bubble losses.25) with the 8D6060 (Fig.109} The 8D6060 was designed to be an improvement over the E374 for crosscountry flying.42 Polar plot: Figs.27 and 12. xjc = 20%. By comparing the E374 (Fig. 12.0% (Fig. the 8D6060 was designed with a longer bubble ramp than the E374. Also see: E374. hand-launch RC sailplanes) the 8D5060 is an improvement over the 83021. although in the same regime the plain 8D6060 has the lowest drag of both airfoils.110) • 8D6060-PT u.0% (Fig. tripped or untripped. To alleviate these effects. 12. DF101.107 Lift plot: Fig.30% SD6060 • 8D6060-PT (Fig. 12. much like the NACA 2.108 Thickness: 9.5411.37% Camber: 1. xjc = 40%.17%.111 Lift plot: Fig. The stall characteristics of the 8D5060.t. 12. Also see: 83021. the E374 has a slight advantage over the 8D6060. At high-Rnjlow-Cl. 12.111) A trip placed at 40% reduces the drag at lOOk by shortening the bubble. s. trips were placed on the SD6080 at 10%. the drag over most of the polar is lower.113 and 12.118) • 8D7003-PT repeated (Fig.s.hfc = . When the polars for the 30% trip and the untripped case are compared (Figs. • 8D608'0-PT thickened trailing edge (Fig. the 8D6080 offers improvements at the low-lift end of the flight envelope.115) • 8D6080-PT u.116). Figures 12. As expected.113) The 8D6080 is an improvement over the 84061.17%.113-12. 10. 12.18% SD7003 • 8D7003-PT (Fig.3. the trip at 30% is the best between Rn' s of lOOk and 300k. xfc = 20%.17%. • 8D6080-PT u. the drag is especially low when compared with all the other airfoils tested. it may be considered to span the entire upper surface. Rather. . It should be noted that the "thickened trailing edge" was slightly different for each of the three airfoils. xfc = 30%.114) • 8D6080-PT u. However this does not mean there is no bubble. the bubble losses are small. xfc = 10%.0% (Fig. 12. the trip farthest forward is best at lower Rn's. h/c = .t.117 Camber: 3. 12. 12. a 30% trip should clearly be used for Rn at and below 150k.wfc = 1.119) The 8D7003 was designed to have a very long and gradual upper-surface bubble ramp.w/c = 1. particularly in those areas which are used by a typical RC sailplane.94 and 12.t. and 30% chord.t. The resulting effect is particularly apparent in the overall smoothness of the polar which shows no trace of high drag due to a laminar separation bubble.17%. 12.h/c = .74% Thickness: 9. Besides this extension.0% (Fig.116.80 Airfoils at Low Speeds SD6080 • 8D6080-PT (Fig. Although the high-lift characteristics of the two airfoils are much alike.0% (Fig. 12. 12.113 show that the lower part of the 8D6080 polar is extended over that of the 84061. At Rn's of 60k and lOOk. 20%. Also see: 84061.117) As observed with the DAE51 and E374. The shape of the trailing edge is shown in Fig. Of the three cases. 8D7032 Digitizer plot: Fig. 8D7037. In fact.114-12. but the effect of increased drag is common to all. the thickened trailing edge increases the overall drag on the order of 5%. 5. E387.s. 12. w/c = 1. 12. 12.43 Polar plot: Figs.116) As shown in Figs. It seems unlikely. type A (Fig. The data show that control of the bubble has been achieved. 12. SD7003-PT u. 12.17%.0% (Fig.124) Several attempts were made to reduce the drag through the use of trips. 12.46% SD7032 The SD7032 is among the best of the thermal-duration type airfoils (such as the AQUILA.127) • SD7032D-PT (Fig.t.s. 12.s. such as with hand-launch sailplanes. The design of this new airfoil incorporates what is presently known about mitigating laminar separation problems through the use of a bubble ramp. as shown in Figs. 12. leaving a relatively rough aerodynamic surface.44-10. but they either increased the drag or had no effect.118-12. S4061. One is tempted to ask whether this means that the SD7003 is optimized for ramps such that no further improvements can be made.s.) • • • • • Also see: SD8000. This needs to be considered primarily for wings operating at very low Rn's. SD6080). At moderate to high Rn's. 12. xjc = 70%. (See Section 5. 12.t.wjc = 1. S3021. MILEY Digitizer plot: Figs.121) SD7003-PT u. bumps xjc = 70%. but it is an interesting question nonetheless.123) SD7003-PT u. The model was then covered in Monokote (version B). Other than a few small discrepancies. but considerably better at Rn of 300k. the untripped SD7032 is about the same at the lower Rn's. 12.s. there is little evidence of separated flow even at Rn of 60k. 12.s.133. 10. the performance is generally better than the E214. S2091. For example.124 Lift plot: Fig.17%. RG15. 12.wjc = 1. 12.51% Camber: 1. xjc = 60%.Chapter 5: Comments on Airfoils The polars shown in Fig. The first was initially covered in balsa only (version A).122) SD7003-PT u. bumps xjc =50%.hjc = . As compared to the E214 with a trip (which is another way to control separation). • SD7032A-PT (Fig.hjc = .125 Thickness: 8. It should be mentioned that the angle of attack at stall decreases with Rn. the agreement is quite good and typical of repeated runs (see the E387 A-PT). and for Rn of 60k the stall is relatively sharp. E214.126) • SD7032B-PT (Fig. and finally a 21% flap 81 .119 are from a second series of runs to provide a measure of the overall repeatability.49 Polar plot: Figs. bumps xjc = 60%.0% (Fig. 12. type A (Fig. type A (Fig.134) There were two test sections of the SD7032 built.3. in contrast to the untripped E214.120) SD7003-PT u. What is also of interest is that the -6° flap deflection is too much of a good thing.95 when compared with the SD7032D-PT. 12.82 Airfoils at Low Speeds was added (version C). As with the E214. more accurate model is version D. 12. 82091. 10. 10.10 Thickness: 9. For C1 of 0. See Fig. 12. As shown in Fig. This is because of the way the wing "flies through" the polars.130) SD7032C-PT -3° flap (Fig.6.126-12.133 Aircraft polar: Fig.126 and 12.128) SD7032C-PT 3° flap (Fig.131) SD7032C-PT -6° flap (Fig. For version B. the drag is decreased for Rn's less than 200k. 12. 12.135 Lift plot: Fig. This is still another illustration of how a relatively small surface difference can affect performance. 12. in this case taking advantage of the low-drag areas.60 < C1 < 0.10. 12.95% Camber: 3. The flap configuration is shown in Fig. The second. 10.hjc = .134. AQUILA. Figs.. Having said that.51 Polar plot: Figs. At the leading edge the added thickness was zero and at the trailing edge it was 2 in. -3° provides lower drag at high Rnflow C. The drag curves for the model A at Rn of 200k and 300k are nearly coincident which indicates premature transition at 300k. given an airfoil that is designed with a bubble ramp. the 300k polar exhibits about a 10% drag increase in the range 0.50. This. 12. Also see: E214.129) SD7032C-PT 0° flap (Fig. • • • • • • SD7032D-PT u. 5.127.t. SD7037 Digitizer plot: Figs. in turn. 12. 12. to give a finite trailing edge thickness.35.wjc = 1. adding the smooth Monokote covering reduces the drag by up to 20% at 300k. note that both the A and B versions out-perform the D version at low to moderate Rn (with only a small penalty at high speed).127 show that the balsa surface acts in a manner similar to that of a trip.50 shows that the A/B/C model has a flat spot on the upper surface near 45% which affects the laminar separation bubble. the flap is intended to be used only in the reflexed position. This model used the nominal coordinates. Fig. 5. 12. This is to be expected.s. 5. xfc = 45%. The nominal SD7032 coordinates for the A/B/C model were modified slightly by the addition of a linear thickness increase from the leading to the trailing edge. This small improvement does not warrant the use of trips for the type of flying typically encountered.17%.132. leads to the differences in drag at 300k.135) A trip at 45% offers only marginal improvements.132) Flaps on the SD7032 extend the useful high-speed range.0% (Fig. 12.66% . £ SD7032C-PT 6° flap (Fig. s. the polar is shifted downward in lift. 10. This airfoil reaches a high lift coefficient which produces a low sink speed if flown just above stall speed. the drag is lower.hjc = .2.140 Thickness: 9.137 Thickness: 9. 10. SD6080.3 offers good 1/D performance which. If so. 12.138. To increase the lift range. 12. lOOk. decambered SD7032. Although it was not measured.S.) Also see: SPICA. but with the caveat that the stall. but flaps in this case are almost essential to improve high-speed. but it is reasonable to assume that the tripped version would maintain most of this high lift while substantially decreasing the drag. should make the SD7037 a popular airfoil. E214 Digitizer plot: Fig. SD7032. For weak thermal conditions common to east coast soaring in the U. which is moderately abrupt.Chapter 5: Comments on Airfoils SD7037 • SD7037-PT (Fig. between-thermal performance. and 150k). flaps would be useful.137) An upper-surface trip at 30% reduces the drag at lower Rn' s only slightly. the SD7037 would make an excellent crosscountry airfoil. As would be expected. E214. xjc = 30%.wjc = 1. xjc = 20%.17%. it was not tested with the trip at the lower Rn 's (60k.0% (Fig. 12. the SD7043 could be a good performer on small-to-medium size airplanes. one would expect the drag at 300k to be increased.11%.51% 83 .t. 12.52 Polar plot: Figs. SD7037-PT u. As with the SD7032. 12. probably needs to be controlled with leading edge stall strips. Due to time limitations.0% (Fig.t. The relatively low drag at Cz near 0. 12.136) The SD7037 is a thinner.53 Polar plot: Figs. o Also see: FX60-100. and the lift range is less. (See Section 5. especially for thermal-duration flying.hjc = ..138) o SD7043-PT u.13% Camber: 3.s. 12. 12.139) The shapes of the untripped SD7043 polars are indicative of a laminar separation bubble that would be amenable to control by a trip.wjc = 1.02% SD7043 • SD7043-PT (Fig. SD7032 Digitizer plot: Fig.136. trips are of little benefit.20% Camber: 3. 12. together with low drag at high lift.139 Lift plot: Fig. As Fig. the 84233 has lower drag. xjc = 15%. 12. At low lift coefficients and high Rn's.0% (Fig. 12.141) • 8D7062-PTu.84 Airfoils at Low Speeds SD7062 • 8D7062-PT (Fig.t. See Fig. its lift range is greater. but when untripped it performs considerably better than the 84233. The primary differences are that the 8D7084 is about 10% thicker. 12.144) The 8D7080 has a lift range much like the 83021 but with slightly lower maximum lift.12.17%. 12.11 Camber: 2.144 Lift plot: Fig. 5.102 and 12.wjc = 1.t. 5.143 Camber: 3. 12.11.17%. 12. the model would have a little wider lift range and slightly higher drag than the nominal. The interesting result is that the two airfoils perform almost identically provided the 8D7084 wing loading is about 10% greater. E205. 10. although not necessary. is recommended to improve thermalling performance. DF101 Digitizer plot: Figs.98% SD7080 • 8D7080-PT (Fig.wjc= 1. 8D5060. 83014. xjc= 15%. 12. If the usual trends apply. In particular.15% SD7084 • SD7084-PT (Fig.56 Polar plot: Fig. and the minimum drag is higher than the 8D7080.97% Thickness: 13. 10.142) • 8D7062-PT u. the situation reverses. MB253515 Digitizer plot: Fig. When both airfoils are tripped (see Figs.48% Thickness: 9.143 shows. the drag is quite low. Also see: FX63-137. E193MOD. 12. This should be considered when building.55.08% and .142).143) The 8D7062 is a thick (14%). A flap. highly cambered (4%) airfoil. There are some significant differences between the actual and nominal airfoils. . 83021. Also see: 8D7084. 12. which will lead to excellent wind penetration. at 200k a trip placed at 15% increases the drag with respect to the untripped case.54 Polar plot: Figs.s. the model was too thick (except at the leading edge) and had a turned-up trailing edge.146) The SD7084 is a slightly modified 8D7080. 10. 84233.h/c = .0% (Fig. except at high Rn.141-12. In some ways it is comparable to the 84233.h/c= .145 Aircraft polar: Fig. although not much.s. 12.11 Camber: 2.04% (0.Chapter 5: Comments on Airfoils Also see: 8D7080. with a wavelength longer than that of the raw balsa grain. 8D6060.148) At first glance the 8D7090 appears to be a lower lift version of the 83021. Rn = 300. A comparison plot of the profiles (see Fig. the 8D7090 has a fairly sharp stall.9 85 .63% SD7090 • 8D7090-PT (Fig. 12. unlike the 83021. E205 Digitizer plot: Fig. 12.005 in) and had only a small effect. 11. is somewhat thicker. 12.151 shows. NACA 2.9) shows that the 8D7090 has less camber than the 83021. After the model was tested in this condition the Monokote was shrunk tightly over the surface sheeting and made to follow the balsa grain by firmly pressing a cloth over the hot Monokote as it cooled.) • 8D7090-PT trips.5411. 10. the covering on the surface was typical of plastic film over balsasomewhat wavy. These tests illustrate how even small roughness elements effect performance.147 Aircraft polar: Fig. and the thickness distribution is significantly different along the forward. 5.149) As received. • 8D7090-PT loose/tight covering.000 (Fig. 12. This should mainly be considered for F3B and cross-country flying where Rn's of 300k and higher are common. (The shift in the angle of attack was caused by an inadvertent sharp jolt to the angle of attack sensor prior to the start of the run.31% Thickness: 9. But as the trip height increased so did the drag.150) At 300k various size trips were tested at 30% chord. especially at low angles of attack where the flow would otherwise have been laminar to almost 60% chord.000 (Fig. As Fig. Rn = 300. 11. 12. CLARK-Y Digitizer plot: Fig. upper surface. 12. 83021.149 shows for a Rn of 300k. The lowest trip was 0. This is caused by a sudden "bursting" of the leading edge separation bubble at high angle of attack. The resulting surface showed a marbled appearance.57 Polar plot: Fig. this difference between the surface finishes did not affect the performance. Also see: E374. 10.58 Airfoil comparision plot: Fig. For the 8D7090 this part of the airfoil plays an important role in the stall characteristics. As Fig.146 Lift plot: Fig. xjc = 20%. SD2030 Digitizer plot: Fig. 12.60 Polar plot: Fig.10).w/c = 1. Also see: J5012.t. i. xjc = 70%. 11.158 that are most interesting. 12. At 60k.157 Lift plot: Fig.155) Although the shape of the SD8000 is quite different from the HQ2/9-RG15S2048 group of airfoils (see Fig.59 Airfoil comparision plot: Fig. There is no significant deadband. even at a Rn of 40k. the J5012.17%.10 Polar plot: Figs. 12. 12.148-12.86% SD8020 • SD8020-PT (Fig.wjc = 1.155 Lift plot: Fig. xjc = 40%. separation at the higher angles of attack limits Cz~ .s. As shown in Fig. making it more like the NACA 64A010. 12.0% (Fig.150 Lift plot: Fig.10% Camber: 0.17%. NACA 64A010 Digitizer plot: Fig.s.157. trips do not produce dramatic improvements at the higher Rn's.151 Thickness: 9. NACA 0009 and NACA 64A010. RG15. (The offset in lift for Rn of 40k was probably caused by operator error when the tare-weight entry was made. But it is the lift characteristics shown in Fig. 12. 12.00% .h/c = . 11. 12. the performance is strikingly similar.71% Thickness: 8. 82048.152-12.hjc = .t.87% SDSOOO • • • • SD8000-PT (Fig.t.156 Camber: 1. The lift range between Rn's of lOOk and 300k is comparable to the J5012. however. 12. 12. 12.e. 82055.152) SD8000-PT u. .153) SD8000-PT u.17%. 10.w/c = 1. At and below 150k.158 Thickness: 10. the lift and drag characteristics compare favorably with the other symmetric sections tested..0% (Fig. there is some advantage to using a trip.154) SD8000-PT u.0% (Fig.86 Airfoils at Low Speeds Polar plot: Figs. As with the RG15 and HQ2/9.) The nearly linear Cz vs a characteristics and acceptably low drag should lead to smooth handling qualities. 12. 10.157) The 10% thick SD8020 is a symmetric airfoil specifically designed for use on tail surfaces at low Rn's. Also see: HQ2/9.s. NACA 0009.h/c = . 12.99% Camber: 1. 161 Thickness: 11.164} These two airfoils were designed by "Woody" Blanchard Jr. however. Consequently the second has more aft thickness which leads to substantially different performance. does have a dragproducing laminar separation bubble at lOOk. 12. 12. however.Chapter 5: Comments on Airfoils SD8040 • SD8040-PT (Fig. 10. 12.162) • WB140/35/FB-PT (Fig. Unfortunately. 12. but the SD8000 along with the HQ2/9.159) The SD8040 was an attempt at an F3B type airfoil. Other than for ease of building. the SPICA airfoil has high drag at low lift coefficients. whereas the WB135/35 shows just the opposite trendrelatively low drag at lOOk. He reports that his SPICA sailplane which uses this airfoil has consistently placed high in the standings at thermal-duration contests. The modified version.160) The SPICA airfoil was designed by Chuck Anderson in 1978 as a fiat-bottom.99% Camber: 2.72% Camber: 4. RG15. This is probably a result of the high lift coefficients produced and the airfoil's gentle stall characteristics. trainer airfoil. fiat-bottom version with the same camber line as the original.159 Thickness: 9. low-level turns in small. 12. turbulent thermals.160 Lift plot: Fig. The first is the original and the second is a modified. 12. and S2048 show better performance. Also see: HQ2/9.61 Polar plot: Fig. It is interesting to note that although the WB135/35 and MB253515 have similar upper-surface contours. S2048 Digitizer plot: Fig. 10. These two qualities combine to improve the chance of successfully completing tight.62 Polar plot: Fig. Also see: SD7043. The MB253515 shows the strong effects of a laminar separation bubble at Rn of lOOk. 12. which would make cross-country flying at high speed difficult. low Reynolds number airfoil design is still far from a precise science. 11. the behavior of the airfoils is quite different (see Fig.11). As for windpenetration. E214. 87 . the fiat bottom does not appear to be advantageous. SD7032 Digitizer plot: Fig.74% WB135/35 and WB140/35/FB • WB135/35-PT (Fig. RG15.65% SPICA • SPICA-PT (Fig. for a typical tapered wing (which closely approximates the "ideal" elliptic planform). To illustrate both good and bad stall characteristics.64 Airfoil comparision plot: Fig. While this airfoil exhibits a reasonable rounding of the lift curve.63 exemplifies the opposite end of the spectrum. A typical example is the 52091B-PT at Rn = lOOk. 12.164 Lift plot: Fig. the sudden drop in C1 is over 0. 12. 12. 12. What are the desirable stall characteristics that one should look for in the shape of the lift curve near stall? First of all. The stall characteristics of the MB253515 at Rn of lOOk are excellent (see Fig. For the MB253515 this angle of attack margin is quite large in comparison to most other airfoils.5411 shown in Fig. 10.88 Airfoils at Low Speeds Also see: MB253515.2 Stall Behavior In the quest for lower drag it is easy to forget the importance of the stall.11 Polar plot: Figs.92% Camber: 3. the MB253515 and NACA 2. Left uncorrected. What is more. "Phase 1"-the rounding of the lift curve-occurs at approximately 10°. The 8° difference provides a broad. serve as examples.55). far more altitude will be lost trying to recover than might have been gained by the marginally lower drag. 11. An airfoil may have a very low drag coefficient.84. The 52091 is an example of the intermediate case. 5D7062. many "full-size" airfoils have similar problems. 12. Fig. Thus the two-dimensional lift data shown in Chapter 12 provides a good indication of what will be experienced in flight.5411 respectively. There are two phases to most stalls. E193MOD. easily flown plateau with only a slight loss of lift from the maximum. the three-dimensional stall will be quite similar to the two-dimensional one.162.3.63.70% 5. there is hysteresis of almost 4°. Of course these problems are not unique to low-Rn airfoils. 12. This airfoil loses lift with no warning whatever (essentially zero angle of attack margin). the angle of attack margin of the NACA 2. 54233. FX63-137 Digitizer plot: Figs. and some exhibit hysteresis as well at the stall. poor stall characteristics seriously compromise the controllability at low speed.75% WB140/35/FB Thickness: 13.53% Camber: 3. 10. Over the years several remedies . On the other hand. The first occurs when the lift developed by the wing no longer increases with angle of attack-the lift curve rounds over. but the actual stall which marks "phase 2" does not occur until at least 18° (the limit of the measurement). even though flight in this region is inefficient because the drag is very high. Many airfoils show intermediate behavior. but if the airplane abruptly tip stalls while thermalling.163 WB135/35 Thickness: 13. The second phase occurs when the lift suddenly falls back to a lower level as the angle of attack increases-the true stall. On this airfoil the pressure recovery region is so gradual that most of the upper surface may be considered a bubble ramp. Flight tests on a model sailplane using a NACA 2512 (which is very similar to the NACA 2. trips were generally effective only below 300k. when placed upstream of this point. the simple two-dimensional trip strip is satisfactory. 5. In summary. and two of these methods are often used together. For model sailplanes the most common approach is to use spanwise twist. it appears that if the boundary layer was sufficiently tripped. on the other hand. It is possible that different heights and chordwise locations might further improve the airfoils tested. it is not always effective. In addition. The SD7003 represents the other extreme. turbulators placed upstream of the laminar separation bubble may improve performance. as may be deduced from the upper-surface shape. While it is known that some airfoils benefit from trips at and above a Rn of 300k. dramatically improve the performance. At a Rn of 60k. drooped leading edges) and stall strips are common. and over the entire Rn range the presence of a bubble cannot be deduced from the polars alone (which is in sharp contrast to the MILEY airfoil). The MILEY airfoil.Chapter 5: Comments on Airfoils have been developed to improve poor stall characteristics. no "miracle" trip shape was found. however. We should point out that other investigators 3 1" have found differences in performance and recommend specific types of trips. given the ability to improve the stall with these techniques. the SD7003 had the lowest drag of any airfoil tested. For Rn below 300k performance gains were achieved. Progressive spanwise twist. has a steep pressure recovery region. but this was not systematically studied in this project.g. The conclusion is that elaborate turbulator devices are not warranted. many low-drag airfoils that might otherwise be reluctantly discarded because a deficient stall can be made usable. the bubble is so shallow and reattaches so quickly that the drag of a trip is about the same as the drag of the bubble it might prevent. were quite effective in alleviating the adverse stall characteristics. Rather. for the airfoils tested in this study. of trips. regardless of how. Stall strips. As a result.5411) showed almost no improvement with changes in spanwise twist. 89 . but these gains were highly dependent on the trip height and position. change of section (e. the performance gain was independent of the particular method.3 Trips and Surface Roughness As amply demonstrated in the airfoil polars. Consequently the laminar flow separates to form a large bubble. The polars show that trips. The MILEY airfoil and the SD7003 are two extreme examples of the effectiveness. where one would normally expect bubble drag to be high. errors two to four times this amount (which are more common) do begin to effect performance. Nonetheless trips are useful to "repair" an otherwise poor performer. low-drag airfoils presented here generally have a gradual recovery. When they are properly used in this context they can produce substantial improvements. For a 12 in chord an error of ±0.4 Trailing Edge Thickness Test on three airfoils (DAE51. The effects of a discrete trip can also be produced by distributed roughness. In order to achieve maximum performance. Since the newer. bumps. essential. typically an airfoil designed for a higher Rn.5 Surface Waviness and Contour Accuracy So long as an airfoil surface is smooth. distributed roughness apparently does not provide any advantage over discrete roughness elements. Much like single roughness elements-trips strips.033% chord. Furthermore some low-Rn airfoils are designed expressly for use with trips 31 • In these cases trips are. one would expect trips to be of marginal value. 5. the airfoil really has no single shape and one can expect its performance to be significantly different from the nominal. Based on our measurements. the Monokotejbalsa waviness has a peak-to-peak amplitude on the order of 0. surface waviness of the type produced by Monokote over balsa appears to have no measurable impact on performance. at least at the higher Rn's. In such cases. 5. produced results qualitatively like that found for the trip strips. all indications are that errors of this order have only a small affect.004 in is only 0.free of sharp edges protruding into the boundary layer. In these experiments "sheet balsa roughness" (no plastic film covering) was tested and. E374. Error in contour is undoubtedly one of the factors that explains differences in performance between two different models of the same RC sailplane. For a 12 in chord. . and SD6080) showed that thick trailing edges produce measurable drag penalties. the usefulness of a trip (on either surface) depends on the severity of the pressure recovery. On the other hand. that is.02% chord.90 Airfoils at Low Speeds Finally. On the other hand. it is necessary to have the thinnest possible trailing edges. of course. the best modern-day construction techniques used by modelers are capable of yielding contours accurate to ±0. blowing through holes-distributed roughness acts to promote transition. the warping of the covering in open-bay construction as it stretches over the cells alters the airfoil contour beyond what is normally considered waviness. as expected. Although at present there is no criterion on the accuracy necessary to meet the nominal performance.004 in. However. 5. 91 .1 DAE51-PT with thickened trailing edge.Chapter 5: Comments on Airfoils DAE51 ------ --J -·--. C======~s~os~o~so~====~==::::::::::~-I·obl 7 t-------3. 5.3 SD6080-PT with thickened trailing edge.2 E374B-PT with thickened trailing edge./6--l Fig.~-!()\ \. Fig. 5.\D ~ Fig. 004 SEAL E214C-PT .g> "'"' ~ Fig.4 Flap configuration: E214C-PT. "' ~ ..310 1 Fig. 5.(0 I>) ..004 HINGE .004 HINGE t-< 0 ~ . t. 5.~.5 Flap configurations: S2091B-PT.6 Flap configuration: SD7032C-PT. .145 C = . 520918-PT A= . ~ /~ C' ~. SD7032C-PT Fig..068 8 = . 5. 93.\~0 N· &.1 (bl 'frlp -"""""'"'" ' •4"" tff u. .chapter 5: CoiJllilents on Airfoils . . -ru-e:.7 (c) Trip configuration: Bumps and holes (blowing).8 Ram configuration.._ Fig. ¢·D31 -B Fig.IO \.'Z.f:>.t. ~ HQ2/9B-PT .:::::7 _l t..94 Airfoils at Low Speeds ¢ \.-t D \__ 1:..c. .?_ 0::.... 5. 5.D.S . .. CG Arm 22 Fus Area• dCm/dCl • -- 100:1: 11..--------. "l i-':' 16 0 ~ 14 § 1t 12 . g 1 10 1. Spaed: ft/s 28 .0 10 ~ ~ ~ ~ ~ ~ ~ 30.--------.055 - L ..• 24 L.12 .---------.20 0 2-E205B 18' ~ 100% {. 5..:. .____.60 2._..· Ci' ~ .--------. 1 S3021A Span • 26 L.. Area • Weight AR Taper R.08 2... Arm L.. Area • s.25 16..--------..20 7.____j____~~_j Fig. s.5 .co .87 -2.--------.--------.9 (a) Aircraft Polar: S3021A-PT vs E2058-PT ::.60 -.14 -.___J_____J_________L_.--------. Span s.....44 4.86 . 44 4.14 -.60 -.86 . . 5..12 ..87 -2.co Ol 10 20 30 40 50 60 70 80 I I I I I I I I 001 Speed: ft/s ~ ~. ~ 1 S3021A Span - ~ Weight AR Teper R.08 2. 1.. Ci' ~.g> "'"' Q.• s. Arm CG Arm Fus Area• dCm/dCl • 100:1: 11. "' s p e e d -2 -3L_______-L--------~--------L--------L------~~------~L-------~------~ Fig. Area s..9 (b) Aircraft Polar: 53021A-PT vs E205B-PT .055 2--E205B 100:1: Aree • - -1 v e r t -- t-< 0 ::. Span S.20 7.60 2.25 16..: . Span • s..12 .14 -. {..• 24 L s.5 l ' 10 ~ 0 ~ .87 -2.>--· C' ~ 10 ~ Fig.08 2..44 4.g "'g .. Arm 22 L CG Arm Fus Area• dCmldCl • - -- 60 70 80 ttls 100% 11...20 0 2--100% 18 L S070320 I I I J \ Q .055 L .25 16.60 2. Area s. 5. "' 16 L I 14 L I 12 L I 1 I I I I I J \ J " ~ I 1.10 (a) Aircraft Polar: S070328-PT vs S070320-PT .60 -...20 7.30° 10 20 30 40 50 Speed: 28 l 1 S070328 Span 26 ~Area • Weight • AR Teper R..86 . .10 (b) Aircraft Polar: ______ 5070328-PT vs S070320-PT _ L _ _ _ _ _ _~ .10 0<> 10 20 30 40 50 60 70 I I I I I I I 001 Speed: 1 S07032B Span • Area • Weight • AR Taper R.87 ft/s 80 I 2--S070320 ~ - "'t-< 0 'l:1 .20 7... -2.86 .· Ci' 100X -2 -3L-______-L--------~--------L--------L--~----~--------L- Fig. Arm CG Arm Fus Area• dCm/dCl • - -1 v e r t -- 100X 11.44 4."' "' "' .g> .08 2.60 2..12 . Span S.25 16... Area • s.• s. 5.60 -.14 -.---.055 s p e e d ~ :::. 12 .Area • Weight AR Taper R. Speed: 28 1 S07084 Span • 26 1-.60 2..14 -.08 2.---------.0 10 ~ 30 40 ~ ~ ~ 80 30r--------.07 .11 (a) Aircraft Polar: S07084-PT vs S07080-PT ~ 10 10 .---------.055 L . Arm L CG Arm 22 tt/s I- -- Fus Area• dCm/dCl • 100% 11.--------.67 16.· 0' ~ Fig.20 7.. Area • s. Span S.60 -. Weight ~ 100% 4.5 ~ \ ~ r § ~ bi ~ g ~ ~ ::.~-------. ~ ~ 16 14 12 1 '' 1.---------. 5.---------..25 - ~.87 -2.44 4..---------.20 0 "I 2---- .= 24 L s..86 . 08 2.25 Area • -1 v e r t s p e e d "' ~ 100% 11.. Area 5.12 .055 -- t-< 0 "' . -· 0' t. Span • s.20 7..87 -2.• S.86 . 5.60 -..11 (b) Aircraft Polar: S07084-PT vs S07080-PT ..14 -.67 16.44 4.g> "' "'f} -2 -3L-------~-------L------~--------L------l~-------L------~------~ Fig. 1 S07084 Span • Weight • AA Taper A..""" 0 0 10 00 20 30 40 50 60 70 80 I I I I I Speed: ft/s .-·.60 2. Arm CG Arm Fus Area• dCm/dCl • 2-S07080 Weight • 100% 4. S. John Wiley & Sons. H. and von Doenhoff. and Boermans. Inc. Hot. Indiana. Vol. Indiana. 1982. Notre Dame. D. Notre Dame. "Boundary Layer Characteristics of the Miley Airfoil at Low Reynolds Numbers. Jones. L." Proceedings of the Conference on Low Reynolds Number Airfoil Aerodynamics. F. W. [11] McGhee. Oxford University Press.. Notre Dame." Memo M-276. 288. London. Profilpolaren fur den Modellflug.J. [10] Althaus. [5] Perry. Necker-Verlag." AIAA Paper 88-0607. Theory of Wing Sections. June 1989. I. "Performance Characteristics from Wind-Tunnel Tests of a Low-Reynolds-Number Airfoil.." AIAA paper 83-1795. VillingenSchwenningen. Low Reynolds Number Vehicles. J. [2] Proceedings of the Aerodynamics at Low Reynolds Numbers 104 < Re < 10 6 International Conference. 101 .. 1980.E. [14] Bastedo. L. October 1986. [12] Volkers.J.J. D. 1985. [7] Althaus.. 1985. Necker-Verlag.J. Delft University of Technology. R. [6] Rae. London. London.. June 1985..J.. FRG. The University of Notre Dame.Wire Anemometry. and Mueller. Indiana. October 1986. "Preliminary Results of Wind Tunnel Measurements on Some Airfoil Sections at Reynolds Numbers between 0.." Proceedings of the Aerodynamics at Low Reynolds Numbers 10 4 < Re < 106 International Conference. 1977. T. [4] Mueller.Chapter 6: References Chapter 6 References [1] Proceedings of the Conference on Low Reynolds Number Airfoil Aerodynamics. Feb. D. FRG. UNDAS-CP-77B123... G. T. October 1986. UNDAS-CP-77B123. A. A. and Jouty." Proceedings of the Aerodynamics at Low Reynolds Numbers 10 4 < Re < 10 6 International Conference.6 x 10 5 and 5..H. "Aerodynamics at Low Reynolds Numbers: A Review of Theoretical and Experimental Research at Delft University of Technology. Profilpolaren fur den Modellflug. M. "Performance of Finite Wings at Low Reynolds Numbers. [9] Althaus. [15] Abbott. [3] Proceedings of the Conference on Low Reynolds Number Aerodynamics. 1959.. 2. Low-Speed Wind Tunnel Testing.. AGARDograph No. Villingen-Schwenningen.. D. New York. and Mueller.0 x 10 5 . [13] Pohlen. 1984. L. and Pope. M. "Recent Wind Tunnel Experiments at Low Reynolds Numbers.G. A. June 1985. R. second ed. T. [8] van Ingen. Dover Publications. The Netherlands. January 1988.. E... "An Efficient Tripping Device.L. John Wiley & Sons.M. [30] Miley.000 and 500. [23] Selig. 10. July 1984." AIAA Paper 74-1017.. and Flight Mechanics.. S. 1989. and Somers. January 1986. VA 23451. Vol. [24] Maughmer.. Vol.. and Giles." AIAA Journal. Stuttgarter Profilkatalog I. van Ingen. and Somers. "Low-Reynolds-Number Airfoil Design for the M. R." Proceedings of the Conference on Low Reynolds Number Airfoil Aerodynamics.. B. "Two-Dimensional Transonic Aerodynamic Design Method. 1983. M. M. [28] Hama.. 24.. Vol..J.M." Proceedings of the Conference on Low Reynolds Number Aerodynamics. The University of Notre Dame. "Airfoil Design for Reynolds Numbers Between 50. D. Daedalus Prototype: A Case Study.A. 25. No.X. August 1980. "A Computer Program for the Design and Analysis of Low-Speed Airfoils. August 1988. UNDAS-CP-77B123. [31] Boermans.. D. "The Design of Airfoils at Low Reynolds Numbers. Including Transition." Soartech 3. M.9. Braunschweig. M.T. "On the Design of Airfoils for Low Reynolds Numbers." AIAA paper 88-4416.000. L. Friedr. October 1986. . Aeronautics. "Experimental Aerodynamic Characteristics of the Airfoils LA 5055 and DU 86-084/18 at Low Reynolds Numbers." Journal of the Aeronautical Sciences (Readers' Forum). R. [22] Drela. U. and Giles. [27] Althaus. private communications. Donker Duyvis. 25. Notre Dame.. Also AIAA-85-0074. No. M. June 1985.D. "Recent Developments in Boundary Layer Computation." AIAA Paper 87-0424.. 1504 North Horseshoe Circle. B. and Wortmann F." Technical Soaring. No. M. Verlag fiir Technik und Handwerk GmbH. R.A.l. and Timmer. Virginia Beach. London. [19] Eppler. [21] Drela. R." NASA TM 80210. Stokely. January 1987...M.102 Airfoils at Low Speeds [16] McCormick. W.. B. D." Proceedings of the Aerodynamics at Low Reynolds Numbers 104 < Re < 10 6 International Conference. [18] Eppler.W. pp 236-237. "Figures of Merit for Airfoil/ Aircraft Design Integration. S.S.. September 1987. F. Indiana. [29] Modell-Technik-Berater 7: HQ Profile. Notre Dame. and Somers. September 1974. M. published by H. F. [17] Althaus. [20] Eppler. M. Vieweg & Sohn Verlagsgesellschaft mbH. 3. A. Aerodynamics. Vol. [25] Drela... J. J. No. [26] Drela. June 1989. "Effects on the Polar Due To Changes or Disturbances To The Contour of the Wing Profile." AIAA Journal. 1981. D. "ISES: A Two-Dimensional Viscous Aerodynamic Design and Analysis Code. M. Indiana. 8. 1. D. 1979.. M. 00000 0.00057 0.00302 -.03806 0.87333 -.92667 0.65333 -.12408 0.39333 0.00961 0.99333 -.00000 0.06405 0.07799 0.00758 -.05313 0.00862 -.93301 0.03065 0.00688 -.08996 0.04077 0.22221 0.02242 -.02000 0.76000 0.01739 0.00920 -.08427 0.00171 0.93301 0.00025 0.02652 0.01566 -.05768 0.50000 0.02000 0.98667 0.37059 0.00000 0.08263 0.07214 0.01000 0.62941 0.02653 -.00000 ----------- CLAAK-Y CLARK-Y 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 1.00939 -.06552 0.03400 0.01131 -.01704 0.02734 -.14645 0.08105 0.08735 0.75000 0.03352 0.44000 -.04171 0.06914 0.01792 -.03806 0.07319 0.01704 0.00972 0.08438 0.02587 0.04677 0.89668 0.00279 0.00898 -.00741 -.00000 103 .00929 0.00115 0.62941 0.02839 -.04394 0.02680 -.96194 0.04945 0.02475 0.00000 0.56526 0.89668 0.00190 -.02214 -.00430 0.00330 0.09312 0.05412 0.60000 -.65333 0.02559 -.33928 -.00940 -.00006 0.00928 -.25000 0.99572 0.02795 -.10332 0.85355 0.06699 0.00866 -.01959 -.54667 0.92667 -.98000 -.76000 -.02653 0.69134 0.00000 0.25000 0.54667 -.85355 0.00000 0.08055 0.99000 0.12667 0.07577 0.00810 -.33928 0.27886 0.06012 0.00572 37 38 39 40 41 42 43 44 45 46 47 48 49 50 51 52 53 54 0.01651 -.01656 0.02351 55 56 57 58 59 60 61 62 63 64 65 66 67 68 69 0.D1333 0.02667 0.82000 -.00616 0.98296 0.00114 0.02458 -.27886 0.00480 0.09128 0.08113 0.09158 19 20 21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 0.30866 0.03333 0.08655 0.31333 0.16000 0.01296 -.08483 0.04786 0.80438 0.96194 0.06000 0.00173 0.08710 19 20 21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 c 0.19333 0.70667 -.70667 0.99667 0.08719 0.03045 0.43474 0.08427 0.30866 0.00333 0.01240 0.06530 0.99667 -.00000 -.99333 0.00429 -.00000 0.69134 0.00575 -.98667 -.00428 0.14645 0.Chapter 7: Airfoil Coordinates L: AQUILA --==--===-=-= AQUILA 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 1.07134 0.98296 0.00019 0.56526 0.98000 0.80438 0.02763 -.02832 -.02018 -.00229 0.37059 0.00961 0.27333 0.02816 -.02567 -.22221 0.49333 0.04667 0.04667 0.60000 0.19333 0.39333 0.82000 0.01345 -.00070 0.02653 0.00009 0.87333 0.00013 0.27333 0.34667 0.05156 0.23333 0.02150 0.00574 -.00428 0.00107 0.00821 -.00333 0.16000 0.01333 0.09333 0.00003 1.17033 0.01000 O.00530 55 56 57 58 59 60 61 62 63 64 65 66 67 68 69 0.03333 0.10332 0.00453 37 38 39 40 41 42 43 44 45 46 47 48 49 50 51 52 53 54 0.02667 0.06699 0.12667 0.00380 0.00917 -.00000 -.01943 0.06000 0.43474 0.00340 0.50000 0.00725 -.05939 0.00618 -.49333 -.99000 -.09333 0.23333 0.05156 0.99572 1.00228 0.75000 0.17033 0.03454 0.08551 0.19562 0.440QO 0.00650 -.00107 0.00787 -.07660 0.00094 -.34667 0.19562 0.09266 0.00047 -.00667 0.02414 -.00122 0.00947 -.01377 0.09318 0.12408 0.00724 -.08774 0.01254 0.00448 0.04027 0.02052 0.31333 0.00667 0. 06717 0.32082 -.64912 0.08527 65 0.07605 0.07580 0.02398 -.01314 97 0.42284 0.92759 87 0.29833 0.43463 -.06051 79 0.00229 -.76130 0.00369 0.62431 11 0.00002 -.09341 -.07793 57 0.06508 51 0.00089 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.03629 -.16912 -. c - DF101 -.91022 0.00000 -.38500 0.00000 Note: Several of the original 120 coordinates were removed due to space limitations.03970 0.50533 0.07652 0.07149 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 0.06279 0.05141 0.00647 0.16715 0.00281 90 0.51050 5 0.00300 0.20557 0.29125 0.00012 -.48520 0.58637 9 0.03611 -.73072 0.61295 0.07436 0.95030 0.94943 0.01617 101 0.08389 99 0.00102 0.00210 0.00015 -.49848 0.00043 0.06757 0.28289 -.08605 0.04484 0.08312 61 0.00209 -.09940 0.63525 0.05331 47 0.01689 -.01353 117 0.79857 0.98041 0.13122 -.00172 0.06211 0.53623 0.03547 -.92195 0.66225 13 0.23394 0.07109 0.03224 41 0.66543 0.03082 0.08460 63 0.00777 55 0.72399 0.45360 -.45162 0.57394 0.03957 43 0.08505 67 0.77558 0.06807 77 0.88981 25 0.00369 33 0.61158 0.01637 0.02988 -.00435 0.01818 0.00000 0.81399 21 0.02477 69 0.00155 -.73812 17 0.03205 -.24708 0.04827 0.00300 -.00200 ------- DFlOl 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.00431 0.00991 95 0.15916 0.01107 111 0.00000 0.02301 0.00021 0.70019 15 0.01444 -.34758 0.00569 93 0.00795 0.88918 0.00155 -.07431 85 0.04877 0.39883 0.97163 0.104 Airfoils at Low Speeds c== ~ OAE51 DAE51 1 1.03397 -.00414 0.01226 0.13128 0.97128 0.00596 89 0.01514 45 0.30935 0.08088 59 0.39670 -.05638 -.89245 0.92375 0.71618 0.38578 0.07192 0.33765 0.00010 0.00376 120 1.00286 0.01512 0.08172 71 0.03475 -.00793 -.00478 0.03046 0.02575 -.76453 0.07397 0.53552 0.83580 0.80990 0.03860 83 0.07845 73 0.01729 37 0.35876 -.00454 -.01544 39 0.16755 0.00000 31 0.00158 -.04975 0.03809 0.05581 0.02450 0.01893 -.77606 19 0.99200 30 0.98715 0.07308 0.00453 119 0.02109 0.00040 0.00418 0.00046 -.99677 1.03216 -.05086 81 0.13208 0.54844 7 0.01218 -.02496 -.01258 0.04663 75 0.43513 0.81783 0.00107 -.00767 -.07396 105 0.00000 0.19643 0.20782 0.68375 0.02458 115 0.00187 0.94717 0.58567 0.01002 35 0.12227 0.68660 0.24496 -.02987 0.07004 53 0.02226 49 50 51 52 53 54 55 56 57 58 59 60 61 0.47256 3 0.46067 0.00793 0.25158 0.00009 91 0.00585 -.99680 0.27159 0.01350 -.85191 23 0.02751 -.20704 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12 13 14 15 16 1.18780 0.80882 0.01740 0.09061 0.01252 0.02450 -.26696 0.01182 49 50 51 52 53 54 55 56 57 58 59 60 61 62 0.34035 0.99661 0.00239 -.06049 0.28998 -.00018 0.02434 -.85050 0.68859 0.99661 0.00606 0.00824 -.00000 0.08332 0.15697 0.18780 0.00039 0.89414 0.54306 0.00000 0.95023 0.07897 0.08213 O.06242 0.01061 0.00010 0.06049 -.07436 0.D7805 0.76339 0.44673 0.00177 -.09527 0.64052 0.98674 0.00000 -.06525 0.03791 -.03025 0.71479 0.85945 0.94794 0.01888 -.19778 0.98674 0.88788 0.01955 0.00000 105 .02136 -.02010 0.00465 0.02558 0.01841 0.00059 0.02652 0.03214 0.05385 0.00034 0.59186 0.89414 0.04642 0.04540 0.03791 0.77923 0.05381 0.64052 0.01224 0. 29699 0.01087 -.07606 0.06822 0.98635 -.48265 0.04258 0.13874 0.00000 0.06507 0.37645 0.00756 0.07111 0.43410 0.00778 0.00988 0.04291 -.71040 0.00210 0.67149 -.02833 0.00422 0.01668 0.08344 0.83461 0.93081 0.00916 0.00155 0.01097 0.02086 -.02435 0.01345 0.51991 0.31449 0.51174 -.00740 0.00039 0.00461 0.79626 0.97312 0.92195 -.68108 0.02236 -._ E214 E214 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.47142 0.00506 0.01689 49 50 51 52 53 54 55 56 57 58 59 60 61 62 0.07813 0.05349 0.00000 0.03321 0.00678 -.00252 0.00684 0.97017 -.00000 .81684 0.30097 -.00000 0.01721 -.02786 0.53217 0.09241 0.81343 -.07207 0.72866 0.00497 0.07342 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 L 0.97049 0.76911 -.06782 0.106 Airfoils at Low Speeds c: E205 ==-- E205 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.02011 -.89034 -.02830 0.04211 0.00000 0.05168 0.96809 0.85400 -.05102 0.03909 0.10406 0.00174 0.01196 0.05811 0.09332 0.01023 0.02436 0.38680 0.89175 0.00495 0.05040 0.06147 0.45751 -.02378 0.98649 0.77412 0.73959 0.00380 0.02021 0.00614 0.07785 0.07532 0.78883 0.03419 0.58218 0.63204 0.16816 0.09281 0.03344 0.03434 0.08639 0.00235 0.00002 0.08214 0.91265 0.61682 0.11157 0.01100 49 50 51 52 53 54 55 56 57 58 59 60 61 62 0.05841 0.99651 -.17764 0.41714 0.00331 0.05937 0.D7390 0.00000 0.42346 0.20510 0.02252 0.02292 0.06669 -.00000 ~ .00923 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0.00003 1.56591 -.01420 0.00994 0.01934 0.94916 0.02292 0.07970 0. 00186 0.01100 0.86035 0.82183 0.00000 0.00268 0.26813 0.00071 0.00624 -.76874 -.61147 0.07529 0.02017 -.03589 0.73567 0.50182 0.00045 0.03230 -.01184 0.01726 0.24083 0.04349 0.04949 0.59272 0.8!560 0.01970 0.12094 0.07037 0.01679 0.07454 0.46580 0.00432 0.00091 0.00000 -- E387 E387 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.00074 0.33968 0.06268 0.00014 0.00000 0.38939 0.53397 0.98705 0.05775 0.01427 0.01264 0.07546 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 0.0!451 0.00605 -.08569 0.03408 0.13845 0.00717 0.64136 0.00882 0.01423 0.88944 0.00000 -.67382 -.00457 -.08035 0.01602 -.00090 0.02768 49 50 51 52 53 54 55 56 57 58 59 60 61 0.01133 -.00000 0.01183 0.03540 0.01441 -.00286 -.15490 0.10428 0.00337 0.04943 0.07020 0.05628 0.26374 0.06390 0.01890 0.31078 0.Chapter 7: Airfoil Coordinates c -- E374 E374 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.00245 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.29824 0.92554 0.00437 0.17669 0.08247 0.96958 -.01177 -.04249 0.98594 -.39252 0.04245 0.89510 0.07936 0.21415 0.00043 0.85202 0.97118 0.04238 0.02003 0.99674 1.25510 0.00643 0.02866 0.03240 -.02662 -.58308 0.03177 -.92214 0.02892 -.00016 1.02297 0.00196 0.02935 0.99637 -.05827 0.49549 0.03065 -.02242 0.66472 0.03256 -.00021 0.05071 0.85254 -.00091 0.07908 0.06448 0.06317 0.19599 0.94942 0.06643 0.00846 0.95128 0.00320 0.01265 -.22742 0.07684 0.34312 0.00178 0.08156 0.04493 0.02748 0.00000 0.00000 0.08173 0.00065 0.05605 0.03559 0.76475 0.36159 0.00804 -.97198 0.03062 -.00258 0.04975 0.02562 0.63210 0.068!5 0.01747 0.02778 0.18906 0.00000 107 .57150 -.51877 0.88892 -.11800 0.00000 0.92205 0.48511 0.00000 0.00700 0.00000 -.01502 -.00423 0.07506 0.00342 0.15345 0.68922 0.00204 0.81027 0.97000 0.98610 0.0!760 0.78007 0.00180 0.10860 0.01153 0.00763 0.03578 0.72243 -.00000 0.03596 0.01500 -.07215 0.98729 0.06820 0.00241 0.89077 0.00410 49 50 51 52 53 54 55 56 57 58 59 60 61 62 63 0.07669 0.02334 0.01329 -.02060 0.2!861 0.94864 0.07543 0.55694 0.62336 -.43681 0.09185 0.00998 -.44679 0.00682 -.54394 0.85508 0.14053 0.03165 -.00307 0.05696 0.81228 -.05033 0.02371 -.31158 0.00228 -.02932 -.01017 -.92085 -.00485 0.00124 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.77292 0.72769 0.44767 0.07251 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 c= 0.17583 0.00132 0.28892 0.41320 0.40077 0.01425 -.94783 -.35505 0.05732 0.68053 0.02570 0.00862 0.71602 0.99640 0.02905 0.99677 0. 00344 0.12408 0.02937 0.12408 0.01123 0.09788 0.37059 0.01879 -.43474 0.80438 0.06699 0.69134 0.75000 0.00296 .08427 0.00000 0.10680 0.75000 0.03754 0.00000 0.43474 0.02653 0.01235 -.89668 0.07552 0.03806 0.00170 0.30866 0.02510 -.01092 0.05982 0.25000 0.50000 0.33928 -.96194 0.00000 FX63-137 FX63-137 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 1.03615 0.30866 0.04584 0.00066 0.00107 0.25000 0.03517 0.02381 -.03548 0.02267 -.22221 0.02199 -.12031 0.12408 0.69134 0.01342 0.02661 0.02273 -.80438 0.01884 -.01617 -.00250 0.02096 -.00000 0.00961 0.19562 0.07880 0.07543 0.69134 0.02653 0.10332 0.01874 55 56 57 58 59 60 61 62 63 64 65 66 67 68 69 0.27886 0.02230 -.85355 0.o2115 -.17033 0.04064 0.10332 0.00991 -.00751 -.06255 0.85355 0.02636 0.01305 0.96194 0.08596 0.03806 0.62941 0.06699 0.37059 0.11122 0.69134 0.00428 0.17033 0.80438 0.11562 0.96194 0.00000 0.12408 0.75000 0.19562 0.01503 -.00503 -.50000 0.01281 0.89668 0.11492 0.06001 0.25000 0.00961 0.07402 0.30866 0.11785 19 20 21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 0.10799 0.05264 0.02402 -.05156 0.06699 0.56526 0.00000 -.08427 0.62941 0.43474 0.00428 0.06052 0.14645 0.00404 0.98296 0.02272 -.07819 0.27886 0.99572 1.07735 0.00107 0.37059 0.06812 0.85355 0.00808 0.56526 0.17033 0.22221 0.00765 0.98296 0.99572 0.00623 0.75000 0.01841 0.03806 0.01704 0.01675 -.00341 0.05491 0.89668 0.99572 1.01719 -.50000 0.02248 -.07876 0.14645 0.08955 0.05078 0.02018 -.00927 0.01016 -.01642 0.00000 0.02056 0.00241 -.01118 -.25000 0.02875 0.22221 0.00702 0.01275 -.56526 0.05156 0.01715 0.27886 0.01418 55 56 57 58 59 60 61 62 63 64 65 66 67 68 69 0.22221 0.07059 0.02653 0.89668 0.02518 -.37059 0.27886 0.06699 0.14645 0.33928 0.30866 0.01029 -.85355 0.01704 0.01482 -.01080 0.01367 0.08427 0.93301 0.01760 0.10332 0.02265 -.17033 0.00514 0.00428 0.02483 -.00428 0.62941 0.00000 0.12180 0.09591 0.56526 0.00961 0.00107 0.03806 0.98296 0.06383 0.01704 0.02653 0.00107 0.62941 0.05156 0.10332 0.01742 0.02287 0.05545 0.12130 0.14645 0.02581 0.108 Airfoils at Low Speeds c:: ======-- FX60-100 FX60-100 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 1.19562 0.05156 0.50000 0.02123 -.00142 -.06902 0.33928 0.33928 -.07708 19 20 21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 0.01706 -.04430 0.00216 37 38 39 40 41 42 43 44 45 46 47 48 49 50 51 52 53 54 0.00226 0.00000 37 38 39 40 41 42 43 44 45 46 47 48 49 50 51 52 53 54 0.93301 0.02331 0.02464 -.08427 0.00192 0.01229 0.00961 0.08270 0.01081 0.98296 0.04643 0.93301 0.07326 0.80438 0.00000 -.10169 0.43474 0.07305 0.02758 0.01933 -.00421 0.02343 0.02043 -.04757 0.99572 0.01704 0.00589 0.02157 -.93301 0.00822 0.02629 0.00702 -.00087 0.96194 0.06743 0.11998 0.19562 0. 83456 0.05514 -.00000 21 22 23 24 25 26 27 28 29 30 0.04359 0.00814 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.29663 0.25000 0.05621 0.90000 0.03102 0.55226 0.02580 31 32 33 34 35 36 37 38 39 0.90000 0.00659 0.00000 0.00043 0.04359 -.02380 0.50000 0.16543 0.04323 0.03880 -.83457 0.05928 -.98907 0.03102 -.06370 0.12843 0.04323 0.70337 0.65451 0.60000 0.00870 -.79389 0.00377 0.00169 -.02320 -.01416 0.05849 0.01612 -.00170 0.93301 0.20000 0.00000 0.00000 0.79389 0.40000 0.05849 -.09549 0.06699 0.25000 0.09549 0.50000 0.00510 -.02950 0.25000 0.02447 0.00710 0.06000 -.05240 0.01560 0.29663 0.85000 0.02500 0.35000 0.05294 0.20000 0.02864 -.00000 -.60000 0.00000 0.03766 0.01490 0.06090 0.50000 0.05928 0.70337 0.01250 0.01490 -.80000 0.04450 0.99726 1.01093 0.20611 0.00760 -.03377 0.34549 0.34549 0.00000 0.06060 0.00377 -.02864 0.15000 0.03377 -.06000 0.40000 0.70000 0.02357 -.00000 -.00000 ==- J5012 J5012 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.97553 0.02480 -.00950 0.00010 0.05360 0.95677 0.30000 0.99726 0.04280 0.02500 0.10000 0.30000 0.00000 -.05621 -.02080 -.95677 0.00500 0.25000 0.60396 0.00000 0.04800 0.00110 0.20611 0.05514 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 0.35000 -.01630 -.95000 0.Chapter 7: Airfoil Coordinates c HQ2/9 HQ2/9 1 2 3 4 5 6 7 8 9 10 1.00169 0.50000 0.02380 -.95000 1.00110 0.01416 -.05189 0.05975 0.06699 0.02570 -.12843 0.70000 0.04871 0.15000 0.05189 -.06300 0.02480 -.98907 0.00274 0.02230 0.01250 0.75000 0.05760 0.00659 -.39604 0.01220 -.44774 0.93301 0.01093 0.05975 -.85000 0.01009 -.04359 -.01870 0.04871 -.02610 -.01612 0.03190 0.02357 0.75000 0.60396 -.03880 0.01009 0.05763 -.05000 0.05294 -.65451 0.04800 49 50 51 52 53 54 55 56 57 58 59 60 61 0.00814 0.80000 0.39604 0.97553 0.00000 109 .00043 0.03766 -.55226 0.00500 0.16544 0.02240 0.06390 11 12 13 14 15 16 17 18 19 20 c 0.87157 0.00274 0.44774 0.02070 -.90451 0.90451 0.10000 0.01870 -.05763 0.05000 0.04359 0.87157 0.02447 0. 42500 -.01333 -.07005 0.06225 0.20000 0.00452 0.00147 -.60000 -.00000 0.00972 -.72493 0.15942 0.06078 0.00000 0.49338 0.95718 0.77354 0.01077 -.07146 0.80073 0.65000 0.30000 0.15000 0.65541 0.02083 0.07718 0.02461 -.00809 -.00014 0.110 Airfoils at Low Speeds MB253515 MB253515 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 1.00652 -.05274 0.10309 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 0.01287 -.01200 -.01260 -.02139 0.43510 0.01809 -.02930 0.52265 0.18796 0.25000 0.75000 •.00228 -.09904 0.04660 0.02803 0.65000 -.04974 -.00098 0.19328 0.88053 0.37500 -.80000 -.01053 -.04866 0.91379 0.04776 -.09487 0.10814 0.12500 0.05807 0.04383 0.17500 0.00000 .50000 0.80000 0.35000 0.01339 -.02960 0.28111 0.01129 49 50 51 52 53 54 55 56 57 58 59 60 61 62 63 64 0.05305 0.15000 0.06460 0.03472 -.30000 0.70652 0.01203 -.01349 -.02033 0.07558 0.27500 0.04941 0.70000 -.07844 0.43862 0.00000 MOB-13-12~ L M06-13-128 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.02636 0.02500 O.01223 -.00073 -.00372 -.12857 0.27500 0.23296 0.00198 0.33185 0.01203 0.01308 -.03001 -.99393 0.00005 0.10057 0.38456 0.01002 0.98232 0.03246 -.22500 0.04622 -.95000 0.99896 0.09002 0.05095 0.09898 0.90000 0.99152 0.47756 0.00994 0.03169 0.96496 0.97756 0.00577 -.04425 -.09473 0.39263 0.04987 0.03878 0.03750 0.00901 -.50000 -.00878 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.08861 0.01070 0.03628 0.90000 -.00370 -.57180 0.00041 35 36 37 38 39 40 41 42 43 44 45 46 47 48 49 50 51 0.40000 0.09660 0.86112 0.07584 0.00233 0.37500 0.17500 0.08582 0.04072 0.32500 0.01000 0.10000 0.11256 0.03393 0.06250 0.02500 0.35032 0.05035 0.08750 0.05000 0.11531 0.84269 0.93068 0.04079 0.25000 0.00088 0.06296 0.04122 0.40000 -.06250 0.32500 -.05026 -.02144 0.04175 -.03551 0.30873 0.85000 -.09868 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 33 34 0.81923 0.00631 0.45000 -.04997 0.12500 0.09602 0.00302 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0.02653 0.00120 -.00000 0.01704 0.00520 0.00961 0.03005 -.02823 -.01436 -. 03151 -.29842 0.79257 0.85355 0.05156 0..02061 -.00218 0.04466 -.00000 -.5411 NACA 2.08427 0.98296 0.02415 -.04431 -.98918 0.04509 -.12408 0.02531 -.75138 0.79520 0.00798 -.17033 0.00218 -.98296 0.89668 0.44807 0.03638 0.00013 0.17033 0.00862 0.00770 0.00172 0.00000 0.03379 -.55132 0.43474 0.03291 -.02039 0.97528 0.03305 0.04295 0.03023 0.83577 0.00000 0.02170 0.14645 0.25255 0.02384 0.19562 0.96194 0.03247 0.03023 -.03247 -.37059 0.01214 -.00767 0.20932 0.00059 0.03564 0.80438 0.04248 -.04427 0.00349 37 38 39 40 41 42 43 44 45 46 47 48 49 50 51 52 53 54 0.04484 55 56 57 58 59 60 61 62 63 64 65 66 67 68 69 0.04397 -.00375 0.02653 0.02384 -.03994 -.04295 -.03564 -.30866 0.01945 -.95720 0.01290 0.01927 0.07831 0.12408 0.44740 0.03553 0.39612 0.03638 -.22221 0.01214 0.00000 -.04431 0.00794 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.Chapter 7: Airfoil Coordinates c NACA 0009 NACA 0009 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 1.13238 0.01945 0.65583 0.02723 0.03994 0.D7080 0.00463 -.07830 0.02823 0.07606 0.05253 0.87053 0.00463 0.03310 -.99723 1.02395 0.10332 0.24745 0.39597 0.50000 0.00994 -.33928 -.00349 0.01704 0.69134 0.06006 0.00770 -.01127 0.00235 0.05777 0.56526 0.00000 -.05156 0.00107 0.08427 0.93238 0.56526 0.05158 0.00897 0.65320 0.00057 0.06317 0.99572 0.04662 O.75000 0.89668 0.04509 0.74862 0.55322 0.03368 -.96194 0.00000 0.09148 0.70475 0.37059 0.02532 0.00000 111 .62941 0.03175 0.03165 -.90368 0.04506 0.70198 0.00621 -.30866 0.07276 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 0.03806 0.03978 -.02039 -.07193 0.20292 0.00961 0.07588 0.04504 -.00428 0.00000 -.01752 0.01677 -.03806 0.03795 0.07912 0.69134 0.01127 -.00000 0.01522 -.90535 0.01380 0.85355 0.62941 0.12448 0.04484 0.00000 ------ NACA 2.01483 -.01646 0.00463 -.02654 0.95633 0.99730 0.16912 0.06659 0.01522 0.02720 -.34642 0.04161 -.14645 0.27886 0.98895 0.33928 0.50000 0.93301 0.06699 0.01925 49 50 51 52 53 54 55 56 57 58 59 60 61 0.03839 0.29485 0.5411 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.04397 0.06347 0.02720 0.02830 -.49933 0.83337 0.00906 0.01208 -.02893 -.00211 -.75000 0.06699 0.25000 0.25000 0.93301 0.03305 -.06852 0. 02973 0.03124 -.00000 0.00221 0.04274 45 0.02331 84 0.00735 41 0.07000 59 0.00189 27 0.01117 53 0.55040 0.04683 39 0.93180 0.00737 0.02103 56 0.00735 15 0.01112 0.03052 109 0.04274 -.40000 0.606~7 0.30012 0.00101 0.04021 -.00573 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.99732 0.00956 -.00815 43 0.00000 87 0.01772 0.74728 0.00886 67 0.39590 0.34730 0.01467 0.20000 0.60167 -.00000 57 0.01284 0.00844 -.00000 83 0.98930 0.00000 28 0.13607 0.00400 0.02724 0.00025 -.10086 0.00950 47 0.95760 0.03124 35 0.00720 0.50132 0.09000 0.06982 0.86957 0.02102 33 0.03471 107 0.02103 Note: Several of the origina1 120 coordinates were removed due to space limitations.01871 0.00000 -.00000 0.05378 0.00737 0.00900 -.00075 0.00300 -.01017 0.09000 61 0.30000 69 0.05889 0.00225 -.03000 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-.00300 0.00900 0.02331 -.00000 0.05968 0.00815 13 0.08874 0.25498 0.04683 -.98883 0.00700 0.00700 -.30000 0.97603 0.09584 0.00000 0.01010 49 0.03471 -.90288 0.01929 49 50 51 52 53 54 55 56 57 58 59 60 61 NACA 64A010 NACA 64A010 .01065 63 0.44707 0.79647 0.00327 31 0.04995 -.01451 0.04995 99 0.112 Airfoils at Low Speeds c ~ NACA 6409 NACA 6409 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.90000 81 0.17257 0. -.39618 0.01080 -.02720 -.99722 1.07153 0.01634 0.01063 0.00886 17 0.93423 0.08684 0.01063 89 0.75272 0.00500 -.05000 0.00600 -.06283 0.01639 0.08780 0.12080 0.70000 0.15830 0.01603 0.04464 0.00000 O.01780 0.00125 0.00125 -.07000 0.04970 0.80000 79 0. 80575 0.00060 0.40189 0.50176 0.00534 -.00307 0.03418 0.04146 0.05649 0.98681 0.02152 -.02877 0.03343 0.03256 0.02345 0.02600 -.30225 0.03985 0.02076 0.49702 0.74243 0.00028 -.97059 0.07010 0.54753 0.02123 -.89810 0.88712 0.01006 -.03658 0.04467 0.01935 0.02696 -.06004 0.94748 0.00048 -.78557 0.50596 0.97237 0.50713 -.01247 0.97174 0.00691 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0.02603 0.02559 -.00000 0.14161 0.04457 0.04518 0.02600 -.99660 1.06123 0.03715 0.00064 0.61084 0.06086 0.91925 0.02106 0.55944 0.01366 -.00865 0.00000 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.97003 0.95126 0.00127 0.05486 0.06162 0.00001 -.06190 0.00413 -.05621 0.00235 0.00018 0.59778 0.00506 -.03954 0.05318 0.76037 0.02546 -.00045 0.44676 0.02762 -.03654 0.39727 0.05068 0.16963 0.40257 0.35127 0.00000 0.35156 0.02553 -.17730 0.02590 0.40212 0.00021 0.00034 0.02654 0.02785 -.00932 0.02590 -.74165 0.00001 0.98652 0.00101 0.69691 0.21624 0.04564 0.95248 0.05765 0.05966 0.25531 0.01717 -.00552 -.64968 0.Chapter 7: Airfoil Coordinates c ----- RG15 RG15 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.05727 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 c 0.99671 0.21878 0.05673 0.01395 -.11087 0.00950 0.99671 0.00208 -.71302 0.13224 0.80700 0.01934 -.09896 0.01200 0.00040 0.69537 0.60115 0.00192 -.78572 0.05039 0.94833 0.88583 0.89683 0.00943 -.06036 0.84779 0.64723 0.84914 0.26132 0.82623 0.01667 -.06025 0.34902 0.00724 -.02644 0.02152 0. 49629 0.01580 -.97150 0.13818 0.23417 0.02187 -.26043 0.99674 0.00164 0.00389 -.89548 0.21835 0.05397 0.00354 0.81694 0.76843 0.82383 0.01374 0.97126 0.05187 0.00000 0.05557 0.00058 0.00000 0.03168 0.00738 0.04688 0.68455 0.97110 0.22002 0.01374 0.34805 0.00403 0.69297 0.00397 0.07830 0.85595 0.00019 0.14765 0.02081 0.02354 -.01713 0.01955 49 50 51 52 53 54 55 56 57 58 59 60 61 0.57154 -.54654 0.25546 0.66007 0.10633 0.17385 0.64527 0.10338 0.94912 0.05478 0.04135 0.03263 0.39625 0.05624 0.38556 0.02390 -.02292 -.00537 -.00870 -.00367 0.00517 0.46591 0.00401 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.02645 0.00575 0.67376 0.00642 -.02205 -.17622 0.00295 0.02043 -.00174 0.00391 0.08063 0.00182 0.99675 1.77501 0.85037 0.49529 0.01890 0.05421 0.55090 -.01315 -.01704 0.00312 -.00001 0.26365 0.01607 0.06384 0.05420 0.88861 0.73071 0.59634 0.98671 0.18232 0.03888 0.00001 -.98706 0.08028 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 0.99670 0.89147 0.00025 0.72225 0.86156 0.00000 ~ 52091 S2091 l 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.07349 0.00144 0.00803 0.53966 0.60607 0.01794 -.00003 0.00000 0.00864 -.01855 -.02358 -.01750 -.00035 0.08804 0.51885 0.01844 -.01269 0.34755 0.03577 0.00699 0.00020 0.88835 0.03433 0.04939 0.01150 0.02387 -.02571 0.00037 0.95075 0.02204 0.01141 -.00075 -.05623 0.00423 -.08823 0.94970 0.99666 1.00617 -.01872 -.00296 0.06231 0.98702 0.05059 0.00367 0.01935 -.00397 -.02092 -.43996 0.01380 -.71218 0.00000 .08676 0.76169 0.02992 0.44602 0.03594 0.00101 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.02855 0.39450 0.44220 0.02293 -.00076 0.00894 -.97046 0.85210 0.01604 -.114 Airfoils at Low Speeds c - 52055 52055 l 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.00073 0.01645 -.62337 0.02356 0.01051 0.05215 0.04585 0.04320 0.00170 49 50 51 52 53 54 55 56 57 58 59 60 61 0.30044 0.58856 0.01373 -.49073 0.33272 0.80796 0.00009 0.05552 0.04884 0.11237 0.04504 0.28206 0.06708 0.01662 0.08793 0.30323 0.04061 0.21310 0.04130 0.98707 0.94834 0.00222 0.01939 -.41332 0.07381 0.92142 0.05408 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 L 0.92529 0.00952 -.01130 -.01151 0.13776 0.06920 0.00012 0.03074 0.07933 0.18960 0.36169 0.00150 0.05927 0.07525 0.63701 0.00000 0.01321 -.31161 0.11622 0.03107 0.00400 0.73898 0.08415 0.08355 0.92292 0.08643 0.78279 0.02373 0.81187 0.92087 0.14884 0.00205 -.01620 -.04846 0. 02221 -.01042 0.43275 0.85493 0.81531 0.00027 0.48405 0.13239 0.10585 0.07235 0.04741 0.04658 0.05971 0.00075 0.88906 0.09879 0.01183 0.38438 0.00000 0.01731 -.00000 0.06689 0.00000 -.05971 0.00526 -.97122 0.66399 0.00278 0.94959 0.00042 0.07556 0.97044 0.05655 0.07796 0.02610 -.07192 0.92405 0.97104 0.01894 -.09180 0.03607 0.99663 0.85625 0.00564 0.05360 0.71495 0.06434 0.01620 -.00264 0.04545 0.07870 0.05337 0.99668 1.01246 -.01064 -.00749 -.02203 -.01048 -.33750 0.01433 49 50 51 52 53 54 55 56 57 58 59 60 61 115 .02766 0.35564 0.89090 0.00072 0.03259 0.07990 0.99674 0.01769 -.03997 0.06322 0.02329 -.07184 0.00911 0.06541 0.01506 -.29376 0.01801 -.00005 0.02385 -.55743 -.07422 0.68001 0.02531 -.85297 0.13785 0.00124 -.98699 0.00367 -.12435 0.00010 0.17206 0.07864 0.06381 0.02006 -.02610 0.02088 0.24635 0.02617 0.61127 0.68021 0.94905 0.00013 0.02399 -. 77261 0.62242 0.81577 0.00021 0.53368 0.01960 0.99674 1.63158 0.98703 0.03958 0.07902 0.01394 0.00936 0.00682 -.00736 -.00086 0.04012 0.05535 0.30520 0.05850 0.00472 -.44899 0.00978 0.51527 0.07043 0.01970 -.00003 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.00932 0.53204 0.00172 -.00595 -.00972 -.00927 0.94840 0.89250 0.02244 -.06790 0.00031 0.01970 0.05403 0.03713 0.92264 0.48213 0.00994 -.02239 0.34371 0.07706 0.03683 0.00118 0.29250 0.80901 0.81342 0.01405 0.20944 0.07029 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 c 0.72739 0.00526 -.00893 -.03190 0.46129 0.04061 0.02204 -.85087 0.38629 0.05301 0.06973 0.00000 0.02009 -.06704 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 0.00332 -.58203 0.25011 0.67431 0.01835 0.00000 0.95049 0.00277 -.20924 0.25713 0.56920 -.88851 0.21197 0.58321 0.02567 -.20200 0.00052 -.00543 -.00181 0.02335 -.03310 0.02625 -.00000 0.40787 0.00001 -.10706 0.02173 0.77158 0.92140 0.00051 0.98667 0.24970 0.97151 0.06076 0.00199 -.00293 0.72423 0.00319 0.39565 0.02585 0.00000 - 53014 S3014 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.01514 -.00532 0.06988 0.02427 -.04556 0.00099 0.77104 0.92108 0.16119 0.17136 0.01383 -.43477 0.76350 0.02386 -.33908 0.02344 0.98706 0.17024 0.02120 -.01252 49 50 51 52 53 54 55 56 57 58 59 60 61 0.29354 0.00100 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.07166 0.00000 0.01492 0.04685 0.13682 0.63214 0.72642 0.00284 0.50312 0.Chapter 7: Airfoil Coordinates c 53010 S3010 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1. 13666 0.17816 0.06243 0.61834 0.03393 0.16726 0.00602 0.00095 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.01115 49 50 51 52 53 54 55 56 57 58 59 60 61 0.04459 0.02059 -.01386 0.00769 0.02807 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-.00000 0.00344 0.01957 -.00063 0.51047 0.01592 -.06252 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 c 0.25118 0.02053 •.20504 0.01266 0.98679 0.30857 0.00039 0.94996 0.02526 -.06624 0.48891 0.98700 -.16464 0.00606 -.05631 0.02251 -.01543 -.03706 0.76840 0.21611 0.06528 0.06542 0.00963 -.06066 0.53839 0.29999 0.07151 0.25685 0.67865 0.02810 -.67514 -.06753 0. 77035 0.51705 0.89005 0.94843 0.97154 -.02146 0.00763 0.00001 -.12712 0.00000 0.02400 0.01160 0.40237 0.04337 0.02442 -.01363 0.38265 0.02358 0.00365 -.00852 -.01588 0.00012 1.00000 .00374 0.17444 0.00984 0.02964 0.01892 0.34513 0.00003 0.24593 0.10216 0.01232 -.00476 0.00000 0.81283 0.00570 -.01742 -.05341 0.33517 0.00193 -.01670 0.98669 0.43996 0.20614 0.03743 0.77202 -.00278 -.00839 0.02776 0.02891 0.00690 ·.02320 0.04888 0.07312 0.00172 0.85355 0.00017 0.07529 0.01876 -.03047 0.04930 0.00035 0.00821 ·.07632 0.02604 -.03329 0.88898 0.92398 0.26085 0.02085 •.48109 0.62352 -.43144 0.04917 0.01941 ·.72079 0.06707 0.10296 0.05092 0.00000 0.94876 0.01088 -.35862 0.85840 0.28946 0.116 Airfoils at Low Speeds c 53016 83016 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.63730 0.00203 0.63040 0.02472 0. 02838 0.71357 0.80895 0.31661 0.92297 0.08024 0.06173 0.99678 1.23566 0.98695 0.00608 0.85115 0.00400 0.01224 -.06828 0.00495 0.98709 0.08301 0.00107 0.00700 -.94869 0.30862 0.14793 0.00090 0.01986 0.00178 0.00283 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.00000 ~ 54062 84062 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.39979 0.07876 0.06147 0.08097 0.03765 0.88751 0.77610 0.04627 0.05470 0.00363 0.00279 -.00146 0.00502 0.00681 0.04487 0.66164 0.08491 0.33414 0.85637 0.00787 -.06881 0.09326 0.50397 0.55168 0.43788 0.97086 0.00441 -.92304 0.04604 0.92184 0.01499 0.00273 0.01907 -.22460 0.92078 0.23092 0.03945 0.89171 0.08542 0.07414 0.73609 0.01482 -.00491 0.00578 0.59189 0.15074 0.77881 0.98696 0.00965 -.01724 -.02876 0.65845 0.00698 0.81765 0.54909 -.07974 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 c: 0.00147 0.05435 0.11855 0.00405 0.01659 -.63971 0.69165 0.11475 0.01800 -.00081 0.01729 0.07513 0.33049 0.60777 0.00069 49 50 51 52 53 54 55 56 57 58 59 60 61 0.99671 0.01291 -.05339 0.00034 0.08227 0.00178 0.04616 0.94954 0.02799 0.01209 -.01719 0.84932 0.71066 0.03342 0.01937 0.73227 0.08239 0.07864 0.59909 0.00000 0.15086 0.26554 0.00046 0.35348 0.06848 0.49704 0.18619 0.12259 0.00074 0.04752 0.01462 -.01145 -.88894 0.36217 0.00196 0.00301 0.01545 -.01864 -.05437 0.38336 0.19171 0.00419 0.27271 0.49336 0.01124 0.00000 0.03151 0.00339 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.06769 0.01156 0.76289 0.94940 0.08399 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 0.00000 117 .00437 0.07393 0.23229 0.01844 -.00511 0.05541 0.97129 0.01034 -.97127 0.00438 0.01380 0.01835 0.00771 -.08673 0.01716 -.55267 -.54359 0.00225 0.15550 0.49522 0.00836 0.00342 0.45626 0.76031 0.06186 0.28346 0.00199 0.00439 0.00000 0.00262 -.07519 0.00364 0.64586 0.27993 0.98714 0.08329 0.01868 -.00041 0.08453 0.44716 0.00320 -.02403 0.38708 0.Chapter 7: Airfoil Coordinates c-= --------- 54061 S4061 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.00079 0.01069 0.00245 0.00000 0.80673 0.00725 -.85727 0.08988 0.44161 0.81936 0.03741 0.18813 0.02676 0.60436 0.19130 0.08510 0.00025 0.00020 0.08630 0.00000 0.94977 0.05666 0.99673 1.08717 0.00061 0.04385 0.00001 0.02843 0.06283 0.03628 0.01780 -.11211 0.01527 -.97107 0.02036 0.01374 0.89200 0.40887 0.01932 -.99675 0.00516 -.03499 0.00337 0.06493 0.68665 0.00015 49 50 51 52 53 54 55 56 57 58 59 60 61 0.06949 0.00363 0. 07351 -.00607 0.38846 0.03714 -.43446 0.18140 -.85861 0.99684 0.30354 0.79967 0.98608 0.65200 0.53609 0.01986 0.98746 0.00043 0.01444 0.59939 0.37937 -.04497 0.12392 0.48820 0.06322 0.05055 0.00407 0.394ll 0.02138 0.00602 0.01523 0.08523 0.01606 0.14066 -.00456 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.26227 0.99650 1.94864 0.02605 0.00129 0.96949 0.08751 0.76933 0.96905 0.00881 0.18457 0.01584 0.01661 0.14915 0.65661 0.00337 -.02934 -.00000 0.30450 0.03768 -.04791 -.03559 -.00000 0.00541 0.09405 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 0.75290 0.00042 0.00432 0.03497 -.68660 0.39426 0.00274 0.00177 0.08471 0.00061 0.32586 -.00183 49 50 51 52 53 54 55 56 57 58 59 60 61 0.00709 0.84849 0.04398 0.44203 0.00643 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 1.08881 0.09867 0.34856 0.00814 0.72503 0.44078 0.00696 0.07420 0.95106 0.04069 0.04319 0.07443 0.06585 0.88703 0.00001 -.97208 0.08934 0.97094 0.08218 0.00040 0.00801 0.03672 -.04740 0.60221 0.03003 -.18382 0.88694 0.09ll7 0.04553 0.09770 0.00433 0.00131 0.05280 0.07095 0.09570 0.92482 0.27460 -.10453 -.01680 0.118 Airfoils at Low Speeds c: S41BO ==---===----=-= 84180 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.94576 0.01300 0.81970 0.00910 -.70909 0.07996 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 c 0.02602 0.00461 0.98618 0.06475 0.09ll8 0.01201 0.09017 0.00186 0.54018 0.Oll32 0.06437 0.00377 -.07798 0.26125 0.08934 0.33802 0.02971 0.03817 0.06761 0.00036 0.06040 0.03718 0.01286 -.02144 -.14971 0.00017 0.63196 0.88220 0.98704 0.00001 0.49043 0.00775 0.80565 0.02237 0.10001 0.00000 0.84300 0.00200 0.20013 0.05468 0.04033 0.44032 0.24338 0.92048 0.00000 ------------ S4233 84233 1 2 3 4 5 6 7 8 9 10 ll 12 13 14 15 16 0.02939 0.67908 0.00156 0.00916 0.81146 0.01032 0.73313 0.00058 0.02558 49 50 51 52 53 54 55 56 57 58 59 60 61 0.00149 -.03249 -.00035 0.11756 0.99676 1.06388 0.00787 0.08329 0.05674 0.02279 0.22218 0.00156 0.16018 0.00026 0.00702 O.03321 -.01466 0.01234 0.54659 0.00736 0.49082 0.00000 0.00488 -.01306 -.77768 0.06963 0.02609 -.00263 0.75899 0.00014 0.00592 -.34797 0.09171 0.91927 0.00652 0.03782 -.28950 0.00956 0.589(>1 0.94698 0.07770 0.22126 0.00016 -.22624 -.03596 0.00376 0.11788 0.06041 0.58414 0.01669 0.08828 0.91665 0.02139 -.02779 -.70342 0.01621 -.01703 -.49308 0.09969 0.00000 .54626 -.02999 0.01416 0.02725 0.01058 -.89384 0.85089 0.99650 0.63859 0.09049 0. 00152 0.82542 0.01176 0.05328 14 0.00001 -.00210 ·.86252 0.01306 -.01997 -.06655 0.01438 0.04320 0.78512 0.06624 0.50273 0.07251 0.00094 -.26288 0.04664 0.00212 0.01440 0.20818 0.05798 0.01741 0.00000 0.01977 -.06460 0.00501 -.45282 0.00017 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.99670 1.00716 0.00304 -.05718 0.01634 49 50 51 52 53 54 55 56 57 58 59 60 61 0.35182 0.00060 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.61905 0.11158 0.76373 0.00066 0.67117 0.81350 0.00226 -.03475 0.00049 0.01371 0.07236 0.00338 0.85428 0.12941 0.92270 0.00853 -.01071 0.73412 0.98643 0.51038 0.21990 0.00044 -.09515 0.64208 0.02183 -.00939 -.40396 0.00035 0.03078 0.06203 0.85666 0.00510 0.54546 0.02369 0.35168 0.02213 -.72136 0.71670 0.99659 1.02106 -.09584 0.06478 0.97089 0.00637 -.49645 0.05724 15 0.00059 0.96985 0.99669 0.01439 -.08273 0.00782 -.69709 0.03010 0.35488 0.06415 0.02018 -.00374 0.35208 0.05737 0.01837 -.01213 -.05143 0.07121 0.00000 0.02803 0.01844 -.03952 0.94955 0.45750 0.00004 0.25331 0.97108 0.02113 -.25401 0.56563 -.00249 0.30137 0.00509 -.00015 -.40406 0.98686 0.02096 0.01286 -.01113 0.01491 -.00930 0.00698 0.11313 0.00157 -.00437 -.04234 0.81846 0.77745 0.03772 0.06360 0.61621 0.45693 0.01829 -.01725 -.76901 0.06964 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 0.06038 16 0.02224 0.01921 -.03791 0.98686 0.80802 0.66738 0.59415 0.30788 0.00995 0.40336 0.56368 SD2083 -.14532 0.92265 0.94728 0.91925 0.03752 0.55259 0.01076 ·.02384 0.00017 0.06601 0.00024 0.00001 ·.16649 0.05157 0.92535 0.89593 0.01696 -.39913 0.00848 -.74219 0.14432 0.00281 0.99665 0.18069 0.08464 0.20896 0.30568 0.02150 0.89158 0.01908 -.04100 0.02564 0.05146 0.06693 0.00398 -.89084 0.00220 0.01105 49 50 51 52 53 54 55 56 57 58 59 60 61 0.65019 0.01992 0.01601 -.04489 0.06243 0.04296 0.00000 ===---==--====- SD2083 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.22034 0.98693 0.26160 0.01592 -.00000 119 .51157 0.88632 0.68886 0.02419 0.00397 ·.18101 0.03150 0.44755 0.04489 0.30191 0.00657 0.00616 -.02207 -.04856 0.02170 -.06268 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 c 0.84906 0.06908 0.00935 -.00000 0.00000 0.00003 0.07162 0.01471 ·.02319 0.03742 0.16727 0.95059 0.01274 -.97128 0.Chapter 7: Airfoil Coordinates c ---- 502030 SD2030 1 2 3 4 5 6 7 8 9 10 11 12 13 1.601~9 0.00049 0.01937 -.94929 0.06004 0.00000 0.01780 -.05576 0.06179 0.05917 0.01032 0.12868 0.01970 -. 120 Airfoils at Low Speeds c 505060 SD5060 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.00000 0.99678 0.98720 0.97149 0.95002 0.92326 0.89173 0.85600 0.81664 0.77422 0.72929 0.68236 0.63396 0.584_57 0.53465 0.48468 0.00000 0.00023 0.00102 0.00256 0.00500 0.00838 0.01268 0.01778 0.02351 0.02967 0.03600 0.04227 0.04824 0.05370 0.05847 0.06239 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 c 0.43508 0.38632 0.33882 0.29301 0.24931 0.20814 0.16992 0.13504 0.10383 0.07656 0.05343 0.03444 0.01954 0.00878 0.00223 0.00003 0.06536 0.06730 0.06816 0.06793 0.06664 0.06433 0.06105 0.05683 0.05168 0.04565 0.03875 0.03104 0.02280 0.01443 0.00637 -.00066 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.00334 0.01244 0.02641 0.04507 0.06835 0.09613 0.12829 0.16475 0.20524 0.24935 0.29661 0.34651 0.39845 0.45183 0.50602 0.56034 •.00656 -.01208 -.01729 ·.02178 ·.02538 ·.02795 •.02938 ·.02973 -.02920 -.02795 -.02611 -.02384 -.02126 ·.01851 ·.01568 -.01288 49 50 51 52 53 54 55 56 57 58 59 60 61 0.61416 0.66680 0.71760 0.76592 0.81113 0.85265 0.88992 0.92244 0.94974 0.97144 0.98721 0.99679 1.00001 ·.01020 -.00772 ·.00551 ·.00361 ·.00206 -.00089 ·.00008 0.00039 0.00057 0.00051 0.00032 0.00010 0.00000 - 506060 SD6060 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.00000 0.99661 0.98660 0.97033 0.94829 0.92100 0.88905 0.85301 0.81346 0.77096 0.72602 0.67917 0.63091 0.58174 0.53222 0.48283 0.00000 0.00023 0.00108 0.00283 0.00559 0.00941 0.01419 0.01977 0.02595 0.03248 0.03912 0.04563 0.05177 0.05738 0.06225 0.06606 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 0.43386 0.38566 0.33862 0.29316 0.24976 0.20883 0.17076 0.13589 0.10456 0.07700 0.05344 0.03399 0.01879 0.00790 0.00148 0.00025 0.06866 0.07003 0.07020 0.06922 0.06715 0.06402 0.05988 0.05480 0.04887 0.04218 0.03486 0.02710 0.01913 0.01132 0.00411 -.00159 33 34 35 36 37 38 39 40 41 .42 43 44 45 46 47 48 0.00495 0.01525 0.03068 0.05114 0.07648 0.10645 0.14078 0.17909 0.22096 0.26592 0.31347 0.36306 0.41413 0.46614 0.51852 0.57073 -.00647 ·.01148 ·.01612 -.02025 ·.02381 -.02678 -.02919 ·.03105 -.03238 -.03321 ·.03354 ·.03338 ·.03273 -.03159 ·.02995 ·.02784 49 50 51 52 53 54 55 56 57 58 59 60 61 0.62223 0.67254 0.72116 0.76761 0.81133 0.85176 0.88838 0.92070 0.94818 0.97032 0.98661 0.99662 1.00001 ·.02527 -.02231 -.01906 -.01568 -.01236 -.00922 -.00638 -.00399 -.00214 ·.00090 ·.00024 ·.00002 0.00000 Chapter 7: Airfoil Coordinates c-: ----- 506080 SD6080 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.00000 0.99676 0.98716 0.97155 0.95032 0.92396 0.89301 0.85799 0.81945 0.77793 0.73396 0.68807 0.64077 0.5925,7 0.54399 0.49546 0.00000 0.00037 0.00159 0.00383 0.00717 0.01160 0.01704 0.02332 0.03025 0.03758 0.04508 0.05250 0.05960 0.06613 0.07185 0.07646 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 c 0.44730 0.39986 0.35351 0.30859 0.26547 0.22453 0.18616 0.15071 0.11851 0.08984 0.06490 0.04387 0.02684 0.01391 0.00509 0.00045 0.07980 0.08176 0.08234 0.08154 0.07941 0.07607 0.07160 0.06609 0.05965 0.05238 0.04442 0.03595 0.02720 0.01844 0.01004 0.00260 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.00082 0.00692 0.01860 0.03565 0.05796 0.08539 0.11771 0.15462 0.19573 0.24059 0.28870 0.33948 0.39233 0.44662 0.50168 0.55682 -.00296 -.00710 -.01060 -.01322 -.01494 -.01581 -.01592 -.01539 -.01433 -.01285 -.01109 -.00916 -.00717 -.00519 -.00334 -.00166 49 0.61136 0.66463 51 0.71596 52 0.76470 53 0.81024 54 0.85200 55 0.88944 56 0.92207 57 0.94944 58 0.97121 59 0.98707 60 0.99675 61 1.00001 -.00020 0.00099 0.00189 0.00250 0.00282 0.00287 0.00267 0.00228 0.00175 0.00118 0.00063 0.00018 0.00000 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.00457 0.01408 0.02839 0.04763 0.07182 0.10073 0.13407 0.17150 0.21268 0.25719 0.30456 0.35426 0.40572 0.45837 0.51161 0.56484 -.00741 -.01285 -.01759 -.02141 -.02438 -.02660 -.02809 -.02888 -.02900 -.02852 -.02752 -.02608 -.02428 -.02217 -.01980 -.01723 49 0.61748 so 0.66898 51 0.71883 52 0.76644 53 0.81118 54 0.85241 55 0.88957 56 0.92210 57 0.94952 58 0.97134 59 0.98718 60 0.99679 61 1.00001 -.01450 -.01167 -.00887 -.00628 -.00403 -.00220 -.00082 0.00008 0.00052 0.00057 0.00037 0.00011 0.00000 so 507003 SD7003 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.00000 0.99681 0.98745 0.97235 0.95193 0.92639 0.89600 0.86112 0.82224 0.77985 0.73449 0.68673 0.63717 0.58641 0.53499 0.48350 0.00000 0.00031 0.00132 0.00310 0.00547 0.00824 0.01139 0.01494 0.01884 0.02304 0.02744 0.03197 0.03649 0.04086 0.04494 0.04859 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 0.43249 0.38250 0.33405 0.28760 0.24358 0.20240 0.16442 0.12993 0.09921 0.07244 0.04978 0.03130 0.01702 0.00697 0.00127 0.00025 0.05171 0.05415 0.05581 0.05658 0.05639 0.05518 0.05292 0.04961 0.04526 0.03993 O.Q3372 0.02677 0.01932 0.01172 0.00438 -.00186 121 122 Airfoils at Low Speeds c= - ------- 507032 SD7032 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.00000 0.99674 0.98712 0.97155 0.95054 0.92464 0.89436 0.86021 0.82264 0.78208 0.73892 0.69356 0.64646 0.59812 0.54902 0.49967 0.00000 0.00048 0.00204 0.00485 0.00894 0.01420 0.02041 0.02731 0.03460 0.04199 0.04925 0.05620 0.06270 0.06861 0.07381 0.07816 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 c 0.45058 0.40222 0.35506 0.30953 0.26604 0.22499 0.18671 0.15146 0.11948 0.09105 0.06627 0.04524 0.02812 0.01502 0.00606 0.00115 0.08154 0.08385 0.08500 0.08493 0.08359 0.08096 0.07703 0.07182 0.06548 0.05809 0.04976 0.04078 0.03145 0.02206 0.01293 0.00448 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.00038 0.00532 0.01649 0.03308 0.05491 0.08180 0.11351 0.14974 0.19010 0.23420 0.28153 0.33154 0.38364 0.43724 0.49176 0.54659 -.00223 -.00701 -.01088 -.01403 -.01635 -.01787 -.01862 -.01867 -.01810 -.01699 -.01547 -.01363 -.01152 -.00922 -.00678 -.00430 49 50 51 52 53 54 55 56 57 58 59 60 61 0.60112 0.65469 0.70664 0.75634 0.80313 0.84635 0.88534 0.91942 0.94797 0.97054 0.98684 0.99670 1.00001 -.00190 0.00030 0.00224 0.00379 0.00485 0.00535 0.00526 0.00458 0.00350 0.00226 0.00113 0.00030 0.00000 ---- 507037 SD7037 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.00000 0.99672 0.98707 0.97146 0.95041 0.92450 0.89425 0.86015 0.82261 0.78201 0.73865 0.69294 0.64539 0.59655 0.54693 0.49706 0.00000 0.00042 0.00180 0.00436 0.00811 0.01295 0.01865 0.02490 0.03141 0.03788 0.04413 0.05011 0.05572 0.06085 0.06538 0.06917 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 0.44745 0.39862 0.35101 0.30508 0.26125 0.21989 0.18137 0.14601 0.11410 0.08586 0.06146 0.04102 0.02462 0.01232 0.00418 0.00021 0.07211 0.07410 0.07504 0.07488 0,07358 0.07113 0.06754 0.06286 0.05715 0.05049 0.04300 0.03486 0.02632 0.01770 0.00936 0.00185 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.00127 0.00806 0.02038 0.03800 0.06074 0.08844 0.12084 0.15765 0.19850 0.24296 0.29055 0.34071 0.39288 0.44643 0.50074 0.55519 -.00393 -.00839 -.01227 -.01541 -.01777 -.01934 -.02017 -.02032 -.01987 -.01891 -.01754 -.01586 -.01396 -.01190 -.00976 -.00760 49 50 51 52 53 54 55 56 57 58 59 60 61 0.60914 0.66197 0.71305 0.76178 0.80752 0.84964 0.88756 0.92071 0.94859 0.97077 0.98690 0.99671 1.00001 -.00549 -.00349 -.00168 -.00014 0.00104 0.00182 0.00220 0.00218 0.00185 0.00132 0.00071 0.00021 0.00000 Chapter 7: Airfoil Coordinates c - =--===--- 807043 SD7043 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.00000 0.99681 0.98736 0.97202 0.95121 0.92544 0.89522 0.86108 0.82351 0.78301 0.74004 0.69500 0.64825 0.60019 0.55127 0.50199 0.00000 0.00046 0.00191 0.00451 0.00828 0.01315 0.01897 0.02553 0.03256 0.03979 0.04690 0.05362 0.05974 0.06515 0.06975 0.07347 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 0.45282 0.40422 0.35665 0.31056 0.26639 0.22457 0.18549 0.14954 0.11702 0.08821 0.06331 0.04249 0.02581 0.01334 0.00509 0.00083 0.07620 0.07789 0.07850 0.07802 0.07645 0.07382 0.07015 0.06549 0.05989 0.05340 0.04612 0.03818 0.02975 0.02106 0.01236 0.00404 33 34 35 36 37 38 39 40 41 42 43 44 45 46 0.00052 0.00555 0.01669 0.03324 0.05501 0.08183 0.11345 0.14956 0.18979 0.23374 0.28094 0.33088 0.38302 0.43678 47 0.49153 48 0.54665 -.00278 -.00770 -.01150 -.01452 -.01669 -.01802 -.01856 -.01839 -.01756 -.01617 -.01429 -.01204 -.00951 -.00683 -.00412 -.00148 49 50 51 52 53 54 55 56 57 58 59 60 61 0.60146 0.65530 0.70751 0.75740 0.80430 0.84755 0.88650 0.92047 0.94881 0.97110 0.98713 0.99678 1.00001 0.00096 0.00311 0.00487 0.00615 0.00688 0.00703 0.00654 0.00546 0.00400 0.00248 0.00119 0.00031 0.00000 0.00027 0.00425 0.01414 0.02991 0.05099 0.07710 0.10798 0.14328 0.18265 0.22570 0.27199 0.32105 0.37238 0.42547 0.47979 0.53477 -.00327 -.01115 -.01701 -.02205 -.02623 -.02948 -.03179 -.03319 -.03367 -.03327 -.03206 -.03010 -.02747 -.02426 -.02064 -.01680 49 50 51 52 53 54 55 56 57 58 59 60 61 0.58975 0.64403 0.69692 0.74768 0.79557 0.83990 0.87998 0.91518 0.94491 0.96864 0.98593 0.99646 1.00001 -.01295 -.00927 -.00593 -.00307 -.00079 0.00087 0.00189 0.00230 0.00216 0.00163 0.00092 0.00027 0.00000 SD7062 SD7062 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.00000 0.99652 0.98634 0.97004 0.94818 0.92127 0.88967 0.85381 0.81427 0.77166 0.72662 0.67973 0.63154 0.58254 0.53322 0.48402 0.00000 0.00057 0.00242 0.00571 0.01036 0.01615 0.02289 0.03049 0.03885 0.04778 0.05702 0.06626 0.07517 0.08347 0.09089 0.09723 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 0.43540 0.38779 0.34159 0.29720 0.25496 0.21521 0.17825 0.14432 0.11363 0.08637 0.06269 0.04272 0.02650 0.01410 0.00562 0.00103 0.10229 0.10592 0.10801 0.10846 0.10722 0.10428 0.09965 0.09340 0.08562 0.07647 0.06616 0.05491 0.04300 0.03081 0.01871 0.00711 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 123 124 Airfoils at Low Speeds c SD7080 1 1.00000 2 0.99671 3 0.98701 4 0.97133 5 0.95018 6 0.92413 7 0.89372 8 0.85943 9 0.82169 10 0.78084 11 0.73720 12 0.69117 13 0.64326 14 0.594Q2 15 0.54399 16 0.49371 0.00000 0.00037 0.00162 0.00395 0.00743 0.01195 0.01728 0.02313 0.02919 0.03519 0.04097 0.04647 0.05163 0.05635 0.06052 0.06404 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 c SD7084 1 1.00000 2 0.99658 3 0.98650 4 0.97026 5 0.94845 6 0.92172 7 0.89064 8 0.85575 9 0.81748 10 0.77619 11 0.73218 12 0.68584 13 0.63770 14 0.58829 15 0.53815 16 0.48783 -- SD7080 0.44370 0.39448 0.34656 0.30040 0.25644 0.21509 0.17670 0.14157 0.10996 0.08210 0.05814 0.03821 0.02237 0,01068 0.00322 0.00003 0.06680 0.06871 0.06969 0.06966 0.06858 0.06641 0.06312 0.05875 0.05334 0.04699 0.03983 0.03203 0.02383 0.01558 0.00763 0.00061 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.00216 0.01016 0.02347 0.04198 0.06549 0.09383 0.12676 0.16397 0.20508 0.24962 0.29710 0.34700 0.39877 0.45183 0.50559 0.55945 -.00492 -.00948 -.01354 -.01694 -.01962 -.02156 -.02279 -.02337 -.02338 -.02289 -.02194 -.02059 -.01891 -.01696 -.01480 -.01250 49 50 51 52 53 54 55 56 57 58 59 60 61 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 0.43785 0.38874 0.34098 0.29505 0.25139 0.21042 0.17254 0.13804 0.10710 0.07986 0.05642 0.03687 0.02132 0.00985 0.00263 0.00000 0.06786 0.06978 0.07074 0.07068 0.06955 0.06734 0.06399 0.05945 0.05374 0.04698 0.03934 0.03109 0.02256 0.01413 0.00631 -.00016 -.01015 -.00783 -.00564 -.00366 -.00201 -.00072 0.00017 0.00069 0.00087 0.00076 0.00047 0.00015 0.00000 --- 507084 0.00000 0.00030 0.00139 0.00361 0.00706 0.01163 0.01708 O.D2311 0.02938 0.03557 0.04152 0.04716 0.05245 0.05727 0.06151 0.06508 0.61281 0.66505 0.71557 0.76376 0.80900 0.85068 0.88820 0.92104 0.94871 0.97079 0.98689 0.99671 1.00001 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.00294 0.01167 0.02575 0.04512 0.06955 0.09876 0.13245 0.17030 0.21190 0.25680 0.30452 0.35455 0.40633 0.45931 0.51289 0.56647 -.00551 -.01031 -.01440 -.01786 -.02071 -.02295 -.02458 -.02560 -.02607 -.02601 -.02547 -.02448 -.02312 -.02143 -.01951 -.01741 49 50 51 52 53 54 55 56 57 58 59 60 61 0.61943 0.67116 0.72105 0.76852 0.81296 0.85379 0.89047 0.92250 0.94942 0.97088 0.98671 0.99661 1.00001 -.01520 -.01298 -.01079 -.00873 -.00685 -.00522 -.00384 -.00272 -.00182 -.00105 -.00039 -.00006 0.00000 Chapter 7: Airfoil Coordinates c ---- 507090 SD7090 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.00000 0.99655 0.98664 0.97113 0.95062 0.92522 0.89500 0.86036 0.82176 0.77972 0.73479 0.68754 0.63860 0.58850 0.53777 0.48692 0.00000 0.00050 0.00219 0.00512 0.00882 0.01284 0.01718 0.02188 0.02691 0.03218 0.03760 0.04304 0.04834 0.05329 0.05773 0.06153 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 c 0.43649 0.38699 0.33890 0.29269 0.24878 0.20759 0.16945 0.13467 0.10352 0.07624 0.05297 0.03384 0.01891 0.00827 0.00196 0.00005 0.06457 0.06674 0.06795 0.06814 0.06724 0.06522 0.06206 0.05780 0.05249 0.04621 0.03907 0.03125 0.02298 0.01458 0.00637 -.00097 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.00345 0.01238 0.02624 0.04514 0.06903 0.09771 0.13087 0.16817 0.20926 0.25371 0.30106 0.35077 0.40227 0.45499 0.50832 0.56166 -.00734 -.01318 -.01834 -.02262 -.02605 -.02873 -.03067 -.03188 -.03240 -.03229 -.03161 -.03046 -.02889 -.02696 -.02472 -.02221 49 50 51 52 53 54 55 56 57 58 59 60 61 0.61442 -.01948 0.66605 -.01658 0.71605 -.01363 0.76381 -.01083 0.80869 -.00829 0.85007 -.00609 0.88735 -.00426 0.92001 -.00279 0.94757 -.00160 0.96974 -.00066 0.98622 -.00009 0.99650 0.00004 1.00001 0.00000 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.00440 0.01370 0.02780 0.04677 0.07058 0.09914 0.13219 0.16941 0.21041 0.25477 0.30202 0.35163 0.40307 0.45576 0.50913 0.56263 -.00749 -.01315 -.01814 -.02225 -.02544 -.02776 -.02929 -.03008 -.03020 -.02969 -.02864 -.02710 -.02514 -.02284 -.02024 -.01744 49 50 51 52 53 54 55 56 57 58 59 60 61 0.61566 0.66757 0.71773 0.76556 0.81047 0.85185 0.88910 0.92170 0.94916 0.97105 0.98700 0.99673 1.00001 508000 SD8000 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.00000 0.99674 0.98711 0.97148 0.95032 0.92413 0.89343 0.85871 0.82042 0.77899 0.73481 0.68831 0.63998 0.59034 0.53991 0.48921 0.00000 0.00030 0.00130 0.00321 0.00607 0.00984 0.01434 0.01936 0.02466 0.03000 0.03521 0.04017 0.04478 0.04894 0.05256 0.05553 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 0.43875 0.38905 0.34062 0.29395 0.24951 0.20775 0.16906 0.13380 0.10229 0.07476 0.05142 0.03238 0.01766 0.00729 0.00136 0.00022 0.05780 0.05929 0.05996 0.05978 0.05872 0.05675 0.05389 0.05012 0.04548 0.04000 0.03377 0.02686 0.01948 0.01194 0.00460 -.00175 -.01459 -.01179 -.00910 -.00662 -.00445 -.00268 -.00132 -.00040 0.00013 0.00032 0.00026 0.00009 0.00000 125 126 Airfoils at Low Speeds c 508020 SD8020 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.00000 0.99646 0.98625 0.97017 0.94885 0.92247 0.89118 0.85538 0.81560 0.77237 0.72627 0.67789 0.62777 0.57647 0.52456 0.47261 0.00000 0.00027 0.00131 0.00330 0.00591 0.00876 0.01188 0.01532 0.01908 0.02308 0.02722 0.03135 0.03535 0.03909 0.04246 0.04536 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 c 0.42116 0.37077 0.32196 0.27523 0.23106 0.18987 0.15203 0.11789 0.08774 0.06179 0.04024 0.02318 0.01065 0.00276 0.00000 0.00276 0.04770 0.04938 0.05034 0.05051 0.04982 0.04824 0.04574 0.04233 0.03802 0.03287 0.02697 0.02041 0.01345 0.00645 0.00000 -.00645 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.01066 0.02319 0.04024 0.06180 0.08774 0.11790 0.15204 0.18987 0.23107 0.27524 0.32197 0.37077 0.42117 0.47262 0.52457 0.57648 -.01345 -.02041 -.02697 -.03287 -.03802 -.04233 -.04574 -.04824 -.04982 -.05051 -.05034 -.04938 -.04769 -.04536 -.04246 -.03908 49 50 51 52 53 54 55 56 57 58 59 60 61 -- 508040 SD8040 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.00000 0.99667 0.98684 0.97093 0.94946 0.92299 0.89205 0.85717 0.81881 0.77739 0.73331 0.68699 0.63893 0.58966 0.53969 0.48953 0.00000 0.00037 0.00162 0.00397 0.00748 0.01208 0.01757 0.02369 0.03018 0.03675 0.04321 0.04942 0.05527 0.06060 0.06530 0.06925 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 0.43970 0.39070 0.34300 0.29706 0.25332 0.21218 0.17398 0.13903 0.10761 0.07994 0.05620 0.03653 0.02099 0.00967 0.00271 0.00004 0.07233 0.07444 0.07552 0.07550 0.07434 0.07201 0.06850 0.06386 0.05814 0.05144 0.04389 0.03565 0.02694 0.01809 0.00938 0.00101 0.62778 -.03534 0.67790 -.03135 0.72629 -.02722 0.77238 -.02308 0.81561 -.01907 0.85539 -.01532 0.89119 -.01188 0.92248 -.00876 0.94886 -.00591 0.97018 -.00330 0.98626 -.00131 0.99647 -.00027 1.00001 0.00000 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.00157 0.00845 0.02120 0.03925 0.06241 0.09046 0.12317 0.16023 0.20125 0.24576 0.29326 0.34324 0.39513 0.44836 0.50232 0.55640 -.00582 -.01097 -.01542 -.01911 -.02202 -.02413 -.02546 -.02609 -.02610 -.02558 -.02456 -.02311 -.02130 -.01921 -.01690 -.01445 49 50 51 52 53 54 55 56 57 58 59 60 61 0.60999 0.66249 0.71327 0.76174 0.80726 0.84921 0.88700 0.92010 0.94803 0.97036 0.98668 0.99665 1.00000 -.01193 -.00943 -.00704 -.00487 -.00301 -.00153 -.00044 0.00026 0.00062 0.00064 0.00044 0.00014 0.00000 Chapter 7: Airfoil Coordinates c:= ~ SPICA SPICA 1 2 3 4 5 6 7 8 9 10 11 12 13 1.00000 0.95000 0.90000 0.85000 0.80000 0.75000 0.70000 0.65000 0.60000 0.55000 0.50000 0.45000 0.40000 0.00000 0.01195 0.02270 0.03275 0.04260 0.05238 0.06190 0.07107 0.07970 0.08773 0.09490 0.10053 0.10415 14 15 16 17 18 19 20 21 22 23 24 25 26 0.35000 0.30000 0.25000 0.20000 0.15000 0.10000 0.08000 0.06000 0.04000 0.02000 0.01000 0.00000 0.01000 0.10583 0.10530 0.10242 0.09690 0.08805 0.07400 0.06634 0.05704 0.04560 0.03042 0.02053 0.00000 -.01280 27 28 29 30 31 32 33 34 35 36 37 38 39 0.02000 0.04000 0.06000 0.08000 0.10000 0.15000 0.20000 0.25000 0.30000 0.35000 0.40000 0.45000 0.50000 -.01510 -.01632 -.01590 -.01564 -.01530 -.01445 -.01360 -.01275 -.01190 -.01105 -.01020 -.00935 -.00850 40 41 42 43 44 45 46 47 48 49 0.55000 -.00765 0.60000 -.00680 0.65000 -.00595 0.70000 -.00510 0.75000 -.00425 0.80000 -.00340 0.85000 -.00255 0.90000 -.00170 0.95000 -.00085 1.00000 0.00000 29 30 31 32 33 34 35 36 37 38 39 40 41 42 0.01250 0.02500 0.05000 0.07500 0.10000 0.12500 0.15000 0.17500 0.20000 0.22500 0.25000 0.27500 0.30000 0.35000 -.01200 -.01800 -.02670 -.03275 -.03685 -.03930 -.04050 -.04085 -.04060 -.03995 -.03890 -.03750 -.03575 -.03185 43 44 45 46 47 48 49 50 51 52 53 54 55 0.40000 0.45000 0.50000 0.55000 0.60000 0.65000 0.70000 0.75000 0.80000 0.85000 0.90000 0.95000 1.00000 WB135/35 WB135/35 1 2 3 4 5 6 7 8 9 10 11 12 13 14 1.00000 0.95000 0.90000 0.85000 0.80000 0.75000 0.70000 0.65000 0.60000 0.55000 0.50000 0.45000 0.40000 0.35000 0.00000 0.01200 0.02310 0.03425 0.04500 0.05510 0.06465 0.07360 0.08175 0.08875 0.09425 0.09790 0.09975 0.10000 15 16 17 18 19 20 21 22 23 24 25 26 27 28 0.30000 0.27500 0.25000 0.22500 0.20000 0.17500 0.15000 0.12500 0.10000 0.07500 0.05000 0.02500 0.01250 0.00000 0.09900 0.09800 0.09660 0.09470 0.09200 0.08825 0.08325 0.07700 0.06915 0.05900 0.04650 0.03000 0.01875 0.00000 -.02775 -.02350 -.01925 -.01532 -.01200 -.00940 -.00750 -.00600 -.00500 -.00400 -.00325 -.00250 0.00000 127 08325 0.02975 -.10225 0.27500 0.00000 0.128 Airfoils at Low Speeds WB140/35/FB WB140/35/FB 1 2 3 4 5 6 7 8 9 10 11 12 13 14 1.12500 0.04075 -.22500 0.60000 0.03675 -.01250 0.65000 0.00000 0.04630 0.02650 -.10450 0.85000 0.45000 0.55000 0.04050 -.03225 ·.03000 0.06950 0.01350 -.09225 0.00000 29 30 31 32 33 34 35 36 37 38 39 40 41 42 0.70000 0.07700 0.70000 0.25000 0.02750 -.04000 0.60000 0.10000 0.01800 ·.05000 0.17500 0.90000 0.03700 -.08830 0.06275 0.01575 -.40000 0.20000 0.03800 -.10000 0.09550 0.10125 0.80000 0.95000 1.01800 -.30000 0.80000 0.40000 0.03275 -.15000 0.01300 ·.75000 0.25000 0.30000 0.10365 0.02275 -.05150 0.35000 -.00250 .75000 0.00500 -.500QO 0.50000 0.02050 -.09725 0.95000 0.09775 0.55000 0.10050 0.04025 -.00250 0.35000 0.02500 0.01250 0.02500 0.12500 0.07500 0.17500 0.04050 -.07100 0.05900 0.00000 -.00800 -.09100 0.22500 0.15000 0.45000 0.27500 0.05000 0.07500 0.03950 ·.01875 0.01025 -.90000 0.17500 0.02475 -.02775 0.08300 0.65000 0.85000 0.01475 0.10450 15 16 17 18 19 20 21 22 23 24 25 26 27 28 0.03950 -.03475 43 44 45 46 47 48 49 50 51 52 53 54 55 0. 00748 -.00659 0.03522 0.01362 0.01200 0.84305 0.54165 -.01470 0.02187 0.00621 0.00885 -.01070 0.58169 0.01274 0.01978 -.02123 0.32064 0.01684 0.39368 0.19912 0.82217 0.00553 -.25556 0.02394 0.90040 0.00000 0.05844 0.17826 0.61490 0.67731 0.09622 0.14185 0.00000 0.01307 0.02780 0.08548 -.00238 0.40164 0.97928 0.03394 -.02806 0.00665 -.23278 0.07520 14 15 16 17 18 19 20 21 22 23 24 25 26 0.01071 0.74751 0.00767 34 35 36 37 38 39 40 41 42 0.59641 0.14456 0.00229 1.03337 23 24 25 26 27 28 29 30 31 32 33 0.07769 0.01676 0.00211 0.98966 0.01024 0.37724 -.69655 0.22347 0.00593 -.04498 0.88509 -.07581 0.79285 -.05531 0.00000 -.01375 0.00460 0.07190 0.07510 0.00123 0.83603 -.76590 12 13 14 15 16 17 18 19 20 21 22 0.00714 -.71380 0.10424 0.29035 0.01570 0.43373 -.97ll0 0.55527 0.64510 0.99603 0.02142 0.62167 -.02908 0.00464 0.00000 0.00991 -.00000 -.85031 0.02485 34 35 36 37 38 39 40 41 42 43 44 0.38540 0.06980 0.99390 0.00000 0.92545 0.72358 0.08578 12 13 14 15 16 17 18 19 20 21 22 0.04364 0.08876 0.04595 0.04165 0.09013 0.03378 -.36835 0.06607 0.01053 -.07430 0.08866 0.25879 0.29820 -.00375 -.06937 0.01506 0.00237 0.00435 0.03506 0.96663 -.02863 0.01887 0.02003 -.98544 -.00169 0.00665 -.01576 0.89405 0.49163 0.89796 0.00546 0.01245 0.00000 0.00884 0.89329 0.49162 -.31864 -.00657 -.13491 0.00000 0.00034 0.01685 0.00417 0.07922 0.04483 0.01618 0.93910 0.00770 0.01619 0.00000 0.00920 0.05238 0.95178 0.07482 14 15 16 17 18 19 20 21 22 23 24 25 26 0.45825 0.00383 0.00721 0.97400 0.00753 0.08444 0.02007 -.05379 0.00382 0.00208 0.02491 0.00968 0.36264 -.01346 0.02286 0.88972 0.02248 -.01289 0.01238 0.02940 40 41 42 43 44 45 46 47 48 49 50 51 52 0.00694 27 28 29 30 31 32 33 34 35 36 37 38 39 0.01570 -.03677 -.02329 0.01848 0.62461 0.00453 0.49234 0.05340 0.00000 0.06429 0.02138 0.07741 0.98829 0.94486 0.00259 -.02951 0.00347 -.00250 40 41 42 43 44 45 46 47 48 49 50 51 52 0.79090 0.01848 0.00764 0.98501 0.00027 -.65555 0.05817 0.92453 0.03800 -.99935 1.00792 0.59105 DFIOI-PT 1 2 3 4 5 6 7 8 9 10 11 12 13 1.00402 -.00670 0.01683 -.09016 0.99331 0.99933 1.07805 0.06409 0.00375 0.94190 -.03647 -.01067 0.00073 -.03831 CLARK-Y-PT 1 2 3 4 5 6 7 8 9 10 11 1.54502 0.00083 1.77287 0.00000 0.02767 0.05931 0.97800 -.00111 0.42578 0.99236 0.03795 0.00000 0.11136 0.00318 0.24816 -.00651 0.00229 -.00385 0.00167 0.17815 -.08215 0.24506 0.11253 0.70417 -.00373 27 28 29 30 31 32 33 34 35 36 37 38 39 0.00083 0.04901 0.00106 0.02811 -.00891 0.96357 0.08685 0.90607 -.24055 -.02590 0.03677 0.05829 0.99740 -.02337 -.08069 0.94727 -.03216 0.06157 0.00000 0.00375 0.00000 -.00000 0.00235 0.32284 0.00587 0.97698 0.88698 0.01366 -.83138 0.78060 -.16135 0.00000 -.84337 -.03356 0.94743 0.00314 0.00568 0.12056 -.03144 0.23528 0.66395 0.83894 0.01896 0.06802 0.00442 0.75184 0.73471 0.00000 0.00130 -.00066 -.08634 0.45718 -.40111 129 .00652 0.07096 0.00117 0.31229 0.84249 0.00553 0.03165 -.66160 0.00056 -.02617 0.70704 -.05262 0.04852 0.03554 0.48012 0.00067 0.00769 0.00625 -.02684 -.08807 0.06617 -.04561 0.01583 0.02528 0.93418 0.01510 0.99746 0.01199 -.99430 0.01143 0.92393 0.02951 0.06774 0.00396 0.03854 0.00247 0.00288 0.00000 0.99975 0.52542 0.78351 0.00326 0.00999 23 24 25 26 27 28 29 30 31 32 33 0.00000 0.53849 -.00062 0.02014 0.13250 0.53354 0.97895 0.05409 0.16999 0.31264 0.14826 -.00000 0.02415 0.38900 0.Chapter 7: Airfoil Coordinates AQUILA-PT 1 2 3 4 5 6 7 8 9 10 11 1.00552 0.95507 0.02690 0.06646 0.60846 0.80749 0.02575 0.06080 0.98967 0.00322 0.04312 -.00994 -.49368 DAE51-PT 1 2 3 4 5 6 7 8 9 10 11 12 13 1.08683 0.07634 0.00700 0.82686 0.86567 0.00000 0.00000 0.81039 0.00000 0.08569 0.03199 0.03436 0.00000 0.00064 0. 00985 0.27519 0.45924 0.63510 -.85245 0.00830 0.00015 1.00456 0.85325 -.00710 25 26 27 28 29 30 31 32 33 34 35 36 0.33658 0.00065 -.94810 .01121 0.99197 -.00221 0.02491 0.24998 0.02447 9 0.00000 0.78023 -.97433 -.00000 13 14 15 16 17 18 19 20 21 22 23 24 0.30177 0.68386 0.00921 0.84996 0.97482 0.04653 0.09833 0.01077 0.71830 -.00184 -.53249 0.01178 0.01865 -.01017 0.00324 0.01202 0.01771 0.64366 0.05787 0.00706 0.00084 0.94238 0.02918 0.04070 0.00253 -.18776 0.00280 5 0.00153 0.90822 0.00000 0.00611 0.13321 0.01572 0.00000 0.01817 0.02823 -.00000 0.20590 0.91451 -.12250 0.02758 0.00000 0.00000 -.00519 0.01476 0.03256 -.94788 0.00873 -.01267 37 38 39 40 41 42 43 44 45 46 0.54395 0.55087 0.00000 13 14 15 16 17 18 19 20 21 22 23 24 0.07312 0.09617 0.04620 0.00000 0.41728 0.69587 0.00023 0.02836 -.00687 0.89354 0.06800 0.00000 0.04747 0.06848 0.00321 0.01607 -.97287 0.35733 0.03500 -.07139 0.48896 0.03461 -.02619 0.01512 0.07552 0.83702 -.08774 0.01255 0.99915 -.14176 0.97647 -.01705 -.02193 -.65450 -.02263 -.00659 -.00000 .16040 0.03747 -.77960 0.04978 0.88932 -.00393 0.13114 0.21567 0.04757 12 0.19187 0.78261 -.95862 0.04951 0.01779 0.92933 0.00930 0.99344 0.00654 6 0.06519 0.80730 0.06100 0.74614 0.09504 0.00227 0.83149 0.08802 -.01087 7 0.45255 0.03587 0.02472 -.00358 0.06~89 E193MOD-PT 1 2 3 4 5 6 7 8 9 10 11 12 1.03902 11 0.05987 0.84695 0.00000 l 1.07866 0.01730 -.37094 0.07602 0.02597 0.02959 -.69649 -.11756 0.01291 0.01611 0.15210 0.02609 -.00764 0.41345 0.00276 1.70203 0.71918 0.11235 0.00203 0.06618 0.7852i -.02182 0.01266 0.01512 0.02758 0.99839 0.64174 0.36710 0.02258 -.53910 0.00000 l 2 3 4 5 6 7 8 9 10 11 12 13 1.49857 0.00152 1.28975 0.44983 0.02915 0.00000 -.07591 0.02935 -.00242 0.36103 0.90833 0.61843 0.85031 -.02433 -.98843 0.02133 0.98684 0.01772 37 38 39 40 41 42 43 44 45 0.04478 0.01992 -.00507 0.00000 0.02087 0.04109 0. -.92050 0.76033 0.06074 0.03504 0.07458 0.00000 0.03580 -.00171 4 0.06602 0.00638 0.07721 0.00180 -.00549 0.00632 0.98762 -.01329 -.04527 0.95585 -.99907 0.27865 0.02853 -.00439 0.90275 0.00000 2 0.01256 0.32115 0.05446 0.03145 -.00000 0.00000 -.91774 -.99758 0.03087 0.00062 3 0.03286 10 0.00000 0.00081 -.99393 0.03718 0.00149 0.01003 -.86791 0.99488 -.02828 -.56072 0.00456 27 28 29 30 31 32 33 34 35 36 37 38 39 0.00805 0.02851 0.01343 0.00122 0.11861 0.07723 0.02840 -.00067 0.57978 0.01043 0.01628 0.96137 0.08943 0.02074 0.02677 -.18323 0.22188 0.00000 0.07923 0.01256 -.01863 -.09505 0.02493 -.01190 0.00069 0.00374 0.00064 0.98520 -.98725 0.00605 -.03383 -.05501 13 0.02292 -.02653 0.07474 0.93023 0.02464 0.00000 27 28 29 30 31 32 33 34 35 36 37 38 39 0.00531 0.76683 0.04178 0.95726 -.01389 0.00019 0.01345 -.00407 -.22483 0.07665 0.29604 0.01556 0.02933 0.45198 0.80273 0.00673 0.46035 DF103-PT E193-PT 1 2 3 4 5 6 7 8 9 10 11 12 1.08223 0.97498 0.24545 0.00573 25 26 27 28 29 30 31 32 33 34 35 36 0.01135 -.09129 0.07718 0.01516 0.01907 0.01242 0.02759 0.62900 -.07308 0.06842 0.89421 -.02613 -.00628 0.99023 0.02525 40 41 42 43 44 45 46 47 48 49 0.07208 0.08034 0.23713 0.06266 14 15 16 17 18 19 20 21 22 23 24 25 26 0.01044 0.07729 0.62046 0.04438 0.05207 0.00000 0.38151 0.03502 -.01514 8 0.38465 0.07269 14 15 16 17 18 19 20 21 22 23 24 25 26 0.01428 0.89347 0.17245 0.98913 -.00346 0.00817 -.93667 -.02792 -.35820 0.03677 -.02234 -.93975 0.77809 0.52020 0.05206 0.00000 0.01107 0.57412 -.05414 0.00739 0.29461 0.71890 -.70790 0.17130 0.130 Airfoils at Low Speeds DF102-PT 0.47468 0.00889 -.02171 40 41 42 43 44 45 46 47 48 49 50 51 0.06029 0.00424 -.02631 0.97224 -.07362 0.00405 0.03645 -.99859 1.00387 0.00049 -.00156 0.01228 0.55877 -.00139 -.04856 0.03764 0.63667 0. 04885 0.00000 0.14102 0.04195 0.49264 E214C-PT 1 2 3 4 5 6 7 8 9 10 11 12 13 1.92625 0.01856 0.97037 -.00141 0.02611 0. -.00000 0.00303 0.01680 0.01539 -.76784 0.80160 0.97958 0.75692 -.59147 0.38189 0.09684 0.88020 0.00680 0.02305 0.98606 0.99284 0.60567 0.02293 0.01014 0.00233 0.00499 0.00721 0.00048 0.15454 -.02653 -.76251 0.80668 0.99873 -.00629 0.78619 0.98277 -.00083 1.03241 0.94593 0.68535 0.00000 0.03718 -.87969 -.69937 -.07676 0.08061 0.00250 0.00596 0.79801 0.01842 37 38 39 40 41 42 43 44 45 46 47 48 0.19372 0.00000 0.05523 0.98881 -.77662 0.01411 0.02863 0.04008 0.02057 0.01086 -.05881 0.46089 0.07584 0.00723 0.97290 0.00013 0.00507 0.01491 -.85807 0.32076 0.01952 0.54246 0.01069 0.44143 0.01110 0.00042 -.00511 0.05300 0.13199 0.00905 0.79263 0.01278 -.02353 0.00311 0.00037 0.00070 0.00846 0.00407 0.02114 0.59339 0.87890 0.58021 E214A-PT 1 2 3 4 5 6 7 8 9 10 11 12 1.21066 0.08039 0.02077 -.37788 0.03228 0.21392 0.08847 0.44197 0.00410 0.01654 0.64824 -.96215 0.14521 0.00816 -.00033 -.00783 0.02647 0.08976 0.00317 0.00085 0.01289 0.04043 0.00430 0.00303 0.00000 -.01291 27 28 29 30 31 32 33 34 35 36 37 38 39 0.20528 0.32690 0.00763 0.08894 0.00310 -.00230 0.01429 0.00000 0.01343 0.00968 0.00337 0.92076 -.99974 0.01119 0.02002 -.06478 0.09096 0.01499 0.38703 -.13766 -.91866 0.71377 0.38814 0.00063 -.33678 0.00572 0.85231 -.91581 0.03257 0.08756 0.70930 0.00852 0.00000 0.02659 -.47023 0.00000 27 28 29 30 31 32 33 34 35 36 37 38 39 0.07150 0.03498 0.77621 -.00535 0.98588 0.00391 -.02223 -.01310 -.02546 -.00063 1.04265 0.08723 13 14 15 16 17 18 19 20 21 22 23 24 0.90743 0.99118 0.49880 0.99352 -.05912 0.02078 0.99926 0.05142 0.00223 0.02370 0.69869 0.02150 -.00246 0.07496 0.30797 0.48264 0.00000 0.06127 15 16 17 18 19 20 21 22 23 24 25 26 27 28 0.01480 0.04741 0.99537 0.06613 0.06458 0.02260 -.71106 0.00000 0.06017 0.06897 0.00831 0.00068 0.00000 E205B-PT 1 2 3 4 5 6 7 8 9 10 11 12 13 14 1.06052 -.92416 -.02165 0.01258 43 44 45 46 47 48 49 50 51 52 53 54 0.00572 25 26 27 28 29 30 31 32 33 34 35 36 0.40729 0.82133 -.00430 0.08023 0.00217 0.01119 -.01166 0.55176 0.61533 0.02537 -.00530 0.00000 0.07273 -.00000 0.01253 0.00000 0.01344 0.04164 0.07982 0.07572 0.06792 14 15 16 17 18 19 20 21 22 23 24 25 26 0.01977.00099 0.73829 0.22884 -.64768 0.64112 40 41 42 43 44 45 46 47 48 49 50 51 52 0.01762 0.01274 0.08968 0.04006 0.00000 0.00675 -.02449 0.10372 0.26549 0.86106 0.Chapter 7: Airfoil Coordinates E205A-PT l 2 3 4 5 6 7 8 9 10 11 12 13 1.00053 -.98546 -.78073 0.02927 0.00813 0.07663 0.08003 0.00000 0.01143 0.02021 0.95562 0.00000 -.03272 0.71250 -.00719 0.99392 0.15044 0.00876 0.75295 0.05802 14 15 16 17 18 19 20 21 22 23 24 25 26 0.57345 -.94617 0.20297 -.01526 0.26049 0.00044 0.00291 0.98921 0.00614 29 30 31 32 33 34 35 36 37 38 39 40 41 42 0.81466 0.74147 0.01012 0.00048 0.00681 0.26612 0.05801 0.40011 0.00749 0.83393 0.00258 -.02275 0.00197 0.00000 0.03315 0.03254 0.95318 0.06849 0.65865 0.06536 0.98991 -.99173 0.03006 -.00709 -.00000 0.80921 0.02168 0.84045 0.56883 -.24597 0.00000 0.07291 0.49886 0.05923 0.05061 0.01639 0.90880 0.00146 -.96116 -.05474 0.02076 0.02611 0.05440 0.91110 0.75150 0.34581 -.53232 0.08367 0.09479 -.01469 0.53834 0.00929 0.05939 0.13117 0.01748 0.04985 0.02548 0.04302 0.06079 0.01110 0.44292 0.08233 0.00234 0.06230 0.00000 0.07548 0.74722 0.26877 -.42989 -.01156 0.00361 0.00089 -.97156 0.00505 0.02210 0.98393 0.05249 0.06738 0.01676 40 41 42 43 44 45 46 47 48 49 50 0.04462 0.01879 -.99485 -.00419 0.00000 0.99848 1.82377 0.73699 0.98578 0.04742 0.96317 0.08344 0.00106 0.06819 0.08863 0.00150 0.08186 0.00334 0.99758 -.00620 0.00289 0.01395 -.28427 0.00143 0.31157 -.68560 0.01864 0.00418 -.32774 0.52519 0.01807 -.08093 0.99116 1.63333 0.00307 0.85444 0.02413 0.04368 0.81510 0.00192 131 .63816 -.86509 0.09241 0.13152 0.00000 0.04717 0.49449 -.00767 -.01003 0.00000 0.20579 0.02360 -.19136 0.99655 0.00152 0.00159 0. 24611 -.00146 0.42928 0.66477 0.81249 0.04103 0.01471 -.00174 25 26 27 28 29 30 31 32 33 34 35 36 0.92373 0.99768 -.00082 0.67132 0.95748 0.07109 13 14 15 16 17 18 19 20 21 22 23 24 0.00000 0.89877 -.02648 0.00034 0.07379 0.23526 0.95261 -.00171 0.00000 0.00504 0.00344 0.00888 0.03114 -.03312 0.00999 0.02603 0.69039 -.00694 0.06483 0.07624 0.01950 37 38 39 40 41 42 43 44 45 0.07283 13 14 15 16 17 18 19 20 21 22 23 24 0.04423 0.06574 0.99336 0.00219 0.97104 0.49001 0.05443 0.01573 -.06553 0.07807 0.00000 0.05900 0.07840 0.04991 0.02052 0.03024 -.00380 -.00086 1.64244 0.00000 0.28899 0.00050 0.00489 -.00380 0.15093 0.98133 -. 0.86350 0.99956 0.00203 0.36181 0.03349 -.08724 0.07326 0.00074 0.02246 0.99004 0.05724 0.56905 -.05618 0.01339 0.51772 0.04153 0.02782 -.07685 0.37263 0.04315 0.00129 0.00000 -.01471 -.00650 0.99570 -.01408 0.05709 0.06283 0.02726 0.06652 0.01982 -.32904 0.90475 0.00376 -.00137 0.03260 -.00000 0.02810 0.99099 0.02161 -.99924 0.00066 0.00013 -.95128 0.97517 0.00000 E374B-PT 1 2 3 4 5 6 7 8 9 10 11 12 1.01243 -.63671 E387A-PT 1 2 3 4 5 6 7 8 9 10 11 12 13 1.03663 0.98292 -.00091 -.85747 0.37453 0.00730 37 38 39 40 41 42 43 44 45 46 47 0.02254 0.98425 0.05012 0.00098 0.00528 0.99993 0.00098 0.00252 0.00381 -.70352 0.00006 0.71691 0.03752 0.00000 -.01023 -.00000 0.07237 0.00256 0.50423 0.59367 0.96395 -.00244 0.14024 0.08539 0.00474 0.06883 0.76815 0.86019 -.26789 0.01135 0.21432 0.95595 0.87735 0.00022 0.50854 0.89135 0.03532 0.06062 14 15 16 17 18 19 20 21 22 23 24 25 26 0.00000 0.01736 0.30895 0.05655 0.83110 0.75201 0.00000 0.81443 -.25364 0.67810 -.00983 0.00000 0.02791 0.00023 0.00000 0.01040 -.86352 0.00657 27 28 29 30 31 32 33 34 35 36 37 38 39 0.77048 -.00357 0.00893 -.00200 .02657 -.07677 0.01116 0.17031 0.29099 0.91962 0.08786 -.99789 1.91799 0.42894 0.00033 0.01233 0.03170 0.04263 0.58767 E387B-PT 1 2 3 4 5 6 7 8 9 10 11 12 1.99533 0.02989 0.99772 -.00219 0.00935 0.04590 0.00617 -.51616 0.01771 0.32673 -.07679 0.19283 0.04485 0.71808 0.01036 0.00170 -.02177 0.29601 0.03356 -.09227 0.63885 0.00583 -.01859 -.00909 0.03294 -.01734 0.01115 -.35996 0.00105 0.07653 0.00097 0.45910 0.41722 0.01882 0.00090 0.74877 0.00313 -.00454 0.00295 -.00000 0.05454 0.61172 -.00125 0.53883 -.99939 0.05501 0.75981 0.01407 0.06899 0.03539 0.01131 0.00275 0.00540 -.00000 0.00000 0.00503 0.37851 0.04642 0.01485 -.47124 0.01837 0.00000 -.05087 0.02039 0.11571 0.00000 0.05179 0.00000 0.00150 0.87744 0.00000 0.15840 0.04674 0.03719 -.03570 0.00548 -.65172 0.132 Airfoils at Low Speeds E374A-PT 1 2 3 4 5 6 7 8 9 10 11 12 1.01825 0.70153 0.00669 -.13467 0.76066 0.97236 0.01244 0.00782 -.00237 0.93069 0.87169 -.45261 -.01173 0.43932 0.00192 -.00540 0.00069 -.10183 0.22957 0.01463 0.74960 -.92030 -.99546 0.00700 0.19350 0.01874 -.07440 0.00000 0.00017 0.82139 0.99096 -.01061 -.01222 0.00000 -.02133 0.00448 0.55808 0.00096 1.02689 0.03128 -.22554 0.02386 -.01312 0.98514 0.05230 0.09328 0.00284 25 26 27 28 29 30 31 32 33 34 35 36 0.00343 0.02669 -.98821 0.04501 0.06943 0.98178 0.25247 0.80627 0.00938 0.00927 0.98587 0.02928 -.02037 -.00680 40 41 42 43 44 45 46 47 48 49 50 51 52 0.02851 0.14550 0.03114 -.82622 0.02218 -.00058 0.95621 0.02082 -.33725 0.99182 -.70590 0.07226 0.99459 0.00541 0.94168 0.00097 0.565~4 0.99700 1.00106 0.01467 0.06005 13 14 15 16 17 18 19 20 21 22 23 24 0.80195 0.03197 0.00234 0.02253 0.02446 0.03042 -.00717 0.02327 -.00869 0.00379 25 26 27 28 29 30 31 32 33 34 35 36 0.00339 0.00083 0.98369 0.19237 0.00305 0.00000 -.00108 0.12583 0.02915 37 38 39 40 41 42 43 44 45 46 47 0.41827 0.01442 0.02563 0.00191 0.56611 .01180 0.07836 0.14326 -.99430 0.00000 0.07641 0.02077 0.59964 0. 70060 0.00438 0.99654 1.01050 0.46829 0.06063 0.83405 0.25878 0.50757 -.09386 0.01148 0.10000 0.36648 0.01357 -.00000 0.00321 0.07737 -.01755 0.01728 -.11343 0.00171 31 32 33 34 35 36 37 38 39 40 41 42 43 44 45 0.48085 0.85934 0.76242 0.01990 -.23771 0.00080 0.06269 0.01623 0.84243 0.11115 -.28428 0.57619 0.02577 -.01727 0.03461 0.00021 0.01110 -.02641 0.05209 0.00092 0.02247 0.70945 0.63881 0.11760 14 15 16 17 18 19 20 21 22 23 24 25 26 0.09767 0.98204 0.221 0.01140 -.00295 -.74410 0.01578 -.01261 0.74795 0.02901 -.03577 0.96638 0.00106 0.04675 0.01722 0.00525 0.01833 0.65128 0.82510 0.00017 0.00036 0.00000 0.59521 0.99601 1.00638 0.90673 0.99840 0.00989 0.99817 1.34963 0.10530 14 15 16 17 18 19 20 21 22 23 24 25 26 0.84235 0.05017 0.12658 0.01134 0.01477 0.03047 0.00596 37 38 39 40 41 42 43 44 45 0.00549 0.00000 0.92908 0.01070 0.11855 0.00044 0.86695 0.96502 0.00027 -.01032 0.00215 0.02728 0.00273 0.00661 0.59311 0.06243 133 .02754 0.65019 0.02520 -.57258 0.06262 0.01889 0.00000 0.01489 0.06743 0.00000 -.94303 0.00403 0.02127 -.04396 0.74790 0.12035 0.12111 0.00085 -.00638 -.01627 0.02331 0.78942 0.01684 0.03103 -.11918 0.11902 0.02258 -.17088 0.05743 0.00454 0.20042 -.97848 0.00000 0.95073 0.00000 -.24832 -.02335 0.00044 0.02841 -.00233 0.67207 0.74808 0.00369 0.02797 -.97660 0.17539 0.05437 0.02811 46 47 48 49 50 51 52 53 54 55 56 57 58 59 60 0.01227 0.00000 0.01704 -.00000 -.01920 0.15725 -.00468 0.04377 0.02833 0.00000 HQ2/9A-PT 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 1.34245 0.99004 0.00624 -.19079 0.53463 -.01975 0.11925 0.97472 0.99003 0.02406 -.12022 -.31387 0.29955 0.98807 0.03059 0.97732 0.00000 0.05724 0.04864 0.01304 0.53476 0.04291 0.10.08163 0.92836 0.00320 0.60394 0.06001 0.07591 0.99535 0.79234 0.04263 0.00463 25 26 27 28 29 30 31 32 33 34 35 36 0.00723 0.64950 0.18910 0.00000 FX63-137B-PT (local slight bulge on lower surface near endplate) 1 2 3 4 5 6 7 8 9 10 11 12 13 1.06990 -.02696 -.02044 0.53267 0.92032 0.05027 0.02478 0.02633 0.02927 -.01669 -.00153 -.00031 0.00018 -.00034 -.37828 0.12072 0.00000 -.03783 0.07505 0.01208 40 41 42 43 44 45 46 47 48 49 0.09294 0.97998 0.01479 0.00000 0.00142 0.00015 40 41 42 43 44 45 46 47 48 49 50 51 52 0.40805 0.95596 0.88623 0.02205 0.00000 0.01377 0.00562 -.53121 0.01459 0.00716 0.37524 0.02992 -.00082 0.40448 0.00000 16 17 18 19 20 21 22 23 24 25 26 27 28 29 30 0.04352 0.00953 -.35046 -.02616 0.01977 0.45272 0.03727 0.00679 0.04846 -.97197 0.11169 0.00236 0.42369 -.08367 0.01159 -.00682 0.00933 0.00762 0.02781 0.88326 0.00634 0.07508 13 14 15 16 17 18 19 20 21 22 23 24 0.02898 0.02559 0.11598 0.10971 0.79653 0.77859 0.06153 0.07207 0.42133 0.63054 0.25427 0.00364 0.07273 0.02457 0.88226 0.02773 0.03794 0.49431 0.98950 0.74843 0.99524 0.00083 -.00000 0.83638 0.00376 0.03516 0.07372 0.00160 0.46945 0.02330 -.90432 0.04045 0.11363 0.96160 0.23448 0.95148 0.99158 0.00000 -.00333 0.70056 0.05472 0.21585 0.01147 -.02021 0.03606 -.00000 0.00095 0.02178 0.03103 0.00954 0.28758 0.97458 0.43566 0.87379 0.08139 0.99688 0.01382 0.00165 0.80264 0.00118 0.74044 0.06835 0.36884 -.02414 0.00599 0.08959 0.00900 0.00614 -.00376 0.91782 0.01202 0.03310 0.00047 0.98818 0.99093 0.00914 0.02029 -.68583 0.01889 0.00000 0.45637 -.91819 0.69008 0.86207 0.00000 0.12362 0.01549 0.00387 0.05353 0.00000 FX63-137B-PT (these coordinates used in plots given in Chapters 10 and 11) 1 2 3 4 5 6 7 8 9 10 11 12 13 1.56897 0.17811 -.01780 0.17228 0.08408 0.03192 0.01857 -.99813 0.05552 0.07704 0.93436 0.00336 -.02129 0.04375 0.00654 27 28 29 30 31 32 33 34 35 36 37 38 39 0.01074 -.06850 0.07484 0.07127 0.06298 0.02323 0.00013 -.05880 0.99028 1.03793 0.00244 0.00124 0.06558 0.14628 0.00861 27 28 29 30 31 32 33 34 35 36 37 38 39 0.02089 0.00675 -.00000 0.60357 0.00796 0.04171 0.02692 0.34272 0.28516 -.82271 0.28888 0.79130 0.02006 0.03726 0.01416 0.53717 0.02140 0.04776 0.69668 0.02512 -.00000 0.00141 0.00997 0.01323 -.03054 -.03127 -.Chapter 7: Airfoil Coordinates FX60-100-PT 1 2 3 4 5 6 7 8 9 10 11 12 1.49999 0.00688 0.23277 0.01154 0.42888 0.94758 0.65564 0.01432 0.00378 0.09972 0. 02720 -.06416 0.11702 0.10483 0.05444 -.02482 0.00000 J5012-PT 1 2 3 4 5 6 7 8 9 10 11 1.00779 -.05852 -.05061 -.02788 -.00083 0.00402 0.03726 -.99710 0.71853 0.01039 0.00072 0.23778 0.64657 0.26780 0.97090 0.00110 0.01121 -.03477 0.02401 0.63080 0.02729 -.05897 0.00286 0.64248 0.79569 0.05482 0.01185 -.99875 0.99113 0.99224 -.71886 -.05729 -.00209 0.00000 0.99302 0.16643 0.01046 -.12338 0.32424 0.18302 0.32312 0.50314 0.20474 0.92077 -.01519 -.00229 0.84803 0.01083 0.05744 0.00098 0.94663 -.09456 M06-13-128-PT 1 2 3 4 5 6 7 8 9 10 11 12 1.00218 -.03835 0.08380 0.50773 0.02643 0.00342 0.05222 0.00952 0.01263 -.98737 0.09933 0.00103 0.01795 -.65407 0.00837 -.10986 0.95164 0.01410 -.00078 0.99873 0.05474 13 14 15 16 17 18 19 20 21 22 23 24 0.04989 34 35 36 37 38 39 40 41 42 43 0.98163 0.95851 0.00000 .00000 0.56234 -.89026 0.01533 -.00046 -.11054 0.05829 0.41197 0.01678 0.05673 0.03509 0.71117 0.01102 -.01884 0.91306 0.79608 0.59102 -.85319 0.34297 0.00426 0.01061 0.90593 0.05766 -.08585 0.16957 0.04944 -.74939 0.97217 0.00000 0.00743 0.00549 0.03797 0.01093 0.84383 0.00068 0.03004 0.01000 0.01037 0.34668 0.03944 -.04476 0.05650 0.05965 0.03772 0.02430 0.00380 -.37864 0.87289 0.02655 0.00166 0.09369 0.00565 0.00000 -.04435 0.01306 -.00000 0.78393 0.99856 -.00531 0.62803 0.06139 0.00405 0.03376 0.03082 0.02068 0.00375 0.05972 0.00000 0.01439 -.02874 -.25128 0.04321 0.52797 0.03546 0.24036 0.97999 0.07332 0.00586 0.00665 0.00006 -.02464 -.00112 -.23425 0.01460 -.06272 0.04303 0.01354 0.93888 0. 70325 0.45504 0.95193 0.29318 0.47429 0.00000 0.00747 0.05945 0.67584 0.01982 0.98068 0.00000 0.04146 0.07307 0.00286 0.67530 -.79724 -.84769 -.02874 -.01325 0.91330 0.04821 0.99839 0.98601 0.01337 0.00843 0.02712 0.95825 -.04517 -.00473 0.11754 0.02849 0.86129 -.04844 0.04791 0.05717 0.46570 0.00416 0.99141 0.00000 -.00137 0.04588 -.00539 -.00000 0.12940 0.03266 0.00415 0.04992 12 13 14 15 16 17 18 19 20 21 22 0.00067 -.59104 0.77219 -.01267 0.98771 0.49340 -.00976 -.02158 -.05887 0.40631 0.53876 0.01910 0.10041 0.02234 0.05797 0.93095 0.99320 0.99800 -.09563 0.99127 -.01822 -.01197 0.08641 0.00000 0.99542 1.09826 0.08761 0.00000 0.06305 0.97221 0.00387 25 26 27 28 29 30 31 32 33 34 35 36 0.43053 0.10527 0.06271 0.00741 -.36556 0.00000 13 14 15 16 17 18 19 20 21 22 23 24 0.53817 0.00577 0.11150 0.76389 0.31208 0.57332 MB253515-PT 1 2 3 4 5 6 7 8 9 10 11 1.01340 0.49443 0.18367 0.00722 0.13255 0.04183 25 26 27 28 29 30 31 32 33 34 35 36 0.00236 1.00026 1.00009 0.95248 0.05295 -.86592 0.06064 0.28317 0.01770 0.00000 0.58573 0.03292 0.00284 -.88089 0.01716 0.00880 -.00103 0.00000 23 24 25 26 27 28 29 30 31 32 33 0.01815 0.00000 12 13 14 15 16 17 18 19 20 21 22 0.00627 -.03590 0.14394 0.11964 0.00018 0.08764 0.02140 37 38 39 40 41 42 43 44 45 46 0.03642 0.01929 0.41617 0.01613 -.00035 -.45124 0.08438 0.57808 0.33028 0.07291 -.03782 0.00000 23 24 25 26 27 28 29 30 31 32 33 0.01360 37 38 39 40 41 42 43 44 45 46 47 48 0.01695 0.03290 0.00226 0.04751 -.70484 0.00000 0.81308 0.40392 0.00018 -.00000 0.54312 0.22462 0.99294 1.19551 0.64288 -.01004 0.81780 0.19537 0.05123 -.01665 0.02877 -.00698 0.07353 0.41644 0.00249 0.00190 0.00000 0.00000 -.00669 -.11643 0.12578 0.02552 -.26557 0.90082 -.00255 0.00961 -.97990 -.02427 0.04204 34 35 36 37 38 39 40 41 42 0.00000 0.69189 0.04962 0.32895 0.09873 0.98225 -.03074 0.76178 0.00787 0.134 Airfoils at Low Speeds HQ2/9B-PT 1 2 3 4 5 6 7 8 9 10 11 12 1.49316 0. 00729 0.98608 0.01767 0.02301 -.42977 0.04647 0.69948 0.10438 0.04818 0.00195 -.58407 -.01785 0.04444 -.07337 0.01095 0.00000 0.00050 0.49536 0.47705 0.96862 .05290 0.00617 0.30559 0.01565 0.01308 -.01442 0.02567 0.00039 -.00000 135 .01904 -.04961 -.01617 0.00668 0.62605 0.08205 0.03344 -.36124 -.03601 0.01291 0.00276 0.01832 43 44 45 46 47 48 49 50 51 52 53 0.00353 -.00660 0.03427 -.00504 0.00000 0.96609 0.05676 0.01920 0.00000 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0.09281 15 16 17 18 19 20 21 22 23 24 25 26 27 28 0.00519 25 26 27 28 29 30 31 32 33 34 35 36 0.23126 0.58733 -.01772 0.01569 0.04526 0.00376 -.04961 0.66517 -.94328 0.03794 -.01239 0.62493 0.90758 -.00068 0.82463 0.09709 0.94936 0.06010 0.02918 0.00105 -.06637 0.02049 37 38 39 40 41 42 43 44 45 46 47 0.00220 0.05484 o.00103 0.01758 0.08912 0.00034 0.02177 -.01258 0.00030 0.10279 0.99208 0.00225 -.00171 0.41589 -.05274 0.00430 0.95407 0.05788 0.41487 0.01995 -.00135 0.01579 0.00819 0.98292 0.87370 0.20304 0.01042 0.00087 0.04255 0.00132 0.06183 0.03831 0.89008 0.06348 0.40654 0.02586 -.66435 0.00944 0.98223 -.02955 0.00349 0.81666 -.85668 0.01811 -.99529 0.00528 -.00356 0.01407 -.90309 0.33724 0.43410 0.00112 0.01738 -.00000 0.2 0.06302 0.01617 -.01560 0.00080 -.99147 0.00000 0.28748 0.00000 0.01436 0.00000 -.03573 -.Chapter 7: Airfoil Coordinates 84233-PT 1 2 3 4 5 6 7 8 9 10 11 12 13 14 1.00265 0.01840 0.00949 -.07015 0.50216 -.02207 0.01468 0.03223 -.00705 0.17709 0.76889 -.10215 0.38078 0.00405 0.00058 -.00344 29 30 31 32 33 34 35 36 37 38 39 40 41 42 0.01331 -.85117 0.09805 0.00082 0.80341 0.14375 -.00111 0.33847 0.85175 0.99965 0.00665 0.09827 0.08089 0.98124 0.03498 0.46493 -.99772 1.01064 0.63022 -.00000 0.00869 29 30 31 32 33 34 35 36 37 38 39 40 41 42 0.06460 0.14049 0.99937 -.90598 0.03031 -.01505 -.00755 0.99655 -.00400 -.73356 0.06495 0.00114 0.01768 0.03861 0.02708 0.01964 -.27437 0.03195 43 44 45 46 47 48 49 50 51 52 53 54 0.74226 0.00000 -.00054 -.99836 1.01408 0.06515 0.77953 0.00000 0.02344 0.03299 0.00111 -.56063 -.04985 0.06299 0.06735 0.99136 0.01851 0.00000 0.02310 0.54575 0.00258 0.02719 0.03884 -.63304 0.00780 -.00452 0.46205 0.39894 0.02533 0.01080 0.58574 0.06448 0.09139 0.99965 0.05695 0.05843 0.09223 -.12159 0.00000 -.16431 0.26237 0.04537 0.02403 0.01015 -.85409 -.25685 -.84138 0.6890.02268 -.70793 -.98044 -.02697 0.01967 -.00000 13 14 15 16 17 18 19 20 21 22 23 24 0.02929 0.92852 0.03244 0.00082 -.00743 0.31977 0.98943 0.96157 -.93306 -.49512 0.01286 0.00373 0.00366 25 26 27 28 29 30 31 32 33 34 35 36 0.01225 -.28714 0.00617 -.06730 0.00000 0.03787 -.03583 0.99026 0.00690 0.02080 0.00000 -.89226 0.03888 -.23863 0.05125 0.90526 0.01435 -.00095 1.36002 0.00127 0.01151 -.02201 0.02054 0.63091 0.53047 0.45243 0.00000 0.00291 0.13961 0.59630 0.06059 SD2083-PT 1 2 3 4 5 6 7 8 9 10 11 12 13 14 1.00784 0.02774 0.05681 0.19208 -.03053 -.00151 -.00274 0.98879 0.00423 0.99594 0.00794 -.19371 0.00197 -.99670 0.01087 43 44 45 46 47 48 49 50 51 52 53 0.00309 .00667 0.98663 0.95825 0.00534 37 38 39 40 41 42 43 44 45 0.02635 -.00000 0.01368 0.01073 0.81957 0.05292 0.04663 -.00000 13 14 15 16 17 18 19 20 21 22 23 24 0.00000 0.01948 0.02170 0.02381 0.02369 0.00072 0.00801 0.00335 -.00348 0.02240 -.00023 0.65583 -.00183 0.-.74591 0.00061 0.16785 0.97366 0.01462 0.00027 0.52135 0.00062 0.99484 0.D1075 -.88423 -.00157 0.00848 0.04386 0. 00170 0.01605 -.00079 0.01429 -.02758 -.26720 0.00000 -.04250 0.02325 -.98072 0.87412 0.00464 0.02664 0.02226 10 0.03165 -.Oll78 0.03023 0.00397 0.06533 0.06833 0.00000 12 13 14 15 16 17 18 19 20 21 22 0.00124 -.04562 .87960 0.01018 0.140 Airfoils at Low Speeds SD6060-PT 1 1.01721 34 35 36 37 38 39 40 41 42 43 0.48053 0.00062 0.97337 0.81523 0.00581 29 30 31 32 33 34 35 36 37 38 39 40 41 42 O.Oll77 -.03767 0.00103 -.92005 0.00221 -.98613 0.00632 -.00341 0.95984 0.02146 0.00028 0.06849 0.00274 0.63386 -.00894 0.00988 0.97953 0.00981 0.05552 0.00000 -.02873 -.04492 0.88784 0.00240 -.00000 23 24 25 26 27 28 29 30 31 32 33 0.98941 0.01381 0.00223 -.39456 0.03378 0.08254 0.58943 0.03325 -.00538 0.07984 0.69393 0.00000 SD7003-PT 1 2 3 4 5 6 7 8 9 10 ll 1.00063 0.01093 -.04614 0.17601 0.08323 0.06850 0.02566 -.00160 0.03265 40 41 42 43 44 45 46 47 48 49 50 51 52 0.00836 -.29043 0.77190 0.00860 8 0.07121 0.34746 0.00782 0.07702 0.99843 1.02092 0.00031 -.00369 7 0.00565 -.01582 -.00161 -.00321 0.02649 0.00639 -.65704 0.54297 0.23356 0.05667 0.00008 0.02599 -.00145 0.25357 0.01231 0.00174 0.02738 0.00150 -.00564 0.99144 0.00000 1.69444 0.05214 0.02324 -.00000 0.04066 0.71977 0.00ll6 -.01088 -.06500 0.01202 -.00622 -.00000 0.99328 0.15269 0.03079 0.01496 0.17602 0.00059 0.01323 0.00000 2 0.00913 0.00000 0.76263 0.00000 0.02082 0.00200 0.95181 0.99071 0.Q7083 15 16 17 18 19 20 21 22 23 24 25 26 27 28 0.99790 1.35759 0.40505 0.93863 0.33337 0.67737 0.99686 0.00000 0.72015 0.03416 0.Oll77 -.07606 0.03807 0.01039 0.01690 0.00244 0.99563 0.30489 0.01727 0.03775 0.08137 0.02346 0.02921 -.05335 14 15 16 17 18 19 20 21 22 23 24 25 26 0.99338 0.12326 0.02940 -.00968 0.03821 0.00099 0.01420 -.00346 0.99352 0.00ll2 4 0.00273 0.00029 -.01685 -.01620 0.00000 0.42930 0.01681 0.91167 0.04650 13 0.06446 0.66197 0.49908 0.95190 0.32396 0.00389 -.00721 -.17139 0.08045 0.18604 0.05525 0.00000 -.04465 0.01446 -.01716 -.00169 0.05239 0.02923 0.00046 0.00100 -.27971 0.89391 0.05548 0.09914 0.01483 -.20186 0.00854 0.14632 0.04437 0.96772 0.92852 0.75444 0.86546 0.99980 0.01292 -.98356 0.46829 0.05615 0.76495 0.00850 -.06425 O.00282 0.03269 -.78327 0.08699 0.12224 0.00014 -.01465 0.98784 0.41383 0.00890 -.98023 0.16377 0.05089 0.23877 0.02199 -.63891 0.00000 0.00000 SD6080-PT 1 2 3 4 5 6 7 8 9 10 11 12 13 14 1.97733 0.00578 27 28 29 30 31 32 33 34 35 36 37 38 39 0.01498 0.91340 0.54167 0.00144 0.00545 0.07070 O.21314 0.59273 -.70362 0.58070 0.09126 0.60244 0.02050 0.00184 -.99485 0.03369 0.06175 0.94524 0.00102 0.10826 0.03342 0.00318 0.00000 29 30 31 32 33 34 35 36 37 38 39 40 41 42 0.36166 0.52933 0.00180 0.59815 0.00000 0.96550 0.00006 0.05792 0.05997 0.02753 -.69134 0.39060 0.98981 0.71077 0.00629 0.34441 0.07663 0.00434 -.00093 0.00155 0. 75386 0.53027 0.62282 0.47523 0.00000 0.01049 0.03027 -.00024 0.31688 0.00091 0.008ll -.00000 0.62ll8 0.00346 0.01594 0.03055 0.26926 0.96319 0.99494 0.04409 0.24528 0.99878 1.21524 0.99066 0.06569 0.82285 0.00104 0.46650 0.81964 0.ll221 0.01521 0.02298 0.01401 9 0.07665 0.00608 -.19891 0.02058 -.00990 -.08189 0.07682 SD6080-PT with thickened trailing edge 1 2 3 4 5 6 7 8 9 10 11 12 13 14 1.08341 0.03132 11 0.53137 0.00163 0.99861 0.67179 0.00756 0.00355 0.82496 0.00000 15 16 17 18 19 20 21 22 23 24 25 26 27 28 0.00096 -.00931 -.13981 0.00270 -.00363 0.03156 -.00077 43 44 45 46 47 48 49 50 51 52 53 54 55 0.36452 0.04949 0.00166 43 44 45 46 47 48 49 50 51 52 53 0.00392 -.90699 0.00082 0.05936 0.99835 0.86566 0.75307 0.01503 -.00167 0.00000 0.00787 0.00214 6 0.02692 0.01127 -.91623 0.00400 0.06740 0.00058 0.80014 0.65836 0.93895 0.01172 -.81591 0.05605 0.00502 -.83534 0.00320 0.05708 0.03681 0.03920 12 0.97956 0.00047 -.85870 0.12918 0.00335 0.99844 1.99682 0.05187 0.40646 0.00000 -.27994 0.00074 3 0.00159 0.00438 0.46417 0.54565 0.56973 -.00517 0.03438 0.05970 0.00229 -.00364 -.01316 -.00128 5 0.45370 0.46848 0.74439 0.00000 -.10445 0.88111 0. 81776 0.81578 0.03021 -.01517 0.02924 0.96051 0.93450 -.01377 0.00781 -.00125 0.02639 0.03803 0.26818 0.63904 0.00192 0.69316 0.02037 -.00342 0.01557 0.01460 -.54410 0.99661 0.02489 -.32999 0.15204 -.67726 -.47549 0.00327 0.00716 -.04036 14 15 16 17 18 19 20 21 22 23 24 25 26 0.607lll 0.99777 1.20352 0.04341 0.79496 -.00275 0.02569 0.00235 .repeated) 1 2 3 4 5 6 7 8 9 10 11 12 13 1.02895 -.95540 0.00986 -.00000 0.05639 0.00000 SD7003-PT (A .02222 0.00035 -.00167 0.08120 0.95169 0.05435 0.01788 27 28 29 30 31 32 33 34 35 36 37 38 39 SD7003-PT (C.00000 0.02893 -.92515 0.01591 0.02692 -.41625 0.01329 -.01067 0.20027 0.00597 -.01578 0.05214 0.88421 -.04844 15 16 17 18 19 20 21 22 23 24 25 26 27 28 0.02196 -.99228 0.01082 0.62788 0.00000 0.01086 0.03137 0.62705 0.05645 0.03580 0.02164 141 .03013 -.54957 0.37050 0.99029 0.03367 0.01659 -.02692 0.59240 0.99156 -.94989 0.16565 0.02971 0.05504 0.06677 0.02792 0.97461 0.34123 0.70836 0.00000 0.62939 0.00256 0.00000 0.03228 0.83383 0.41796 0.99177 0.68151 0.00000 -.00249 0.01802 -.00000 0.04324 0.00135 -.00073 -.14505 0.05722 0.00423 -.02798 -.74212 0.02988 -.99691 -.05629 0.05490 0.19715 0.04814 0.84047 -.01046 -.05631 0.1.00146 -.00000 0.07172 -.90913 0.05153 0.90395 0.00052 -.00201 0.81045 0.01110 0.11276 0.05433 0.85627 0.03160 0.02797 -.91842 0.87511 0.5 in off model centerline) 1 2 3 4 5 6 7 8 9 10 11 12 13 14 1.00172 0.48816 0.05479 0.00306 29 30 31 32 33 34 35 36 37 38 39 40 41 42 SD7003-PT (B .01062 0.48492 0.21359 -.04177 0.01254 0.45985 0.00752 0.00000 0.35423 0.00590 0.98089 0.15572 0.00305 0.05338 0.74815 -.00079 0.00000 0.04885 0.00577 0.01929 -.08076 0.01858 0.97330 0.02188 -.00020 0.06766 0.25853 0.01916 -.76441 0.04037 0.41985 0.99706 0.01884 0.10740 -.54404 0.06985 0.00256 0.05008 0.02838 0.01595 43 44 45 46 47 48 49 50 51 52 53 0.02465 -.00168 0.00000 27 28 29 30 31 32 33 34 35 36 37 38 39 0.99714 1.11940 0.98418 -.41691 -.01674 40 41 42 43 44 45 46 47 48 49 50 51 0.68466 0.03496 -.00607 0.00000 0.09631 0.00445 0.00404 -.85674 0.00000 40 41 42 43 44 45 46 47 48 49 50 51 0.03434 0.01049 0.74091 0.4.00000 0.01214 0.58682 -.78119 0.13869 0.02014 -.05619 0.71354 0.02619 -.66555 0.00002 -.01119 -.02150 0.02078 0.29520 0.79920 0.00913 0.02882 -.60577 -.04847 0.Chapter 7: Airfoil Coordinates SD7003-PT (R .00170 -.00000 0.01871 -.47990 -.00092 -.00280 0.01476 -.00091 0.00822 0.01044 0.00347 0.00153 0.03882 0.98155 0.55800 0.55647 0.97274 0.02687 0.15765 0.22369 0.99067 0.80116 0.00000 -.04979 0.14828 0.00588 0.01039 0.04386 14 15 16 17 18 19 20 21 22 23 24 25 26 0.00397 0.25412 0.00112 0.02178 0.01092 -.5 in off model centerline) 1 2 3 4 5 6 7 8 9 10 11 12 13 1.02641 -.00000 -.94514 0.00000 -.00437 0.00200 0.03009 0.01504 0.51957 0.05009 0.01251 -.40624 0.28571 0.40213 0.05380 0.90760 0.00254 0.99845 0.86491 0.01183 0.62442 0.04022 0.05653 0.68241 0.03514 0.56269 0.02979 0.00001 -.00194 0.00024 0.00190 0.75779 0.20728 0.01260 0.35010 -.02220 -.00000 0.00780 -.02884 0.04687 0.5 in off model centerline) 1 2 3 4 5 6 7 8 9 10 11 12 13 1.96228 0.02989 -.11056 0. -.03279 0.05212 0.01480 0.00805 0.04264 14 15 16 17 18 19 20 21 22 23 24 25 26 0.47847 0.05509 0.02108 0.55247 0.44810 0.98594 0.99094 0.04424 0.00649 0.34937 0.37324 0.05318 0.71029 0.00660 0.00477 27 28 29 30 31 32 33 34 35 36 37 38 39 0.02477 -.05855 0.01520 -.05559 0.00000 0.03862 0.65013 0.01175 -.02946 -.85172 0.03029 -.03021 0.21714 0.91000 0.00358 -.00463 -.00085 -.00141 0.04407 0.48216 0.12282 0.02942 -.99628 0.00610 0.01618 0.02683 0.00358 0.50917 0.98972 0.02483 0.03701 0.98521 0.00315 0.00115 -.02129 0.97723 0.02546 0.73689 0.03628 0.00474 0.02818 0.01966 0.99850 0.27384 0.00162 0.00532 0.05237 0.97479 0.02367 0.03993 0.00044 0.16902 0.98686 0.88320 0.02247 0.02458 40 41 42 43 44 45 46 47 48 49 50 51 52 0.91003 0.00148 -.10516 0.01954 0.36098 0.00736 0.99320 0.97019 -.00728 -.-1.05636 0.00143 1.00000 -.31553 0.09394 0.00000 0.99810 1.78332 0.29665 0.02857 -.00125 0.95149 0.02332 -.00000 0.00000 0.02704 -.01006 -.74005 0.02444 0.86108 0.04689 0.27920 -.24408 0.30461 0.01438 -.22473 0.40331 0.02522 -. 99788 0.00286 0.00918 0.08306 0.00000 0.00443 -.84831 0.00285 0.31682 0.01314 0.20312 0.01597 -.00431 0.08379 0.04601 0.00418 0.26972 0.06240 0.01237 0.01518 0.01078 -.00000 -.02482 -.00594 -.00751 0.66816 -.17663 0.08400 0.07985 0.00273 0.00548 0.56361 -.04499 0.00405 -.01983 -.42236 0.28260 -.05977 44 SD7032D-PT 1 2 3 4 5 6 7 8 9 10 11 12 13 1.85561 -.01930 0.00000 0.03232 0.00000 -.05466 0.32451 0.08227 0.04707 0.01279 0.90268 -.85246 0.70058 0.83807 0.00330 0.79457 0.98248 0.99212 0.01220 -.11631 -.78486 0.00356 49 50 51 52 53 54 55 56 57 58 59 60 61 62 0.00000 14 15 16 17 18 19 20 21 22 23 24 25 26 0.02114 -.02725 -.00000 0.17163 0.04074 0.00040 0.05506 0.04768 0.04193 14 15 16 17 18 19 20 21 22 23 24 25 26 0.07258 0.18733 -.00430 0.45692 0.00094 0.00192 0.01955 -.87592 0.12344 0.92939 0.00097 -.01677 0.00000 o".00150 0.07146 0.03382 0.00001 1.09997 0.05294 0.04579 0.03167 0.04509 0.99580 1.01251 0.98064 -.96493 0.02528 0.00000 0.90897 0.01502 -.62807 0.75229 0.00920 0.00074 0.00000 0.76998 0.07486 SD7037-PT 1 2 3 4 5 6 7 8 9 10 11 12 13 1.03883 0.80782 0.00190 0.01252 0.01394 40 41 42 43 44 45 46 47 48 49 0.20211 0.00107 0.72514 0.38225 -.00635 0.90325 0.00251 0.62499 -.00217 0.02138 0.00028 0.32112 0.04380 0.47487 0.07732 0.27218 0.66498 -.02604 0.31076 0.00324 0.07162 -.12205 0.36989 0.06820 0.80111 0.02964 -.00000 27 28 29 30 31 32 33 34 35 36 37 38 39 0.572 0.00317 0.03728 0.51949 0.15228 0.21266 0.97357 0.00135 0.80383 0.64776 0.06568 0.00223 0.00568 0.01936 0.02917 0.00000 0.00593 0.52799 0.01834 -.06320 0.03790 0.01601 -.94144 0.05629 0.00000 0.93649 0.00344 0.73922 0.01081 0.05435 0.99919 -.5 in off model centerline) 1 2 3 4 5 6 7 8 9 10 11 12 13 1.00090 0.02025 -.04779 0.00150 0.07780 0.01972 -.98886 0.02746 0.00478 0.05490 0.00000 27 28 29 30 31 32 33 34 35 36 37 38 39 0.05735 0.01078 0.29160 0.24117 0.01329 -.37360 0.01518 0.00138 0.00805 -.06655 0.08238 0.07028 .00611 0.07285 0.01065 -.00727 0.77441 0.04581 0.02948 0.02071 -.00125 0.06190 0.00040 0.61360 0.01087 -.00279 0.05527 0.98819 0.98247 0.81714 0.40618 0.01946 -.00303 -.64102 0.01410 0.11048 0.02010 -.00131 0.01283 -.00234 0.00000 -.84249 0.01253 0.07272 0.95100 -.00075 -.06081 0.00990 0.00300 0.79819 -.50700 0.02248 0.00100 0.02077 0.23081 0.01376 0.01830 -.00611 -.00908 0.07915 0.99137 0.99251 0.24471 0.03720 0.03686 -.99502 -.43508 0.07913 0.06784 0.32809 0.40798 0.99877 1.03100 0.00234 0.94911 0.00000 0.08308 0.00951 -.00565 0.01451 0.05186 0.95698 0.77805 0.47315 0.97531 0.98676 0.52036 0.01726 27 28 29 30 31 32 33 34 35 36 37 38 39 0.01110 0.44411 0.00278 0.99274 0.12348 0.83486 0.00628 0.78647 0.00285 0.00029 0.90739 0.99597 0.97785 0.99855 0.56683 0.16234 0.00209 -.02294 0.99922 1.00031 -.35584 0.97687 0.73439 0.02038 0.05568 0.34570 0.00111 -.00149 0.00000 0.00684 -.92233 0.76829 0.02430 -.70106 0.02111 0.00145 0.89925 0.48002 -.00602 0.07298 0.01758 0.00029 0.17208 0.02853 -.00039 -.02672 0.06670 0.01044 0.00444 0.02533 -.00281 -.01935 0.78908 0.81691 0.01015 0.00939 40 41 42 43 44 45 46 47 48 49 50 0.51706 0.38172 0.67833 0.08262 0.04116 0.99929 0.04406 0.00040 0.00942 0.01983 0.07208 0.00484 33 34 35 36 37 38 39 40 41 42 43 45 46 47 48 0.05208 0.01948 0.78859 0.58683 0.00120 -.79593 0.98850 0.03221 0.98586 0.00005 0.03716 0.00438 0.99331 0.00000 SD7032C-PT 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.58.02096 0.16788 0.60399 0.42374 0.00000 0.94657 0.02918 -.00052 0.88800 0.00207 0.21907 0.00392 0.01786 -.00000 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 0.142 Airfoils at Low Speeds SD7003-PT (D .00000 -.80814 0.97369 0.69177 0.85655 0.48859 0.01490 -.56021 -.00113 0.52521 0.00000 0.00770 0.13483 0.57299 0.00000 0.78215 0.00208 0.24444 0.02080 0.00195 0.00364 0.78212 0.05585 0.02320 0.08988 0.03961 0.00580 -.73236 -.04006 0.01484 -.00209 -.04907 0.08236 0.04599 0.02568 0.00000 0.00576 0.04524 0.06905 0.01372 40 41 42 43 44 45 46 47 48 49 50 51 0.06577 0.00727 -.06890 0.12085 0.98922 -.00616 -.00000 14 15 16 17 18 19 20 21 22 23 24 25 26 0.03212 0.26289 0.00963 -.89510 0.94450 0.68369 0.-4.00575 0.00087 -.00400 0.73151 0.00185 0.00830 0. 08283 0.04117 0.03550 0.91610 -.06648 0.10473 0.00196 0.00174 0.01397 0.99376 0.03018 -.56668 0.05446 0.00000 0.00250 -.45008 0.00210 0.01892 -.69676 0.37481 0. 73311 0.10829 0.0231l -.54144 0.20519 0.00000 14 15 16 17 18 19 20 21 22 23 24 25 26 0.63631 0.95556 0.06475 0.l1868 0.39758 0.56410 0.00000 -.Dl333 0.00658 0.46365 0.05181 0.04624 0.19497 0.35012 0.00076 0.98891 -.01191 -.15342 0.41309 0.Ol175 0.73237 0.02421 0.00000 0.99990 0.04835 0.86881 -.61179 0.00872 27 28 29 30 31 32 33 34 35 36 37 38 39 0.04483 0.03178 -.00433 -.00290 0.49992 0.47657 0.37210 0.94941 -.97309 -.10497 0.00241 0.00391 0.01287 -.01019 -.63455 0.00000 0.07458 0.96263 0.66158 0.00000 0.01131 -.04165 0.00159 0.06550 0.53611 0.26596 0.92851 0.00861 37 38 39 40 41 42 43 44 45 0.88226 0.05314 14 15 16 17 18 19 20 21 22 23 24 25 26 0.75148 0.47835 O.04850 0.93019 0.21095 0.01553 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.09412 0.0701l O.Dl055 0.01723 -.00116 0.99012 0.00000 0.02278 -.00754 -.14450 0.02582 -.00634 -.03955 0.06332 0.07149 0.90289 0.00542 0.98594 0.01318 -.02637 -.96463 0.01001 0.72084 0.89294 0.00261 0.00290 0.0251l -.01270 -.08l10 13 14 15 16 17 18 19 20 21 22 23 24 0.00009 0.99858 0.01943 -.00717 0.07518 0.02177 0.00394 0.00254 0.01381 0.02221 -.98975 0.15401 0.98036 0.02680 -.03208 0.00000 0.83860 0.04122 0.19377 0.02707 0.07084 0.00080 1.63630 0.04318 0.00165 0.01503 -.03962 0.00378 0.34847 0.01267 0.00000 -.96858 0.95343 0.35134 0.00610 0.04995 0.00350 0.02575 0.03946 0.35724 0.09052 0.00000 SD7062-PT 1 2 3 4 5 6 7 8 9 10 11 12 13 1.75556 0.01167 0.01594 0.00042 0.32848 0.01088 0.06435 0.01185 40 41 42 43 44 45 46 47 48 49 50 0.06581 0.98352 0.03306 -.00000 0.09977 SD7080-PT 1 2 3 4 5 6 7 8 9 10 ll 12 13 14 15 16 1.Chapter 7: Airfoil Coordinates SD7043-PT 1 2 3 4 5 6 7 8 9 10 11 12 1.60846 -.02566 -.00464 0.05878 0.00000 -.27587 0.91890 0.71025 0.01651 0.27330 0.06236 0.13205 0.00039 -.00110 0.00000 143 .19526 0.06139 0.62824 -.02211 -.03430 0.01561 0.00291 0.02164 0.01760 0.01715 0.03589 0.00000 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 0.02601 0.11109 0.00283 0.84994 0.04505 0.93589 0.07916 0.53688 0.12296 0.43739 0.00832 0.37827 0.53834 -.77743 -.05332 0.01089 0.20697 0.03645 0.00181 0.00935 -.02616 0.68145 -.00234 25 26 27 28 29 30 31 32 33 34 35 36 0.00072 0.58628 -.68963 0.00117 0.43149 0.06886 0.77406 0.01725 -.26367 0.00070 0.00128 0.01767 -.05108 SD7080-PT (R .87480 0.03217 0.67807 0.99183 0.05550 0.02710 0.06910 0.01529 -.15520 0.069.99093 0.02526 0.01123 0.45218 0.95485 0.08101 0.00096 0.02892 -.02595 0.Dl003 -.98508 0.00845 0.05120 0.05468 0.00649 -.02546 -.00401 -.03170 0.06782 0.01998 49 50 51 52 53 54 55 56 57 58 59 60 61 0.11281 0.00454 -.00082 0.00487 0.28829 0.99874 0.81907 0.01974 0.98251 0.49176 0.05006 0.46235 0.65950 0.02945 0.81359 0.00442 0.04689 0.96737 0.05018 0.00900 -.00000 0.09144 0.00000 0.41947 0.00014 0.66589 -.00555 O.00526 0.82892 0.82797 0.04793 0.00021 -.01684 0.01040 0.89835 0.24 0.02076 -.00068 0.06140 0.02971 -.96163 0.64469 0.60750 0.71136 0.01982 -.72826 0.81902 0.02299 -.77121 0.01157 0.00098 0.16464 0.06297 0.00456 0.02559 0.00056 -.01555 -.99375 0.05381 0.00000 -.97740 0.01028 0.01436 0.01634 0.00653 0.00000 0.00299 0.00072 -.10941 0.88433 0.06657 0.99747 1.02182 0.60375 0.98462 0.82590 -.02622 -.05456 0.77224 0.00633 -.00181 0.10893 0.06707 0.02600 -.99654 1.01147 0.01564 -.06858 0.08914 0.00930 0.33088 0.03414 0.57352 0.0151l 0.07047 0.01578 -.92416 0.55543 0.repeated) 1 2 3 4 5 6 7 8 9 10 11 12 13 1.07803 0.72884 -.01475 -.00509 27 28 29 30 31 32 33 34 35 36 37 38 39 0.00000 -.05644 0.00438 0.03165 0.05138 0.11012 0.05836 0.04668 0.00000 0.65702 0.06475 0.19765 0.27461 0.02833 40 41 42 43 44 45 46 47 48 49 50 0.91714 0.00840 0.99193 0.18926 0.00000 0.00655 -.00453 0.99711 1.02762 -.03252 -.00120 0.28393 0.04265 0.29930 0.00000 0.02428 -.23366 0.03330 0.00000 0.00285 0.00000 0.03308 0.01025 0.23356 0.08161 0.01602 0.68761 0.00240 0.00634 0.00562 -. 06638 0.00159 0.06555 0.02060 0.03243 -.40677 0.98882 -.33532 0.02834 -.96936 0.00016 -.04507 -.02121 0.99396 0.00000 0.04194 .99882 -.00000 -.95168 -.00922 0.00968 0.82765 0.97538 -.19340 0.95497 0.05010 0.02689 0.93466 0.00087 -.74885 0.99653 0.99114 0.88482 0.57180 0.01766 0.03733 -.00194 0.00000 SD7090-PT 1 2 3 4 5 6 7 8 9 10 11 12 13 1.00576 0.02834 -.00311 27 28 29 30 31 32 33 34 35 36 37 38 39 0.02585 0.02598 -.15766 0.64219 -.99216 0.62509 0.03124 -.04938 0.05173 -.73551 0.07929 0.00736 0.04800 0.00191 0.89561 0.03007 -.11169 0.59401 -.73667 -.93672 0.00985 -.86826 0.04662 0.94846 -.06377 0.00625 -.00195 -.80901 -.02041 -.80447 0.99462 -.03824 0.35147 0.98897 0.00096 0.04972 0.00111 0.00000 0.02593 -.88704 -.01697 -.02675 -.02842 0.00000 0.70619 -.04886 34 35 36 37 38 39 40 41 42 43 44 0.04128 0.98570 0.00830 0.00573 0.06629 0.05096 0.144 Airfoils at Low Speeds SD7084-PT 1 2 3 4 5 6 7 8 9 10 11 12 13 1.04068 0.06612 0.23836 0.00000 25 26 27 28 29 30 31 32 33 34 35 36 0.43092 0.97209 -.00000 23 24 25 26 27 28 29 30 31 32 33 0.26455 0.00201 0.00243 0.02121 0.00072 0.90390 0.04045 -.00013 1.01780 40 41 42 43 44 45 46 47 48 49 50 51 52 0.98366 0.01549 0.00586 0.38273 0.03114 0.96612 0.00410 0.02023 -.13372 0.03382 0.00000 -.12168 0.40596 0.57198 0.02453 0.07125 0.32521 0.00000 0.04777 0.02985 0.01027 0.07007 0.01606 0.02436 37 38 39 40 41 42 43 44 45 46 47 48 0.03329 -.01387 0.62448 0.53740 0.92732 0.04584 0.00000 0.29873 0.86846 -.29657 0.99950 -.78624 0.00899 0.22966 0.37682 0.00000 0.01270 -.05100 -.05399 0.03709 0.02466 -.02091 -.39780 0.00000 14 15 16 17 18 19 20 21 22 23 24 25 26 0.70568 0.00554 0.00000 1.01455 0.02921 0.01307 0.18958 0.00571 -.05930 SD8000-PT 1 2 3 4 5 6 7 8 9 10 11 12 1.06279 0.00248 0.03140 -.00319 -.33493 0.02736 -.07344 0.02978 0.00222 0.01870 -.02773 -.00632 0.01190 -.00083 0.01058 0.25319 0.00046 0.00073 0.02080 0.00412 0.00666 -.01624 0.06130 0.07305 0.05519 0.00113 0.00795 0.06181 0.00000 1.99027 0.00048 0.01528 -.66694 0.04660 0.82063 0.02852 0.00275 0.04409 0.04153 0.02622 -.58645 -.00504 0.01364 0.00147 0.27801 0.02135 0.00753 0.00807 -.03878 0.06811 14 15 16 17 18 19 20 21 22 23 24 25 26 0.00026 0.00000 12 13 14 15 16 17 18 19 20 21 22 0.91556 -.00308 27 28 29 30 31 32 33 34 35 36 37 38 39 0.59887 0.062'52 0.00612 0.00065 0.17594 0.01161 0.97294 0.49319 0.03696 0.02383 0.82595 0.23338 0.00974 0.00226 0.00000 0.02359 -.86804 -.12139 0.13368 0.07055 0.05786 0.00000 -.00000 0.02400 -.00110 0.05965 0.00054 0.05690 0.90654 -.99401 -.00961 -.00873 -.00647 0.02616 0.10990 0.64809 0.04035 0.01823 -.75090 0.02247 -.33492 0.08309 0.06117 0.D2008 -.00889 0.36217 -.52013 0.00497 0.68131 0.00793 0.00248 0.98171 0.82036 0.44099 0.53534 0.01796 0.03230 0.93611 0.05518 0.00195 0.74736 0.00101 0.01708 0.87495 0.00000 0.18226 0.00000 0.02295 -.00622 -.25884 0.22209 0.04970 -.01961 0.93423 0.00130 0.00000 0.13759 0.99665 0.00800 0.19244 0.00275 0.03199 0.04383 0.25288 0.99949 1.16687 0.00255 0.04887 0.00050 1.07629 0.00430 0.03476 0.04514 -.17660 0.01668 0.97424 0.70380 0.01491 0.04503 0.07052 0.00371 0.47817 0.79664 -.03760 0.06299 0.01276 0.96484 0.00115 -.04464 0.89450 0.99724 0.03413 0.00000 0.01575 -.00000 13 14 15 16 17 18 19 20 21 22 23 24 0.02689 -.99809 0.32881 0.11413 0.02480 40 41 42 43 44 45 46 47 48 49 50 51 0.67337 0.01874 0.01900 0.02910 -.02032 -.00634 0.00307 0.01413 -.29527 0.03319 -.00379 0.43854 0.75542 -.00318 0.40845 -.00099 -.01066 0.00000 -.53659 0.00275 -.48977 0.01716 0.01893 0.05264 SD8020-PT 1 2 3 4 5 6 7 8 9 10 11 1.00485 0.75841 0.01106 -.46494 0.66395 -.03672 0.01101 -.52937 -.99014 -.05866 0.00137 0.02676 -.62373 0.00275 0.45167 0.01068 0.03335 -.51865 0.99794 1.01358 -.00182 0.00102 0.68385 0.99198 0.00048 0.00568 -. 00751 -.00097 0.05288 0.00000 0.33433 0.05629 0.64966 0.51953 0.08754 0.09621 0.10427 0.00336 -.02103 49 50 51 52 53 54 55 56 57 58 59 60 61 0.01663 0.01128 34 35 36 37 38 39 40 41 0.48508 0.76755 0.00448 0.01472 -.01631 43 44 45 46 47 48 49 50 51 52 53 54 0.37070 0.03955 0.62232 -.46196 0.06086 0.02979 -.14094 0.00276 0.71153 0.09232 0.83783 0.00000 0.06398 0.00060 0.87438 0.51941 -.02317 0.05044 0.11345 0.01732 -.02308 -.00793 0.00218 0.00413 0.99835 -.98481 0.82209 0.08158 0.89998 0.08771 0.01087 -.01562 0.10466 0.04891 0.09962 0.54497 0.24661 0.02790 0.07278 0.00000 145 .50396 0.00440 23 24 25 26 27 28 29 30 31 32 33 0.00936 0.06120 0.04827 0.62886 0.17375 0.06867 0.00276 0.98316 0.77971 0.06686 0.00000 0.02815 0.00158 0.98203 -.97784 -.08673 0.02869 -.06831 0.23707 0.68977 0.00719 29 30 31 32 33 34 35 36 37 38 39 40 41 42 0.84101 -.00785 0.02154 -.00477 0.07863 0.73298 -.16299 0.03807 -.97029 0.04183 -.01426 -.01151 -.00000 SPICA-PT 1 2 3 4 5 6 7 8 9 10 11 1.86854 0.79536 0.00037 1.30793 0.06855 0.92254 0.00000 -.01297 0.42396 WB135/35-PT 1 2 3 4 5 6 7 8 9 10 11 12 13 14 1.05337 0.Dl358 0.93869 0.00210 0.00000 0.02660 0.98033 -.01244 0.04421 -.06943 0.93238 -.07332 0.05959 0.G3858 -.98865 0.99700 1.01217 -.02226 -.70863 0.00433 -.00056 0.10819 0.99965 0.02907 -.00000 0.06621 0.98927 0.90696 -.00555 0.03279 -.00480 -.00287 -.83607 0.03098 0.03241 -.01653 -.00000 -.04300 -.02078 0.00000 0.00274 0.03509 -.70186 -.98701 0.00305 0.00040 0.03152 0.00000 0.00603 0.59982 0.00566 0.02351 0.09633 0.04372 0.96319 0.00542 0.01496 0.34640 0.99821 0.93184 0.07550 0.30002 0.00254 1.25831 0.03907 -.00569 -.00191 1.02619 0.09367 WB140/35/FB-PT 1 2 3 4 5 6 7 8 9 10 11 12 1.00000 0.02825 -.58684 0.43114 0.24624 0.66984 0.00950 0.37689 0.05425 0.01313 0.75229 0.99747 0.07254 0.00000 0.00293 0.99521 0.01016 0.00972 0.05818 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 0.99542 -.00000 -.01706 -.08172 0.12166 0.95585 0.02704 -.96097 -.73344 -.07618 0.06102 0.03411 -.01781 0.00482 0.07243 0.03555 0.Chapter 7: Airfoil Coordinates SD8040-PT l 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 1.19031 0.00000 0.00000 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 0.99303 0.02054 0.22090 0.01310 -.99738 -.00766 -.85855 0.01887 0.00000 0.06167 0.92894 0.01932 -.01512 -.01721 0.01238 0.99723 0.06491 0.99072 0.98989 -.01374 -.02948 -.08820 0.70000 0.00536 0.01219 0.00123 0.01007 0.00708 0.02146 0.16063 0.59546 0.02546 37 38 39 40 41 42 43 44 45 46 47 0.02962 0.01433 0.17395 0.00231 0.01392 0.89180 0.55472 -.69153 0.02117 -.00331 0.01724 0.89714 -.00671 0.00169 0.09962 0.02237 0.00892 0.01259 25 26 27 28 29 30 31 32 33 34 35 36 0.01812 0.12049 0.94426 -.05706 0.91961 0.01531 0.00000 0.00000 0.37012 0.00259 0.88632 -.03066 -.00864 0.09598 0.00000 13 14 15 16 17 18 19 20 21 22 23 24 0.00630 0.61451 -.01217 0.00046 0.44053 0.02456 -.08378 -.03167 -.67693 -.41105 0.25066 0.04098 -.95575 0.03109 -.02751 -.98199 0.30158 0.06119 0.36440 0.00889 0.00578 0.97036 0.00717 0.07876 0.01048 0.00224 0.43789 0.62781 0.98804 -.03498 0.17643 0.01025 -.00406 0.04431 0.09773 0.00515 0.00520 0.87616 0.38505 0.00375 0.04403 0.03758 0.00457 0.00000 15 16 17 18 19 20 21 22 23 24 25 26 27 28 0.00063 0.78998 -.00308 0.01452 -.53052 0.00920 -.00139 0.01155 0.01908 -.02874 -.02579 0.00000 0.10149 0.27127 0.27413 0.00192 0.05152 0.63629 -.60202 0.01638 -.04901 0.99188 0.28913 0.81262 0.03293 0.98772 -.95165 -.10112 0.06669 0.01662 -.00383 0.02743 0.33180 0.76684 0.20477 0.42357 0.32927 0.04050 0.00686 0.00365 0.00265 -.06122 0.09418 0.81667 -.03837 0.09740 12 13 14 15 16 17 18 19 20 21 22 0.03340 0.00000 0.77828 -.10802 0.03477 O.99763 -.00137 0.16424 0.00545 -.00445 0.00608 0.11668 0.00780 0.06785 0.56368 0.04995 0.21630 0.47150 0.50644 0.01375 -.01423 -.00000 0.85532 -.03637 -.03588 0.02376 0.25293 0.41526 0.99040 -.50241 0.33472 0.03874 0.99297 -.00207 0.75053 0.13855 0.55852 0.02135 -.09128 0. 146 Airfoils at Low Speeds . 0 S3021 S4061 S4233 S8040 16 16 0010 -0.082 17 1001 - n i7 IUO ::>UIU\JU -0 04R4 -0 SPICA 04ll~ -0 1024 Note: Average moment coefficient as predicted by the ISES code for Rn = 200.0758 -0 O~Rn -0.0578 RG15 120 -( 1./4 ~.0 !2 17 -( 1. 138 174 -0.0597 .OR21 H02/9 i15 -0.055 E387 17 FX60-100 -0.( 1.1201 -O. 147 .0873 -0 lOR DAE51 '1 )1 l2 '1 12 -0.y -0.000.052 -0.Chapter 8: Predicted Moment Data Sorted b. ~54 10 -0.0 18 -11.0 lO -0. N arne Airfoil Cm. 148 Airfoils at Low Speeds . 00 2.64 3.00 SD8020 10.51 1.30 DF103 11.94 S2055 7.70 13.5411 11.65 SD8040 9.85 SD5060 9.86 1.32 S3014 9.72 4.46 3.76 S2048 8.00 2.10 4.48 3.87 SD7090 9.24 E387 9.55 FX63-137 13.94 HQ2/9 8.16 NACA 0009 9.03 E374 10.36 S4233 13.10 2.72 3.63 2.37 1.00 0.01 E214 11.85 4.43 MILEY 12.96 84061 9.99 SD8000 8.59 5.47 2.53 3.30 DF102 11.99 11.48 SD7080 9.14 54180 9.82 10.95 SD7037 9.74 SPICA WB135/35 13.51 SD7043 9.02 3.98 3.53 4.15 E205 10.55 DAE51 9.20 3.99 J5012 12.66 SD7032 9.00 RG15 8.71 0.Chapter 9: Airfoil Thickness and Camber Sorted bY N arne Airfoil %Thickness %Camber AQUILA 9.37 3.91 2.63 1.00 MB253515 14.84 3.81 5.92 1.30 E193 3.50 NACA 64A010 10.56 2.75 WB140/35/FB 3.00 2.77 4.13 SD7062 13.46 2.05 CLARK-Y 11.06 3.31 1.92 149 .52 2.00 2.97 1.57 10.45 2.30 SD6060 10.22 E193MOD 11.57 83016 9.00 NACA 6409 6.00 9.00 0.09 S3021 9.00 NACA 2.90 S4062 9.97 3.91 10.80 FX60-100 9.15 SD7084 9.26 SD2030 8.96 2.66 S2091 3.97 2.00 0.99 1.38 4.18 SD7003 8.60 3.96 2.25 SD2083 8.74 SD6080 9.10 S3010 2.98 DF101 11. 00 0.97 1.45 9.63 2.55 FX60-100 1.00 NACA 6409 9.00 9.90 S4061 9.30 DF102 11.71 1.53 WB135/35 5.80 E387 3.01 10.72 CLARK-Y 4.96 MB253515 .99 2.06 3.00 2.24 E374 2.20 3.57 10.99 SD8040 9.97 13.48 807080 9.30 DF101 2.37 1.95 9.55 11.57 S3014 9.75 13.56 8.84 SD6060 3.46 2.14 S4062 3.60 9.96 2.00 S08020 82091 10.63 1.52 2.16 12.15 3.00 10.13 2.66 807032 9.10 3.99 1.91 2.05 AQUILA 2.30 OF103 11.09 9.53 4.38 4.15 E193MOD 0.72 SPICA 11.74 806080 9.91 3.02 807037 OAE51 9.43 14.92 WB140/35/FB 3.82 S3010 10.10 0.46 2.66 82055 807003 8.30 SD5060 9.65 0.70 13.77 4.03 E214 3.22 E193 10.36 S4180 3.47 2.51 807043 9.48 E205 10.81 MILEY 3.85 4.37 3.00 J5012 5.26 13.541! 11.64 84233 3.86 1.76 RG15 8.99 HQ2/9 NACA 0009 9.94 82048 808000 8.97 3.150 Airfoils at Low Speeds Sorted b Thickness %Thickness %Camber Airfoil 7.00 11.98 SD7062 2.59 FX63-137 3.00 2.25 802030 8.96 S3021 S3016 9.94 13.18 9.10 4.98 9.85 802083 8.00 12.31 SD7084 9.00 6.92 8.50 NACA 2.00 11.51 1.74 11.32 2.87 807090 9.00 NACA 64A010 10. 20 3.09 83016 2.60 84061 3.30 805060 9.71 1.26 S4233 13.46 SD7003 7.48 3.02 3.92 10.00 2.10 0.46 9.37 OAE51 11.85 802083 8.90 9.57 E!93 3.96 MB253515 807080 9.36 9.55 10.00 2.00 12.05 9.98 507062 3.74 9.00 0.74 11.91 E374 8.99 HQ2/9 8.51 1.51 9.94 13.72 3.52 2.55 FX60-100 CLARK-Y 11.37 1.00 808020 10.30 OF102 2.01 E205 807037 9.00 NACA 64A010 10.Chapter 9: Airfoil Thickness and Camber Sorted by Camber %Thickness %Camber Airfoil NACA 0009 9.30 11.65 808040 2.98 9.75 WB135/35 9.96 S3021 10.99 2.81 MILEY 5.30 OF103 11.25 802030 2.5411 11.50 NACA 2.00 0.00 J5012 8.59 FX63-137 6.77 S4180 4.99 1.63 1.66 S2055 8.31 2.43 14.82 S3010 10.45 OF101 11.96 9.80 E387 3.97 13.18 SD6080 13.24 10.06 3.72 SPICA 5.00 NACA 6409 151 .95 S07032 13.53 54062 4.63 2.00 2.85 E193MOO 4.92 3.84 806060 1.00 807084 9.99 82048 8.86 808000 1.10 82091 3.87 807090 9.76 RG15 8.15 11.70 WB140/35/FB 3.47 2.14 9.91 10.00 0.97 3.38 AQUILA 4.48 2.97 9.00 9.16 12.13 507043 9.64 3.53 3.94 1.10 4.15 2.22 3.56 2.32 2.03 E214 4.57 S3014 9.66 9. 152 Airfoils at Low Speeds . 0062 ln. Avg.0158 ln. . 10..Chapter 10: Digitizer Plots ~1g. J / 30 ln. • 0.. ~--------- ~· '"·' ocuo-• ' OJ«-•-•• Avg. dlff.. • 0.. dlff.1 AQUILA & AQUILA-PT 153 --. ./ . 030 in. diff.---- --.0172 in .4 - DFIOI G DFIOI-PT Avg.3 DAE51 G DAE51-PT Avg. 030 in.0150 in .154 Airfoils at Low Speeds ~g. 10. --c Fig. . diff. . --. 10. a 0. • 0. • 0.. 10.. dill.0354 In I I . • 0.030 ln. ------- .0189 ln .5 155 E193 &E193-PT Avg...Chapter 10: Digitizer Plots ~ Fig. dill. ~. '''"" ''"""'" Avg.. ---- ----- ---~ ------ / .. .. 10. • 0.B - E205 & E205B-PT Avg.156 Airfoils at Low Speeds ~ Fig. • 0.. d!ff.. ..0151 ln. 10.--- ---- ~ig.0106 in .7 E205 & E205A-PT Avg. ------- . d!ff. 030 ln. __. . • 0.0081 ln.--- 157 . .9 E214 S E214A-PT Avg. 030 ln. diff.Chapter 10: Digitizer Plots ~ 10. =: .0066 ln . diff. • 0.10 E214 & E214C-PT Avg. ~· 10. - E374 & E374B-PT Avg. 1=1 . • 0. dill.12 . ~f!g.158 Airfoils at Low Speeds ~!g. 10.11 E374 G E374A-PT Avg. dill.0063 in. • 0. 030 in. 10. ---- l 1_. --- ..0089 in . 10.14 E3B7 & E3B7B-PT - Avg. • 0. d1ff. c: Fig. 10.13 159 - E3B7 & E3B7A-PT Avg.Chapter 10: Digitizer Plots ~F1g. • 0.0109 1n.0174 in. d1ff. I \ / I . .16 FX63-137 & FX63-137-PT Avg. d!ff. 030 ln.160 Airfoils at Low Speeds ~ Fig. • 0. ~-- --. 10..15 - FXB0-100 S FXB0-100-PT Avg.0074 ln. 10. • 0.. - Fig. ~------­ -- .------. dill.0322 ln • . ----__. 030 in.. • 0.0164 in • .Chapter 10: Digitizer Plots ~ Fig.. I \..------ \_ ___ ___ ~ Fig.- .... 030 1rni7.0124 in .----...- - -----. . . 10.17 HG2/9 G HG2/9A-PT Avg. • 0. 10.18 HG2/9 G HG2/9B-PT Avg. dill. 161 ...... dill..... • 0.0128 1n..20 MB253515 & MB253515-PT Avg.~~:g~~J5:0~1:2~&~J~5:0~1:2-:P:T~~~~::~~~:=:=====:::=:=>a-• Avg...- .. d1ff..l. ..<k. • 0. 10..0047 1n .. 0 030/ _. ...~~O~.030 1n.162 Airfoils at Low Speeds ~~::~~F~lg~.__ ____ \ Fig. d1ff. .0221 ln.. ----- - Avg... • O. d1ff.Chapter 10: Digitizer Plots .OOBO ln .. • 0. --- - 163 . • 030 ln.. Avg.... dill.. d!ff.23 NACA 2. • 0. d!ff. 030 in. . 10. 030 in.164 Airfoils at Low Speeds ~!g.0046 in . Avg.5411 & NACA 2.0065 !n . • 0. . .5411-PT Avg. ---- 165 ... -+ ~ Fig... • 0... Avg. 030 ln..0096 ln .0097 1n.Chapter 10: Digitizer Plots ~···· """ . --.. .26 """ - RG15 & RG!5-PT Avg. d!ff.. '"'' . • 0. d1ff. 10. 166 Airfoils at Low Speeds ~ fig. 10. dill. diff.27 52048 ~ 52048-PT Avg. • 0. ---- .0233 in. ----~ Fig. 10.0246 in. • 0.28 52055 ~ 52055-PT ------- - ·Avg. dlff. ---.----..Chapter 10: Digitizer Plots c Fig. dlff. -------- 167 .29 52091 & 520918-PT =---· ---·- Avg.0329 ln .-------. • 0. 10. • 0.0128 ln. 10.30 53010 &53010-PT Avg. .---' ----- ~ Fig... 0114 ln. . . dlff.0. 10.168 Airfoils at Low Speeds c Fig. 10.32 53016 G 53016-PT Avg. 030 ln.31 - 53014 & 53014-PT Avg. --- ---- ___ ___. .- ..0. ---~c Fig. dlff.0092 ln . 10. • 0. • 0. 030 ln.Chapter 10: Digitizer Plots ~ Fig. . 169 . ~Fig. dlff. 030 ln. diff. .34 53021 & 530218-PT Avg. 10.0084 in .33 53021 & 53021A-PT Avg.0067 in . • 0.. d!tf..36 5406! &S406!B-PT Avg. dlff. !0...35 54061 G S4061A-PT Avg. 10. -..014B !n .170 Airfoils at Low Speeds c:Fig. 030 ln.__---~ ~lg.. ---~ .. • 0. ...0176 ln. ....0141 ln.. .. dlff... 030 ln. 171 .. dlff. Avg. • 0..37 54062 & 54062-PT ·----Avg. ..0043 ln .. c "·" "'" '._. • 0. 10.Chapter 10: Digitizer Plots ~g. 10.172 Airfoils at Low Speeds ~ Fig. 030 in. . • 0. i0.0067 in . . diff. c Fig. • 0. diff.40 5020B3 G S020B3-PT ---Avg.0072 in . • 030 in.39 502030 G 502030-PT -Avg. . 030 in.---- c Fig. .. .------- 173 . 030 in . 10. ---- -..Chapter 10: Digitizer Plots c Fig.41 505060 G 505060-PT Avg.0095 in . • 0. dlff. diff. • 0.0110 in .42 506060 G 506060-PT Avg. 10. ... . diff. 10.43 SD60BO G SD60BO-PT • Avg. • 0.174 Airfoils at Low Speeds ~f!g. 030 in.0052 in • . . 44 507003 &507003-PT Avg. d!ff. . . • 0.Chapter 10: Digitizer Plots F!g. 10. d!ff.0041 ln.45 507003-PT G 507003-PT (R) Avg. 10. 175 .0007 !n . 030 !n. • 0._ J ~ F!g. .47 507003-PT & 507003-PT ~) Avg. • 0.0012 in . diff. 10. 030 in.176 Airfoils at Low Speeds c Fig.46 507003-PT & 507003-PT (A) Avg. . 10. . d1ff. c Fig. • 0. 030 in.0022 in . 10. d!ff.0034 !n. . c Fig.Chapter 10: Digitizer Plots c Fig. 10. 030 !n.49 507003-PT &507003-PT (D) Avg. E 177 .0033 ln .48 507003-PT G 507003-PT (C) Avg. • 0. • 0. d!ff. . • 0... D 0. diff.. .. "------ ~ig.. 10._..178 Airfoils at Low Speeds c : F i g .. Avg.51 507032 &5070320-PT \r--- - ..50 507032 & 507032C-PT Avg. 030 in.__ ..0175 1n .. diff. 10.. F .---- --.0062 in. . • 0.. 030 ln. dlff..0099 ln . --- ---- 179 .Chapter 10: Digitizer Plots ~ Fig.. • 0.53 S07043 &507043-PT Avg..__ C F ! g .. .52 S07037 &S07037-PT -Avg. dlff.. . 10. !0...0120 ln . 030 ln. -----. . • 0..:.____..0093 ln..- . dlff._ - -------. 10.180 Airfoils at Low Speeds Fig. ln.54 507062 Q 507062-PT Avg. -.. dlff.56 507080-PT S 507080-PT {A) Avg. 10.Chapter 10: Digitizer Plots c Fig. = 0. ---~--- c fig.0025 in .0185 ln. • 0. 030 ln.55 507080 S 507080-PT ---Avg. 10. 181 . . dlff. • 0.57 - 507084 & 507084-PT Avg. 10. ------ . dill. --~ c Fig. • 0.0071 in. 10.0097 in. dill.58 507090 & 507090 PT Avg.182 Airfoils at Low Speeds L Fig. d!ff.59 183 - 508000 G 508000-PT Avg.0050 ln . • 0. . dlfl. 030 ln.0037 ln. 10. r __.Chapter 10: Digitizer Plots c Fig. • 0. b030 ln. 3- L . Avg. 184 Airfoils at Low Speeds c Fig. ..61 SDB040 & SDB040-PT Avg.._.. dlff. 10..0139 ln.0143 ln. / ~·· '"·'' '"''' ' . ~ 0. dlff. • 0.. -Avg. 63 HB135/35 G WB135/35-PT Avg. diff. / ----..0097 in ...0144 in. 10..64 WB140/35/FB G WB140/35/FB-PT Avg... 030 in.-. . • 0. 10. • 0. diff.Chapter 10: Digitizer Plots Fig. Fig. ~ ---- - 185 .. 186 Airfoils at Low Speeds . 0065 ln. . • 0.0131 in .Chapter 11: Airfoil Comparison Plots ~ Fig._:F~iQ~-~1~1-~2~~D:F~10:1~-:P~T~G~D~F1~0:3:-P:T~----------------------===:::::::=>> __ Avg. - ~--·_. diff. 11.1 DF101-PT G DF102-PT ------:=-Avg. 030 in. • 0. dlff. 187 . 030 in.0139 in .4 HQ2/9 G 52048 Avg. dill. 11. 11. • 0. . dill. .188 Airfoils at Low Speeds c: Fig. • 0.0134 in.3 HQ2/9 G RG15 =--I Avg. c: Fig. 11. ~ Fig. 030 ln. • 0. 11.0063 ln.6 HG2/98-PT G AGI5-PT Avg. • 0. .0118 ln . ---.5 HG2/98-PT G HG2/9A-PT Avg.- 189 . dlff. d!ff.Chapter 11: Airfoil Comparison Plots ~ Fig. \ ' ... 11. -~ _________ / ~ (_ F1g.030 1n.. 0.---- / / .8 S3010 G DF101 -'-======--· -·=====-·• 0. Avg... 11.. a r~ .0240 1n. d1ff.7 HG2/9B-PT G S2048-PT .--.____--- I ------.190 Airfoils at Low Speeds c F1g...0589 1n. .Chapter 11: Airfoil Comparison Plots Fig. \ \ Fig. • 0. 11..0210 ln. 0386 ln. 11... ...__ ______ / 191 .9 53021 & 507090 Avg....10 - 508000 &HQ2/9 Avg. . dlff. • 0. diff. ---- .. 030 in... .0264 in . • 0. I I 1 11.--- ---.030 I n.... '-. .... I .. \ \ \ ~1g.=:. -----. 97f.~ Fig. diff.11 W8135/35 ·~ &M8253515 I -----_:Av~g~.12 I e-- E205 ll E193-PT Avg.192 Airfoils at Low Speeds ~====~====~~-~·~..:. 11.____. . . ..... ....... .. .... ...~ : : : : : : : : : : : : ::~.... '' .... L...... . ... .... .......... . ............ .... . ··················· ... ...... . ............. ..:l \. '"]'" . 1 : : . lJ Chapter 12: Polars and Lift Plots ~~~~~~~~~~~~~-~~-~-~---~~~~~-~ :::::::::::::::::::::1:::1:::1:::::::::::::::1::::::1:::1:::1:::1:::::::::::::::1::: ' . . ..·......... -0 ... ...f . 0'0 f' 0 N 0 0 Ill 0 I 0 - 0 :- ~·o- .... .... 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' ' ' ~ ' ' ' ' ' ' ' ' ' 0'0 .. ··-···· ... ·················· ... . ....... ....: : : . . ... . . .... .. .. ...~~. ..... .: .......: ~::P.. ... . ......... ~ ........' .. . ...... . . . .!:.' ..... ...~o(J-... O'l II II II II c c c c a:: a:: a:: a:: 0 0 0 0 0 0 0 0 q_oO~O..: : ......·. : ..···········j"-'1~:· ' ' ' o(J' . .... ..' ' ' '' ' ' ' '' ' : : : : '' ' ' ~·o- .. . .. . ~ . : ...: ... ... .. ... : . .. . .... llJ.. : : . : .... . . . ....' ....... .. ........1~~()- .. .. : : 0 "' :::::::!::::::::::: ::t::::::::::: :::\::::::::::: : : ' ..::::::r:::::c ::::::::::::::: : : ~ tMJ I ... :.~~.~ . ·1· . :--t:·... : ...j JJillii • : J~l ltf. ·--~--T·-~---~ -10..:::::: :::~ ::: ···: ···:···: ···:· . . -~--.. -: . .4 CLARK-Y-PT (Watson/Boren) Rn = 60. : : : : Jf..:-..:-.... . :J::yr:c J::fi}. -:T'·[··· .. ... ·: '.~~~~~ -... ..:r: : : <Y.... +·+·· ::-. .. 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' tihl::\:::\::1: : : :::!::: :::!::: :-r~>+-[1> ....... . .:..... ........... .. ....... :.. .. 12... ....... ...... .... ...._. ' : . ... . . .... . . " ' : " ' ! " ' ! ' " ! ' " ' " ! ' " ! ' " ! 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' ...0 MB253515-PT (Bam e) = Rn 70.' . ..0 :···t:·:··· ....... .......... .. ...... : 0 : .. ..: .......000 ~ ~...0 : : . ... : .. ··> ..... -~ .. : ..... : : : ::T:T:r::!::: ::r:.. 0 ::: :::::::c:c...... ·+·+·+·+·· . .. .i... :a . ···:···:···:···:··· ~~~F~t~:r~+~: MS253515-PT (Same) Rn = 90. ::··· --·. ..... 10. ···:··· 0 __.:::l:::~::· ::rrrr: "1 ···:······:···:··· ···:···:···:···:··· ···:···:···:···:··· 0 I : ·= : .-.' .0 249 . +-. .0 : : : : 0... ···j . '' ' . ·:: .._.. ........... ~ ... :.... :--·:"'j"'''.... I--·f··-~·~--........ .. 0 0 0 -i' .. ~ .:~::T:: :::~::~~:::~:::~::: :::~:~~ ~:· :::~ ... : .. . : .0 0.. .O.........~-~. ...-1 0 ......~..... ...... . : : : • : : : : : : ~~~j-~~~~~~~~ "1 ~:: :~(.... :.. 12.. ·.... :...::::::::::::: :::::::::::::::....0 : 20.f ... ... . : .0 .. .. ···1··+-+-+-.. .. .... : ....-. 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( ::: ::: ::: : (: :::: ::...' ........ -~···~··"!"-· ···~···~···~···~··· ···~---~···~··T·· : : .. .: ......:::j::: ::::::::::::::::::: ::::::T::::::::: ::r:::::rr: • • • f• ·• • Ti 1-JJJJ iT if : .... ..................... ...-::::::r::.... Airfoils at Low Speeds "2' Q) (/) 0 Lt '-' 0 0 0 0 II c ~QN(Y) II c 1!1 0:: 0:: 0:: 0:: 0 . ..... j. .· .... .... . . .::: ::::::::::::::::::: :::....·o- ..~ .. .. . . . . ..... ~'~"<. :. ' : . 0'0 ...... c..~~~r·~:::j:::j::: J01...... ...256 .. ' ... : .... :::~:: . ... . . .• ••• . i..:· . . ........ .... ' ...... : 1"'!·+·· ...000 ~ ~-~-~-~-~~~~~~~~ : : : : : : .. ··················· .. .". .. . -: .. -1 . ~ . _:: : : : ::t:c:::::::: ::::::::::::::: ::: I 0.. .. ::: ~ . .: ···:···:···:···:··· ···:···:· :· :···:· ....... .:ir... .. ·+":. o .*:. ···:·---~~J$ ..~....... . . ~ .: . --~--+-.....0 10....:. .1. . 0 0 ~ LLL : : :::::::: . .~. ~.r·-r·-r·-r·· ··-r·-r·-r·· ..._~_.0 20..... . ..i .. '' '"!"''"''"'.. +. 1{ . . T·· .. ...... ··················· ...... . . ~. ...:--p . .0 ::::::::::::::::? ... .... .-.... ~ ---~· --~ ...... :. ···:··T·T·T·· ···1-·T·T-T" ···:···: .... .. ~ .. ···. -1 ....-.. :.:.._. ... ~ ......... ·--~· . .. :.... . ..... ·++--'.0 20. 0 0 .-1 . ::t:. .1. .. . .-i-_... .... .. . . .. = 0 ·-·........ .. .. = -·..·+ ..: .::t:i::: :/j:::i:::. ...... ....:.....:.:: ::::::::: 0 ~ I 11!1 I III J!t! -10. ..::: ::::::::::::::: ::: ... .... .0 20...-. ··················· . ....0 10..·" .5411-PT (Fraser) Rn = 101. .... ' ' ---:---:---:---:-----:---:---:---:------:---:---:---:--: ..... ....~.Chapter 12: Polars and Lift Plots Fig..... ... ....... ~ ~-~-~-~---~-~-~-~---~~~~ ··-r·-r·-r·-r . "'i"'i"'i"' .. ll:. .......:..63 NACA 2....:......... ~ ···\···f .·+·+--1--·1·" j"·1·+·+...... ... ..... +... = ~ 0 = -~ ~ -~....~.000 ~ .:... :... -' cY~ ...000 NACA 2.5411-PT (Fraser) Rn = 59...c ~~rrr: I -10.1. ...::::: ···:···:···:···:··· ···:···:···:··· ... ...: .. . 12. .... .0 257 . ... v·:.. .-:···:···:··· ·++. ... ·:--·:· ..... ···:. . ........ ... -:-.:. . -I ...-.0 10. . ..--+-+··\·--~ . ..0 NACA 2.···:···. ..:::::::::::::::::: ::::::::::::::: ::: : ·u·f . .0 0.-'--1 ::::.0 -10.... f-.... .. = ...... ....... .... . . : : : : : : : : : 0 -' ·+ -:· -:"·1-. : . ·1--.. ~ .. . r~ :{~: ::rrr. ~··: ·+ ...5411-PT (Fraser) Rn = 82.r:*•••~••• ......._'--::lo~-'--.. ..... = = = = = = = ::::::::::::::::: ::::::::::::::::::: :: :::::. -~ .. "'i' ·: .....:.. . .0 0..... (l . " ' ' " ...... ...' ... ... . .... ..........··1 j~~l o 0..' .............. . ....0 $2 \0 N ("') ooogg 000 00 ooo -- ~ 0~ 0 II II II II II 0 0 0 0 0 ...... ... 12.. .... ... -~.. ...... ... .. .... .· :_::.. «•!···]·· j··+· ...... . ... .. .•.. . ... . ' ..'' . . ........ ....... . ... ···:···:"·:···:··· ···:···:···: ...... : . ....+...1··-•i:0..... .. . :.... --... • • • • • • • • • • • • • • • •••:................... ..... ··. ... ···:··· .---... ..-r:>-: .\~'~~: .... . ... ~---j. ..... ... ..i.' ... .......... . ...' . .-j-~~'<i'..' ' . ..... ..... .........::_::.. 1 0 0 0 I 0 I[) 0 .... . .. ' 9'0 ' ··-·· ' ' ' '' . .. ...... Airfoils at Low Speeds c 'Q) ·a. ....' . :....... .. o·z: . ......._. . ......... .... - c c c c c 0::0::0::0::0:: ~ z ' ·--~---~---~---~·-· o·o 9'0- ....--... :---:--- .... .fr.. ............. .. .' ...••T::::::::::::: . L..' ...... .. •• : ••• : •••••• : ...' .. .. . ....... ... E 0 ...."i·'. .. ...... ~0 f. . ··················· ..~r:>-. • g lJ ::::0 .' .. .~\~ . t... . .... .. .... ...-............... ..... :..-0'-..... ...... . .. .. ..... ....... -.-.. . .. ... :r~~~< . 9'0 ....: • : ::~~\ ....i_<S>_j__:... ···:···:···:···:··· : : : : r 0 9'0- . .. ... . . ."': .. .. .....: ::.. .... .. .... "'!···!··-!--·!·...:. ..--.-·.. :. • • : ••••••••• 0.. ...... ~ '" • · :\~::: :_: ........ .: ·•· ·•· ·: ·: : · · : •·• ·:·t& ·::r+rr: ·rrrr· I lI ...... .. ..... .... .0-0-o 00 \.-.. ~J'_·..... ........ ..... .' ....:"' - \. ....kJ~·.. ' .. . . . : ......:..... •: : : :::IT: .0 ... ' .......-.. . : .. .. : • .. .. ..~0~~ . . j ..... . ....-. ....... . ... ... . . ... .:.. .... ...... ... .... . ..... I 0 N Fig. .......... ..64 .(] ......... nr: • • •·• • • • • • ·• • • r• • • [I~~±[rr ........ ....... ..-'···'·· ~ :....··· . .. .....' .. ... .. o·o ' ..... :... .. i........ .........(i <1. 1 ................ . ....c. ........ ttJr.. ...... . . ....._: .. . . ... :. . .' . .... -·· ·-· ..... ...' .: 1-::........ .. j)..... .. ···:···:·· :···:··· ........ ..' . ........ 0 M 0 . :--- : ~~.. .. ....258 .............. --.. ... ... . ....... .......... -.. .....c 8 1 t :• : : : • : : : : : : . .' . :..--. ···:···:···:···.... ....~r....' . ··~ ••• : • ..... ..................... . ... ... ..~._::+::_::. .' ._::. ··' •. ..... .... . Chapter 12: Polars and Lift Plots Fig. 12.65 NACA 64A010-PT NACA 64A010-PT (Champine) (Champine) Rn =60,000 Rn = 80,000 ~ ~~~~~~~~~~~~ ~ ~-~-~-~-.-~-~-~.~-~~~~~ o 1-oo~·:_oo·~:·~oo,_oo~·:·_ool-oo~·:_oo·~:·_ooi~oo~·:·_ooroo~·'_oo·~~·-ooi~OO~·~OO· 0 ::rrrr: ::rrrr: ::rrr: :" ~ . . .. ' .. ... ~ ·--~- --~---~- .. . .. '"(·~---~ ···j··· ... :... :... :... :...... :... :... :... ;,,, ... :... :... :..... . ···j--·t--·t--· ·-- : : : : ::::::::::::::::::: :::::::r::::~~~~ ::: "1 000:00·:···: 00: . :---:·yoo· tY"1 •<i; : . 000 ~ I : : : : it.:: : : 0 0 q o !,!T[if/111••• ::rrrr: :::::::::;t~::~: : : .... r-.. . :. --'-'--.-r-;..--'·'--'-·...;.,-r-;......:......:.....;'--1 ::::::::::::::::::: ::tt::::::::::: ::::::::::::::: ::: IIIIJ:['fffi••• q :+++L ::1.:.::1:1. :.I.EF· : · 000,000';}1=· "1 ' ' ' .. T ' ' ' ' ···:···:···:···:··· ··:···:···:···:··· ···:···:···:··· ... . ·:-:--:--:--1!''·:--:00·:--·:· . 000:·:--:-- 000 o ; ; : :I ; ; : : : : : ···:···:···:···:·· ···:···:· .. :···:··· ···:···:···:··· ... ':' 000 ;oo·:--·:000; 0 ;. •: oo:oo• 000 -10.0~ 0.0 10.0 0 I 20.0 ·"1-J}f :+::::::::::.::: :.·.::::::::::: .:: -10.0 • 0.0 10.0 C( NACA 64A010-PT (Champine) Rn = 101,000 ~ ~~~~~~~~~~~~ : : : : .... : ... ···:···:···:···:··· ···:···:···:···:··· ···:···:···:··· . . . . . .. ' .. . . . . . .. ' .. .. .... ................... .: .: :. .: ··················· .: .: .: .. ................. .: .. : ........................................................ o tY ~ .: .: .: .: .: .: :. .: ... 1... 1.. T .. 1 ...... 1... f.. T .. 1' .. .: ...: .. T..... ---~---I .. . . . ... . .. ... ... . ................... .. . .. .. . .. ................... . . . . ............... . .. :::i:::i:::i:::l::: :::i:::i:::i:::~~-~-~:· ::: 1-oo•'---foo•.:. ._:oo•+--:oo-'-o:oo+ooo_;_o:oo...:.,•:·:~7·' . . :. •:•+ooo. . :. o :•. . :. •~- '-'-+~•~oo• ·++++. 000: /++00 ·+++00 000 :::;::::::::::::::::f.\:::::::::::::;:::;:::;000000 q "·=.... ;:...:...;... ··=:...;:... ;:... ::......;:...;:...::...... .... . . . . 0 . . . . . ... .oo,ooo:oo•:oo·:·i ooo,ooo:oo•:oo•:oo• •oojoo•:oo :00 · ... ~ ::FIJI:: ::rrrr· ::ri::_::: : : 1 -10.0, 10.0 0.0 C( 20.0 20.0 259 260 Airfoils at Low Speeds ' .. . . .. .. .......... 0. l ' ; ' ' ; ; : : · · 0'0 · ; 0'0 "' ~·o- "' ~·o- 0 I 0 "'0 0 0 o 0 (") 0 N 0 0 ;; 0 ~ 0 lJ Fig. 12.66 . ... .. ... '." ........... ' .. . ..... ... ... ... .......... ····-·· .............. . ... ... .... ········· -·· ........ . .. . .......... . ~·o ~~~4-~~~~._;~~.~~~~~~~ .. . . .. . ········· ······ ··········· .................. .. .. . .. ' .. ·········· ... .. .................. . . . . . ........ .. . . . .... . . . .. . .. ... ... ... ...... . .. .'' ... ... ... . ... ... .. .. .. .. ...... .. ... . .. ... . .. .. . ... ..... .......................... ... .. .. .. ..' ... ... ... . .. ........................... .................. .. .. .. .. o·;:: .. :·'!"· ... "'(!"'!"' "'!"'!"' "'!"' ............... ' ' ' ; :.rr: ::r: ::::::::: : · : : '" ......... -·· ··············· ··········· ..................... . ... ... . .. ... ... ... ............................................................ . . . .. . . .. . ... ... .... : . ' ' . . . . . . ................................................ : :::~l:(:l:::~~t~~J~~ ~~f~~ :~"·': ... \..·0.,.. .. .. . ............ ;............... ;............... ;................. . :::1:::1:::(:f~~~~~~~~~,~~~: ::: : : ::::::: ::: : : : : : : ::: j .. j .... j· :,.~.· ;"\!1 ......... 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Ci. NACA 6409-PT (Halsey) Rn = 70,000 ~ ~.-:-..,.-~ . .-~.~.-,-.-,.~_,.....,.,_,...~ l!~~a . 1•••r 0 ~. r•••...rtnr ·l>L +.: ·+. :·++.....:.+ . :. -:-.. ... ; ......... ; 0 . . 0 ' ---:---:---:---:--- ... ;... ;... ;... ;.. . ::::::: ::: :j::: ::::::::::::::::: :::::::!:::=::::::: •= : : : : : : : : ol-'-'-...;-"'-i---lf--"--'--_,_rl....;--7-;..-;.-J 0 l•••rritl·:·r 1 ~_;_~~~~·~·....:...~~....:...~ -10.0 0.0 10.0 Ci. 20.0 Ci. 261 262 Airfoils at Low Speeds : : : ···:··- ···:···:···:··· . . . . ......... . • • • • • • • ~-+·( • • • . • 0. ... .... . ·(- · ............ [-.- .................. ~ ··T·T·-~--T·· ~ •••• " • • • • • • • • c;·o ••••••••••• - ! ' ' ' ., ......... , ... , ... , .... 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SD6060-PT SD6060-PT (Fox) (Fox) Rn = 59.J<.. o : : : : ···:--T·T·T-..... 0 0 Vl S8~oo ~~NM g "'>.Fig.113 (• 0°l o(j [)> 0'0 0 N 0 l1 Chapter 12: Polars and Lift Plots .·o 0 "" 0 0 0 0:: 0:: 0:: 0:: 0:: [!] 0 (") ~ Ul Q) cQ) -§ 1- -5 II II II II II I CL g 0 ccccc 0 1.·o :J . -o o . 12.- ooo88 888o_o_ 0°l tJ 307 . . -....... :· ··:···t· . ........01 0..... ...t. ... ·t· ··:···:· ....---r --------... ... .0% 1!1 0 & "i Rn Rn Rn Rn -·C' r.........eo 0 00 >-~· SD6080-PT (Hohensee/Murray) u... --..q· .. ·: . ::: . ·: . ··t·.. ..s.... -· . ..... .... ·: .. ··-r.... ....{jl~:: 0 ......... ... .. r........... x/c = 10%.... ·t·· ·: ..... . ... ... ·t ···.000 = 200. ....-~---~---~-.04 0.... --------............. ........ ...............17%.... -· .. :..... .-~---:--T·~·-·:··· 'i' -10 10 0 01.. . :···t· ··:· .....000 ~ 0 t:-< c N = 150.. ·: .............. ·: ....... . ...... .. --..--r··---·---·--------+--·---·---·--.....................02 0.... : .. : ... ........--r...... ....... ....... : : . ~ = 100.. . 0 .. ..... . .......g> "' Ill ........000 ... .. ·t···:···:· . ··t···: ... -. . ... .... ...... ........ ............-~ ···: ---~---~---! -... ·:... . ..... . .... ·: . 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(Wagner/Olsen) Rn = 81,000 ~~---,__,.....,.....,.....,,......,....,.....,....,...-r-,-..,..-,.....,.-i -·+·+·+·+·· ... ~ . ·~·-+·+·· ··+··t··+·· ... ., ::rrr:F ::rrrr: ::rrr: :·: 10.0 0.0 01. : : ::r;r:: : : : :[: :;: : : : : : : : : : : : : : I -10.0 f : : : ; 0.0 20.0 : : 10.0 01. SD8020-PT 0 ::: :::t::tj:: :. :::t::t::::::t::: :::t:::t:::t:::t::: : ~ : . . . : : : 20.0 Fig. 12.159 r.. .. ..-.. .. ..,.... .. .. .. ..T... .. .. ' ,·- ' ' ' ' ' ' ' :_r.i ' : . :' :r: .. .. ... .;·o- ::: ::: "]'" .............. . o·o ~ 0 0 o I 0 l() 0 0 N 0 0 0 lJ Chapter 12: Polars and Lift Plots .. .. r.. .:·~:::: . . . . : . :. :. :. ~.-~ ~.-~ ~.-~~.~~.~.- ~~~.-~..-~.- ~.j-...~j.-.. ~.~ .. ~~..,.-...~j.-.. ~~.~~.-, g 1 1 1 : : :·: :·: ::· :. ::: : : : . -:-:~ :::r" ":r::r:· "r:r:: ::r· ::r:c : : : : :' • • • • • I• • .• • I I{ :-.-~ I If • • • ' . ' . . . . . . . ..... . ............ .,............... ,....... ,... ,' ... ,... .,...... .,... ,.... ,... : =riii"' .. . ... ... 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T ..0 0 :::::::::::::1 : : ~= ::rrrr: : : : :. 1. ..:::. :--:· ·: ••( ···:·t··:--t""· ···~···:···:···:··· : : : : ~: : : : : : : :::::::::::::::1::: ~~::!:: :!:::!::: :::::::::::::::!::: 0 0 : : : : .. : --·:···:···:•· ·-· . : : : : .0 10.. T.. ···:···:···:··-. ::: ::: ::: ::: ::: ::: :::::::r:::r::r: :.. .to "' Cl> g:0>- WB135/35-PT (Blanchard) 0 a 0 b.000 't-<" ~ 0 0 ~ "' .......... ·--~·-·:···:···:··· ···:··:-..q· 0...04 0... :· --~---:··..000 Rn = 200.......-··-~ ..000 Rn = 150. ---~---~---~---~--- . .05 I -10 10 20 . .000 Rn = 100.'\--' ·--~·-·j··-~---~--- u 1() 0 "'"' ~ ..... : ... . ...···:·-·:···:···:·--r··:·--:--·:---:--· ..... : ..... ··:· .....---=--·:··-..... ·- .02 eel ... .. :. "" "" . 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III/I III III I ::t:::::!::Ll::!:::i:::!::::::!:::!:::::::!::: ::::::j : : ::Lt:L: :::j:::j::::::L: .. "l 0 I . ...... ..0 WB135/35-PT (Blanchard) Rn = 101...... i· . +·-~·-+·· ··+·+ ..... . .... ~--· ~ ·--~.: . .: :: ···:··· ···:··+·· . .. ::rrrr: ::rrij. .....·t ......0 10... . :.. .. .. ---~---~·-·j--· ---~·-·t--·~·-·t--· II') ••• : ••• : ••• : ••• : •••••• : ••• : ••• : ••• : •••••• : .. .. 0 ::::::: ci ... : .. -10.. .. . ... .000 ~ ~~~~~~~~~~~~ (Blanchard) Rn = 82... .. ... . : .0 20.. .. ---:---:---:--.... : ..0 10.Chapter 12: Polars and Lift Plots Fig... :... :. :.. .---:---:---:---:--- d .:. ..... .. : . .0 0. . 0. .... 0 :::-::: ::::::::::: ... "'!:·.:. ·:--..... ....:-. )"·!~ ci : : : : /. : .... ..j. . : .. ·: .. . . ~ "'" ..-··:···:···:· .. .... :: ::: ::: :: ... .. :· ..· WB 140/35/FB-PT (Blanchard) 8 <!> 6 'i' Rn Rn Rn Rn 0' ~ =100. .03 cd 0... . ... ·: .:.. .. : .........01 0.. ::c::::c:::r:1:::1::: ::: :::1::: ::: :::L: :::I:J::: :::: : .... ' 0 0 ~ . II I{) 0..000 =200. : .···:···:·· : . .... .000 =300... ··: ..._ :::r::. ·: ..04 0. : . ::···:···: .. :· .000 ~ 0 t-< 0 ~ N .go "' "' f} I{) . . :···...... ·: . . ....05 -10 :: ' 10 0 01 : ") 20 oc.02 0. Q1 00 > . .000 =150... ...... . . ..00 0.. .... :· ... ~ ~rlllr:::·I L cl cY ... :.... ::: ::: :: .. ·:· . .· .... : ...... ::...···:···: ...u::::::: 0 .. .... ·: .. ·:· . 450 0.0109 0.85 0.0176 1.426 0.991 1.88 -1. c.522 0. c.0162 1. AQUILA-PT Rn Fig.088 c. 0.166 0.0180 -1.126 0.925 1. = "' c.005 0.35 Rn = <> -4.251 0.538 0.168 0.116 0.0252 0.0095 0. 26 0.015 0.3 Rn 61300.546 0.445 0.537 0.33 1.5 Rn 60400.33 11.439 0.29 0.0202 1.448 0.29 c.100 0.0179 0.0367 Rn = 148600.934 0.33 1.86 0.Chapter 13: Tabulated Data = 301100.407 0.0099 0.113 0.34 1.778 0.20 0.0152 1. <> -1.27 0.18 0.526 0. c.0128 1.0458 1.161 0.0081 0.152 0.133 1.0175 1.691 0.25 5.742 0.85 0.23 3.340 0.87 0.233 0. 12.36 1.90 0.0189 0.899 0. = a -2.0471 13.399 0.299 0.20 2.0294 1.0507 Rn 101100.147 0.84 0.16 1. 788 0.740 0.24 4.065 0.0130 4.0120 DAE51-PT Fig.015 0.31 8.0209 7.0670 c.24 5.002 0. <> -1.0151 1.346 0.0475 14.87 -0. c.217 0.18 0.17 0.18 6.84 0.18 1. 0. cd " c.0651 310400.0269 -1.20 0.23 3.490 0.0153 2.188 0.726 0.0188 0. 764 0.0164 0.0198 -0.0143 -2.0136 0.27 4.31 8.91 -2.89 0.292 0.0132 0.87 0.34 0.210 0.0152 1. c.29 6.210 0.0215 7.116 0.0573 Rn 150500.836 0.782 0.0115 = 1.0130 0.216 0.0104 0.646 0.36 1.170 0.148 0.0097 0. c.061 0.34 1.172 0.21 0.0120 0.198 0.109 0.834 0.34 13.35 12.104 0.0119 0.25 0.28 0.0154 0.691 0.22 2.0197 5.0099 0.25 0.711 0.340 0.6 Rn 100100.984 0.27 5.0134 0.545 0.305 0.916 0.36 1.0233 8.0271 6.0216 1.0303 4.0101 0.041 0.143 0.963 0.595 0.0346 Rn = 203900.0364 Rn 102600.346 0.38 1.748 0.31 7.33 11.841 0. c.0214 1.0223 -0.146 0.699 0.0119 0.86 -0.120 0.0283 8.086 0. c. c.120 0.402 0.86 0.35 1. <> -3.32 1.0191 4.32 = = = a c.0208 -1.0214 8.0099 0.882 -2.15 0.0372 1.0246 1.26 0.235 0.17 0.24 0.90 -1.0385 12.448 0.33 1.092 1.0205 -0.30 6.0111 0.0225 9.0127 1.18 0.0640 Rn = 203200.0221 1.26 0.0213 1.0390 Rn 153500.23 3.929 0.89 0.0108 0.22 0.006 0.0358 6.0091 0.0248 1.33 1.0140 0.0239 1.36 1. 0.0184 3.28 7.28 0.0128 359 .352 0.0104 3.0138 5.529 0. c.82 0.641 0.0190 6.37 1.0268 1.794 0.006 0.0373 8.0206 6.95 -0.27 0.896 0.921 0. <> c.833 0.0224 2.21 4.0120 0.31 6.686 0.0118 0.944 0.605 0.33 9.172 0.0313 1.064 0.0246 0.627 0.92 0.0196 -0.0227 0.28 6.336 0.0724 301200.0198 6.0326 10.33 9.0190 3.26 4. <> -2.0113 0.0098 2.0184 4.283 0. -2.0281 0. cd -2.0182 8. 744 0.86 0.0228 -2.340 0.176 0.29 0.0139 0.93 -3.087 0.0090 0.36 1.004 0.82 9.248 0.0145 0.0087 0.0296 0.496 0.0224 -1.22 0.20 2.0186 5.0166 3.0247 0.34 9.0354 5.0143 0.32 1.31 8.19 0.28 7.0245 1.834 0. -2.0208 0.28 0.33 1.911 0.0427 1.22 3.19 0.014 0.21 2.26 5.34 10.542 0.0197 0.0359 11.92 -1.0134 0. a c.510 0.1 Rn 59400.24 0.0113 2.148 0.30 7.29 0.37 1.30 0.0218 0.29 0.0191 7.0091 0.0286 0.0187 0. c.607 0.298 0.365 0.0078 0.82 0.302 0.36 11.0128 0.0474 DAE51-PT Fig.0238 1.21 0.280 0.918 0.0256 0.0192 1.0084 0.32 8.730 0.0112 0.13 2.268 0.85 0.741 0.34 4.839 0. 0.0272 0.34 1.37 1.989 0.35 1.0103 0.098 -0.0352 10.23 0.245 0.177 0.84 -0.0196 5.22 3.0388 1.0150 6.17 0.0125 0. 0.043 1.240 0.924 0.33 1.0194 1.0121 0.32 1.644 0.0404 7.0314 12. <> -1.0332 11.337 0.25 0.0285 10.159 0.93 -2.829 0.0179 2.87 0.34 9.83 0.166 0.085 0. -2.010 1.0187 1.0098 1.84 0.0349 1.102 0.069 0.31 1.312 0.94 -0.24 4.31 c.0101 0.28 7.079 0.35 1.35 10.28 5.637 0.337 0.0141 0.34 11.916 0.625 0.336 0.90 0.32 1.26 5.166 0.0279 0. a Rn = cd 0.186 0.427 0. 791 0.0311 9.77 6.34 1.0313 2.041 0.0265 11.041 0.011 0. 72 3.34 1.18 0.156 0.184 0.0166 -1.0539 1.874 0.22 0. <> -3.87 0.20 2.138 0.0179 4.91 -1.600 0.90 0.88 0.047 0.0224 0.292 0.191 0.34 cd 0.0209 1.261 0. <> -3.0092 0.043 0.0383 Rn = 203800.241 0.0251 9.0273 0.0344 0.0271 1.87 0.0298 3.481 0.31 7.0193 9.0222 10.28 0.31 7. 12. <> -3.0134 -0.26 4.922 0.137 0.648 0.960 0.23 5.501 0.157 0.174 0.637 0.34 9.92 -0.18 1.0500 1.90 -1.007 0. 732 0.25 0.17 1.33 0.159 1.438 0.0122 0.37 0.0170 7. c.145 0.0282 -2.0124 0.0299 11.0324 3.0131 0.0238 -0.0128 -1.19 3.0259 10.181 0.0120 -0.135 0.14 1.34 10.069 0. cd 1.22 0.32 8.0399 Rn = 101400.27 4.0141 1.0230 0.0266 7.0356 13.91 0.0181 0. c.272 0.232 0.37 1.89 -0. c.0254 0.34 1.0278 1.35 1.83 0.399 0.17 0.137 0.525 0. c.0184 1.30 7. -0.20 0.287 0.0291 4. 12.295 0.0172 2.088 0.96 -0.215 0.235 0.0305 9.35 10.0162 -0.87 -0.163 0.844 0.26 0.91 -1.24 0.481 0.30 1.0179 0.36 14.34 0.0284 5.17 1.23 0.0239 8.646 0.18 3.28 0.198 0.0422 10.35 1.22 0.064 0.742 0.33 1.157 0. c. c.876 0.37 0.541 0.0383 9.213 0.28 6.0152 CLARK-Y-PT Fig.0085 0. 12.31 8.24 5.815 0.0110 0.20 0.0151 1.0124 3.0114 = = c.596 0. 19 c.0190 0.135 0.245 0.563 0.156 0.0456 0.1178 Rn = 100200.171 0.767 0. "' -0.93 -0.16 2.149 0.0182 0.076 0.0270 0.226 0.045 0.17 1.134 0.0204 0.00 -4.861 0.250 -0.950 0.648 0.154 1.0094 0.0488 DF103-PT Fig.0418 1.0163 0.87 0.24 5.Dl!l5 0.137 0. cd 0.0757 DFlOl-PT Fig.22 7.33 12.0273 1.0550 0. c.201 0.450 0.0113 0. 12.19 4.21 4.612 0.440 0.029 1.0149 0. 71 6.424 0.229 0.181 cd 0.900 0.24 5. -0.31 7.473 0.283 0.18 4.0101 0.049 0. cd -0.842 0.351 -0.34 12.0137 -0.20 3.9 Rn = 59100.96 -1.0393 14. c.186 1.88 c.93 -2.29 8.060 0.0107 0.77 Rn = " c.0362 301600.0109 0.93 -1.0111 0. = " -2.0126 0.110 0. 78 6.562 0.181 0.414 -0.872 0. 1.141 0.293 0.649 0.237 0.012 0.888 99700.0096 0.0128 0.0211 0.624 0.613 0.112 0.145 -0.0108 0.80 6.0136 0.106 0.360 0.0326 0.17 3.976 1.0138 0.0100 0.035 0.34 13.0502 Rn = 197900.26 7.0092 0.0232 0.92 -0.0102 -0.0275 1.94 -0.896 0.260 0.434 0.814 0.34 0.008 0.0150 0.598 0. 73 3.0205 0.0267 0.074 0.088 1. 0.28 8.0350 0.34 0.841 0.31 11.0211 0.0401 0.104 0.0114 -0.27 10.0603 305500.22 6.91 0.238 -0.380 0.0193 0.819 0.144 0.26 7.1833 153700.053 0.82 9.640 0.41 0. "' -2.766 0.27 8.375 0.055 0.360 Airfoils at Low Speeds 1.77 9.0184 0.28 9.0216 0.0107 0.524 0.638 0.93 -1.18 3.21 5.022 0.0091 0.97 -3.731 0.43 -0.32 11.31 10.29 9.0157 0.84 c.0093 0.129 0. 70 5.26 7.691 0.65 2.331 0.0268 1.018 0.336 0.701 0.0195 -0.0240 -0.136 0.0174 0.052 0.0135 0.807 0.0463 101700.35 Rn = "' -3.679 0.474 0.29 9.0103 0.008 0.0144 0. 76 8.0110 0.652 0.34 cd cd 0.0174 -0.98 -2.91 -0.0155 -0.908 0.973 0.31 Rn = "' -6.0238 1.126 1.00 -4. 759 0.0270 0. cd -0.0179 0.197 -0.19 3.34 10.30 9.26 7.0137 0.33 13.32 10.855 0.23 5.34 15.202 cd 0.214 0.13 1.20 3.33 11.185 0.016 1.856 0. cd -4.485 0.0217 1.143 0. " -3.0208 0.0110 0.0160 0.0826 149100.171 1.16 1.0146 0.064 0.0184 0. -0.0307 1.0204 -0.22 4.0083 0.27 8.98 -1.66 3. 799 0.32 11.24 6.0159 0.32 10.19 4.036 0.0148 0.24 6.117 0.26 7.29 -5.17 2.0139 0.0204 1.0109 0.23 6.0335 1.0383 1.0115 0. c.571 0.0203 0.31 7.89 0.0429 1.74 3.28 8.0201 DF102-PT Fig.13 1.338 0.435 0.0170 0.121 0.323 0.0147 0.0221 0.526 0.016 0.0123 0.202 0. c.179 0.124 0.156 0.947 1.91 0.0126 -0.923 1.0155 0.16 3.0547 1.265 0.489 0.0129 -0.193 0.354 0.0152 -0.0143 0.0136 1.0195 -0.932 0.33 11.0367 1.0333 1.25 7.0275 1.0149 0.0242 0. -0.939 1.0322 0. c.325 0. cd 0.786 0.12 2. "' -3.0091 0.0127 0.055 0.564 0.0245 0.0078 0.0173 0.31 9.649 0.15 1.646 0.107 -0.0223 0.29 7.25 6.97 -1.96 -1.97 -2.458 0.0132 0.22 5.0143 -1.0128 0.16 2.875 0.0321 1.23 4.613 0.339 0.0178 -0.03 -2.0157 1.0103 0.24 6.95 -2.0245 0.34 13.0110 0.21 4.0214 0.78 6.35 14. 721 0.651 0.0095 0.0152 0.294 0.35 -0.264 -0.31 9.27 7.815 0.896 0.187 0.25 4.26 8.87 0.12 1. = " -2. 71 3.0262 0.0118 0.780 0.0242 0.787 0.730 0.970 0.067 0.469 0.167 0.947 0.83 9.33 Rn c.25 6.95 -2.101 0.31 14.0196 -0.194 0.0206 0.0158 0.0124 0.17 2.0145 0.34 12.115 0.0149 0.0086 0.0462 300900.0195 0.882 0.97 -3.33 Rn = 1.695 0.99 -3. a c.0169 -0.027 0.932 0.17 3.19 4.132 cd 0.29 10.235 0.36 0.435 0.88 -1.0120 0.23 5.0231 0.0184 1.0209 1.375 0.0219 0.0131 0.35 14.037 0.0103 0.463 3.89 0.97 -2.0085 0.88 0.8 Rn = 58100.050 O.0119 0.541 0.99 -4.0214 -0.0122 0.81 9. 737 0.980 0.505 0.232 0.19 1.243 1.0177 0.047 0.14 1. c.30 9.93 0.92 -0.0228 1.28 9.10 Rn 62500.0127 0. cd -0.34 12.24 4.0133 .27 10.26 7.0083 0.226 0.0089 0.0284 0.395 0.0176 0.0214 0.0144 0.808 0.0227 0.0147 0.18 2.040 0.0152 0.0344 0.0175 0.17 2.0358 0.81 11.20 4.18 3.0500 0.12 1.89 -1.95 -1.0294 0.34 13.841 0.0118 0. 777 0.0106 0.0107 0.0316 1.0126 0.31 10.93 -0.85 12.97 -3.0118 0.0165 0.800 0.194 0.97 -2.14 1.89 -0.042 1.354 0.09 .0189 1.0162 0.0099 0.0094 0.24 7.525 0.135 0.29 8.0189 0.0188 0.15 2.24 5.04 71 200900.89 0.32 12.90 -0.0095 0.23 6.31 Rn = "' -6.083 0.86 0.0160 0.0102 0.0347 cd -2.938 0. "' -4.0177 0.123 1.21 4.31 10.097 0.0218 0.0260 0.34 12.0123 0.96 -3.131 0.351 -0.0282 1.95 -1.0115 0.19 3.31 10.31 Rn = " -3.0170 0.338 0. cd -0.22 4.0218 0.34 13.0148 0.112 0.0160 0.0156 0.213 1.293 0.688 0.775 0.16 1.93 -0.079 0.176 0.167 0.235 -0.0194 0.0455 0.556 0.29 8.0220 0.048 1.0420 Rn = 201400.33 11.306 0.95 -1. -0.0297 1.283 0.90 -0.92 -1.079 0.826 0.0144 0.27 8.909 0.0098 0.0395 0.379 0.0379 1.551 0. 12.28 9.093 0.0188 1.0327 0.974 0.0105 0.28 9.17 2. 12.32 10.43 0.0248 0.25 c. Rn = "' -5.0087 0.97 -2.0201 1.699 0.23 5.0105 0.94 -2. cd -0.0265 1.34 c.22 5.0260 1.295 0.175 0.417 0.34 10.25 6.0238 1.99 -4.022 0.26 6.0136 0. 0135 1.97 -5.501 0.0126 0.28 6.980 0.91 -1.86 0.20 3.019 0.0265 0.0111 0.811 0.0241 0.92 -1.0216 0.76 9.715 0.203 0.88 -0.25 cd 0.197 0.911 0.0096 0.0560 13.372 0.0196 0.0120 0.121 204200.097 0.21 4.0166 0. = c.062 0.28 7.0102 0.87 0.0120 0.047 0.0127 0.0177 0.0127 0.0219 1.89 -0.128 0.103 0.0115 0.0197 0.26 6.0130 0.98 -3.42 0.0093 0.0158 1.93 Rn -2.87 0.0310 0.120 0.89 -0.0217 0.618 0. c.24 6.34 10.0185 1.0129 0.720 0.006 0.0143 0.0226 0.0208 0.111 0.21 4.04 70 12.80 Rn 0.516 0.89 -0. "' -2.88 0.0894 El93-PT Fig.0323 0.30 7.0340 0.0223 1.24 6.0539 Rn 97500.0195 0.0141 0.0084 0. " -0.0179 0.88 -0.0519 c.90 -0.23 5.119 1.401 0.205 0.23 7.01.645 0.0183 0.185 0.0279 9. 12.0228 1.072 0.212 cd 0.265 -0.0198 0.0374 203200.34 12.32 10.045 0.18 2.024 0. -0.0358 0.20 3.21 5.92 -1.72 6.030 -1.18 2.0109 0.0489 0.192 0.908 0.0317 0. 12.0097 0.0196 0.0293 0.606 O.961 0.063 0.16 1.0100 0.0195 0.100 0.92 -1.0247 0.30 8. = Cc " -0.33 11.0287 1.33 11.0097 0.126 0.0168 0.914 1.0227 1.0231 1.317 0.180 0.0139 0.66 3.651 0.706 0.277 0.0101 0.34 1.26 7.0288 1.15 1.926 0.15 1.0108 0.00 -6.0415 0.876 0.0107 0.33 8.0260 1.366 0.0170 0.0102 0.264 0.96 -2.598 0.22 4.017 0. = c.021 0.080 0.25 7.29 -0.27 10.0166 0.97 -5.0114 0.26 6.410 -9. -0.210 0.0547 148500. " -0.0221 1.79 12.373 0.0326 1.599 0.13 1.0253 0.29 11.218 0.524 0. = c.449 0. = c.31 10.0222 0.0414 -0. = c.0932 -0.078 0.0116 0.16 2.28 8.0225 0.314 0.34 12.289 0.87 0.15 1.208 0.30 7.19 2.78 10.0149 0.040 0.959 0.515 0.199 1.0160 -0. 12.0339 11.36 10.86 0.557 0.262 0.74 6.380 0.645 0.234 0.35 11.90 0. 727 0.101 0.803 0.0220 1.0174 1.33 Rn cd 0.29 9.29 7.14 1.0157 0.32 10.33 10.93 ·1.924 0.004 1.28 6.372 0.133 0.0099 0.944 0.90 Rn cd 0.0128 361 .110 1.34 Rn "' cd 0.0259 1.0122 0.28 7.25 5.0106 0.0092 0.34 12.91 -1.16 3.33 1. " -0.898 0.0129 0.219 0.0161 0.514 0.89 0.0217 0. " 0.32 9.19 3.034 99500.12 2.32 -0.0096 0.0198 -0.15 2.373 0.923 0.0146 0.144 -3. "' -0.0123 0.43 0.0289 0.27 7.0236 0.0172 0.045 0.821 0. " 0.0148 -0.0402 0.076 0. 771 0.20 4.33 9.33 9.0424 0.0234 0.32 1.555 0.248 0.295 0.23 7.35 Rn 0.048 ·1.0292 0. 710 0.915 0. = " -2.608 0.993 0.73 6.0402 0.416 0.665 0.94 -1.021 1.0192 -0.0102 0.67 3.0242 0.455 0.16 2.090 0.95 -1.86 0.14 1.95 -1.0154 0.357 0.220 -3.0270 0.25 7.21 4.223 0.31 8.0112 0.15 1.0138 0.30 8.77 9.583 0.33 0.0197 0.868 cd 0.0134 0.18 2.972 1.31 9.972 0.0097 0.22 4. -0.17 2.0208 0.0104 0.80 12.30 0.33 1.77 9.31 8.0157 0.0123 0. = ·4.26 6.0147 0.0231 0.90 -0.0128 0.594 0.74 7.94 ·2.24 5.88 0.33 9.003 0.0209 1.207 0.20 3.0138 0.D105 0.0430 0.0190 0.20 4.0382 303100.619 0.0085 0.505 0.0149 0.23 5.414 0.0388 7.0228 0.0109 0.22 5.0173 0.0291 0.0118 0.018 cd 0.258 0.0122 0.0100 0.33 11.0161 0.063 1.238 0.0127 0.298 0.211 0.20 2.34 11.104 0.18 4.0277 0.629 0.0110 0.0144 0.506 0.24 4.29 8.0116 0.0158 0.93 cd 0.962 0.0131 0.0586 304800.0348 207900.22 5.92 -1. = -2.858 0.564 0. = c.Chapter 13: Tabulated Data 0.0389 0.67 3.863 0.20 4. c.096 0. " -0.13 Rn 59800.0232 149400.0334 8.0137 0.32 -0.0091 0.30 Rn = " ·2.0377 0.135 1.188 0.085 0.0082 0.364 0.0098 0.89 -0.0109 0.34 cd 0.198 0.0289 0.276 0.0141 1.0156 0.13 1.104 0.283 0.0226 0.315 0.0154 0.0207 0.991 1.109 0. = c.29 10.89 -0.616 0.161 -0.551 0.0282 1. " -0.179 0.076 0.85 0.20 3.03 Rn -8.34 9.0127 0.477 0.0106 0.181 0.0122 0.29 cd -0. = c.013 0.69 4.31 10.89 ·1.741 0.0242 1.468 0.28 -0.21 4.0210 0.23 4.23 5.11 Rn 60200.18 3.0092 0.0107 0.19 2.817 0.0285 0.960 0. 795 0.17 2.25 8.412 0.324 0.25 6.0170 0.687 0.0083 0.004 0.13 1.502 0.0282 10.074 0.34 0.0178 1.34 1.202 -3.069 -2.32 8.906 0.740 0.0112 El93MOD-PT Fig.18 4.397 0.750 0.0231 0.0179 1.26 5.860 0.25 6.0281 202100.44 1.34 c.650 cd 0.95 -2.27 10.0119 0.046 0.03 -3. = " -2. 732 0.106 -3.0201 1.0140 0.1195 -0.33 10.960 0.0206 0.745 0.996 0.31 8.081 1.164 1.15 1.0194 -0.79 13.090 0.32 Rn 01 cd 0.121 0.0180 0.0107 0.85 0.0226 0.657 0.830 0.987 1.155 0.0177 1.388 0.98 Rn cd 0.19 3.20 3.0085 0.33 1.0210 0.12 Rn 100300.125 1.94 -2.24 cd 0.0115 0.0312 E205A-PT Fig.111 0. 763 0.090 0.512 0.111 0. = c.04 7 0.88 -0.31 9.0134 0.0333 150600.0107 0.19 3.164 0.175 0.0119 0.0088 0.294 0.24 5.21 4.0188 0.0266 0.811 cd 0.16 1.32 9.0334 0.014 7 1.081 0.392 0.35 10.082 0.16 1.94 Rn -2.985 0.0125 0.965 0. 0.0338 307800.0222 0.514 0.31 7.0354 c.828 0.0147 0.32 11.34 11.780 0.125 -3.95 -2.381 0.083 0.079 0.930 0.186 1.098 0. 0201 1.84 0.91 -1.960 0.061 0.13 3.19 2.563 0.087 0.95 -0.0277 -4.0222 1.0274 -1.0242 1.22 0.335 0.0160 " E205B-PT Fig.995 0.0582 " -4.94 -0.0108 0.456 0.33 1.D103 5.0128 1.0864 -0.92 0.0210 0.0763 -4. Ct cd -0.25 4.258 0.186 0.0233 E214B-PT Fig.247 0.392 0.283 0.040 0.39 1.0172 -2.029 0.88 0.19 1.24 5.255 0.86 13.809 0.0280 10.23 0.0177 O.89 0.D136 -0.917 0.92 0.0154 -2.02 -4.267 0.622 0.372 0.126 0.0302 1.0210 8.398 0.672 0.0120 0.25 0.0196 1.861 0.0296 1.36 11. cd Ct -5.002 0.806 0.89 -1.0171 0.149 0.172 0.076 1.493 0.0108 0.95 -0.27 6.243 0.068 0.21 0.28 7.083 0.30 0.32 0.39 1.0136 -1.0665 146700.0200 0.817 0.0207 7.87 0.35 11.0221 1.88 -0.0192 10.614 0.90 -0.206 0. 12.457 0.129 O.25 0.569 0.0091 0.19 0.278 0.28 6.28 0.33 10.734 0.0300 6.35 1.0143 8.199 0. cd Ct -6.86 0.24 5.0092 0.0180 0.24 4.108 0.0313 1.27 5.957 0.20 3.362 Airfoils at Low Speeds 5.34 9.17 Rn = 59800.97 -0.34 1.0421 -5.24 0.16 1.0547 1.652 O.94 -2.844 0.060 0.0137 -1.081 0.997 0.24 5.38 1.28 0.0115 0.16 2.0182 0.23 0.017 0.38 Rn = " -4.19 2.882 0.24 7 0.0308 2.0190 2.282 0.257 0.0463 1. cd Ct -7.0382 1. Ct cd -5.0244 1.32 10.0126 -0.0151 0.16 Rn = 101600.94 -0.0123 0.967 0.0099 0.38 1.37 11.083 0.0235 " " -4.31 9.87 0.0207 2.0327 -2.28 0.912 0.0090 0.32 11.22 0.045 0.0226 9.0106 0.0364 3.219 0. 12.0084 1.215 0.0138 0.0727 -5.0329 11.0252 -3.97 -0.101 0.176 0.19 2.27 0.0081 0.0260 1.37 1.0192 1.0146 0.28 7.988 0.90 -2.91 -0.131 0.0369 -0.19 -0.33 8.0120 0.87 0.118 0.882 0.0456 Rn = 105200.15 1.137 0.87 0.19 0.0101 0.877 0.89 0.0197 0.28 0.22 0.099 0.0098 0.0280 1.28 6.0118 0.0095 0.084 0.30 8.0135 -0.24 4.00 -0.32 Rn = " -4.696 0.0136 0.271 0.0133 0.26 0.26 0.675 0.301 0.279 0.0235 -0.0461 Rn = 202600.303 0.0435 -3.163 0.566 0.305 0.84 0.0099 4.86 0.194 0.14 0.32 9.0481 -0.718 0.0357 -3.209 0.0162 0.0171 -0.30 7.050 0.156 0.0108 0.37 12.0141 1.0178 1.0324 5.230 0.31 0.39 1.019 0.0226 0.0131 6.0276 -2.33 1.98 -4.17 1.0267 9.061 0.979 0.684 0.88 -0.0113 0.378 0.0134 0.383 -5.0082 0.0089 0.30 -0.28 8.28 0.23 5.30 7.31 1.357 0.0588 304300.0091 2.255 0.091 0.118 0.728 0.0206 8.86 10.154 0.0139 0.244 0.541 0.0256 12.33 8.0584 206600.27 0.0319 7.956 0.317 0.082 0.0134 1. Ct cd -5.95 -3.95 -2.662 0.344 0.0116 " " .16 0.15 1.0107 0.23 3.0198 0.651 0.570 0.318 0.16 1.082 0.0104 0.0225 1.417 0.0423 1.32 1.176 0.93 -0.600 0.84 0.30 0.094 0.21 3.364 0.041 0.0258 9.117 0.31 9.16 0. Ct cd " -0. Ct cd -0.916 0.0149 -0.0278 Rn = 301800.88 -0.018 0.096 0.927 0.26 0.19 0.221 0.00 -6.342 0.0109 -0.95 -2.031 0.0179 0.748 0.0116 0.350 0.35 1.31 8.407 0.743 0.0158 0.0104 0.031 0.567 0.294 0.523 0.0185 0.193 -4.496 0.0086 2.0111 0.049 0.37 1.0277 3.0439 1.28 6.625 0.006 0.070 0.98 0.0248 2.86 -0.0275 1.029 0.136 0.342 0.0216 -1. Ct cd " -0.0101 0.836 0.027 0.0117 0.090 0.840 0.096 0.86 0.444 0.0214 3.0498 1.30 0.17 0.0110 0.88 0.828 0.301 0.0101 0.33 1.0172 1.0247 0.059 0.29 6.076 0.00 -0.0118 7.36 Rn = " -7.0186 -2.473 0.070 0.010 0.273 0.19 0.0306 9.081 0.93 -3.838 0.0130 0.0183 1.28 6.261 0.464 0.97 -3.0212 0.795 0.0116 0.37 Rn = " -4.26 5.367 0.33 10.0287 8.0122 0.147 1.0227 1.22 0.193 0. 12.182 0.0127 0.0263 0.109 0.32 1.899 0.0338 1.0207 4.19 0.468 0.0227 0.39 -1.0205 5.94 0.0343 Rn = 151100. Ct cd -0.255 0.0210 7.0120 0.31 1.988 1.30 8.22 0.0132 0.578 0.0380 5.072 0.24 4.467 0.736 0.305 0.0175 1.0817 201800. Ct cd -0.366 0.20 0.0112 -0.29 0.89 -0.86 0.37 12.86 0.935 0.0112 6.38 1.260 0.0289 8.0106 0.025 0.238 0.0208 4.33 9.0209 1.30 7.99 -3.21 0.0148 0.594 0.066 0.0219 6. Ct cd -0.065 0.95 -3.36 10.836 0. 763 0.331 0.16 0.0171 1.99 -0.0118 0.30 8.0227 0.0211 6.024 0.0145 1.91 -2.25 0.0200 0.0230 -3.361 0.005 0.0175 1.0172 1.32 11.264 0.0230 11.068 0.90 0. 775 0.34 1.268 0.0145 1.532 0.90 0.84 0.0266 10.95 -3.20 2.636 0.0175 0.96 -0.31 1.0211 5.84 0.392 0.349 0.014 0.0181 -0.0388 Rn = 99300.0207 1.88 -1.31 1.79 7.893 0.35 10.0110 0.0093 0.34 9.75 4.32 1.169 0.31 Rn = " -5.0105 2.0293 0.22 4.0161 9.27 0.94 -3.473 0.555 0.0152 -1.0138 0.92 -1.0407 4.20 2.0345 6.0105 1.291 0.30 0.0302 7.0135 7.324 0.32 1.0133 8.235 0.24 0.31 1.0133 0.D75 0.01 -0. 735 0.14 Rn = 63600.30 7.733 0.0366 11.0694 -4.0216 -0. 782 0.86 -0.875 0.088 0.0108 3.0159 1.209 1.505 0.0758 1.0336 4.99 -0.0090 0.0166 0.108 0.25 4.054 0. cd Ct -1.0338 1.281 0.31 12.26 5.0310 10.89 -1.088 0.0674 299000.18 0.0093 3.92 -2.0112 0.23 3.1016 E214A-PT Fig.160 0.051 0.91 -1.0200 -3.85 0.88 0.692 0.0195 3.20 3.0094 0.327 0. 705 0.160 0.0089 1.291 0.17 0.0548 301000.0203 0.80 0.026 0.666 0.538 0.84 0.0212 1.975 1.37 11.0561 0.0194 0.150 9.029 0.0098 = c" E214C-PT Fig.30 0.062 0.93 -2.89 -1.25 0.0205 9.131 0.0170 -3.0155 2.35 9.634 0.84 1.070 0.0123 -0.27 0.18 3.908 0.23 0.34 1.24 4. Q Ct -5.112 0.0144 0. 729 0.88 -0.0485 1:271 0.99 -4.978 0.171 0.590 0.925 0.31 cd 0.91 -2.0164 1.239 11.0145 0.0269 0.93 0.0098 -1.0278 0.0191 1.623 4. "' -5.0123 1.0241 0.340 0.01 0.131 1.0307 Rn 201700.87 Rn "' Rn = "' -0.328 0.553 0.178 0.369 0.002 -1.17 0. 77 4 5. 12.0162 8.96 -3.0157 7.0089 -0.o215 8.060 0.94 -0.789 O.29 1.457 0.36 13.790 cd 0.34 c" 0.281 0.0112 5.20 Rn 104200.0369 0.81 5.36 1.0185 3.0183 5.880 0.30 1.876 0.916 0.011 0.0113 0.286 0.36 Rn 0.0134 0.0186 0.011 0.28 1.89 0.0204 -2.0189 9.21 3.095 0.35 1.230 0.089 0.0094 2.0388 1.87 11.261 0.0092 0.95 0.89 0.0133 -2.0232 0.0112 0.0400 1.077 0.692 0.0190 0.0123 0.230 0.0213 1.91 0.003 0.321 0.0102 E214C-PT Fig. Ct "' -0.0103 0.34 8.0305 1.87 8.92 -2.287 0.36 14.38 9.0158 0.0691 198400.598 0.20 0.36 = " -0.287 0.0281 1.899 6.160 0.36 1.36 1.829 O.21 0.0136 1.27 0.32 1.32 1.317 0.0118 6.057 0.0189 0.89 cd cd 0.27 0.0233 0.416 0.22 4.0228 = "' = cd cd E214C-PT Fig.0559 1.899 0.0317 -2.0199 0.0120 0.33 1.462 0.17 0.18 1.99 cd cd 0.878 0.105 0.0086 -0.33 6.29 7.067 0.670 0. 781 0.17 1.28 3.0142 o.88 0.0166 1.0146 7.70 0.202 -0.36 Rn = "' -6.92 -0.25 0.763 0.0203 1.24 4.0545 1.0121 1.0122 5.460 2.115 0. = "' -0.324 0.87 8.347 0.32 1.37 1.28 1.35 1.921 0.82 0.223 c" 0.0797 200800.19 1.0210 7. Ct 0.82 1.34 1.242 -3.89 -0.0296 0.17 0.0561 1.0271 0.075 8.188 0.31 6.26 4.678 0. Ct cd -0.0153 1.262 0.060 1.0251 0.0148 0.0421 0.0184 9.80 0.0202 0.179 0.64 0.093 0.083 0.21 Rn 100600.573 0.166 0.276 0. = " = q " -0.0094 0.27 3.0135 1.0149 0.0327 Rn 302800.34 6.0316 0.0142 = -0.147 0.0370 0.31 1.38 12.36 10.32 9.0373 1.0207 1.106 0.23 5. 711 0.345 0.295 0.008 -4.0233 1.125 0.85 0.0253 10.0155 1.0441 Rn 301800.31 7.187 0.0120 -2.0100 -1.0235 0.273 0.0200 -4. 771 0.89 0.232 14.271 0.276 0.415 0.30 6.0518 0.97 -3.0200 9.248 0.283 0.668 0.0481 Rn 149400.987 0.78 Rn 0.0777 E214C-PT Fig.0344 0.0146 0.266 1.337 0.0087 0.931 0.33 1. 12.210 0.35 1.064 0.392 0.87 0.0288 0.29 9.37 1.492 0.0569 0.37 1.0598 0.0213 1.89 0.672 0.195 0.98 -0.232 0.0133 6.267 0.26 0. 12.18 Rn 102700.44 -0.463 0.0214 5.85 0.0153 6.566 0.0257 -4.0214 1.114 0.82 8.134 0.0283 363 .0150 1.0109 0. 75 2.85 0.23 0.36 1.336 -0.Dl03 4.88 11.0326 0.426 0.152 -0.30 1.293 0.392 0.82 5.0407 1.866 0.0208 0.87 8.27 1.084 0.067 0.285 0.35 12. Ct 0. 779 0.931 0.008 0.28 5.86 1.195 0.20 3.779 0.0145 1.0207 9.0106 0.0089 1.285 -0.24 Ct -0.0156 8.37 1.515 0.237 0.84 0.27 0.0198 -1.77 2.0206 6.25 0.103 -2.34 9.878 0.230 0.0354 0.0187 4.0315 Rn 149900.82 0.29 3.0204 9.35 1.36 1.332 1.804 0.36 1.0091 0.643 0. Ct -4.0172 -0.172 0.0202 1.36 12.0115 0.29 0.289 0.91 0.278 0.0173 3.042 0.565 0.0269 -2.01 -4.217 0.022 0.433 0.0120 0. "' -3.78 o:844 o.0385 151100. Ct "' -0.0193 8.0096 3.0127 7.627 0.19 Rn 59700.88 11.37 1.044 0.30 0.21 0.0084 0.0141 0.17 2.24 0.0221 -0.0122 5.22 2.0272 = 303400.0096 0.0208 11.28 6.20 0.271 0.0344 0. 788 0.082 0.270 0.Chapter 13: Tabulated Data 4.86 -0.0129 0.286 0.0095 2.36 1.0184 0.0296 12.050 -2.35 9.194 0.79 5.220 0.172 0.95 -2.588 0.0216 6.321 1.0335 11.324 1.35 1.064 0.89 -1.0156 -1.0294 0.0148 -3.90 -1.236 0.25 5.0335 0.27 0.93 -1.338 0.086 -0.345 1.91 0.0111 0. q -0. 71 2.0256 0. Ct -3.155 0.0107 3.0211 0.84 0.0275 0.985 7.056 0.0117 4.0105 1.228 0.586 0.87 0.450 0.0209 1.35 10.0162 -0.218 0.516 3.105 0.0252 1.194 0.342 0.0128 0.21 2.0241 10.23 0.85 0.0091 0.85 -0.36 13.0210 0.16 0.178 0.0195 0.310 0.0157 2.95 Rn -3.32 8.36 1.31 1.79 0.0260 0.93 -0.280 0.0147 0.27 3.440 0.0188 0.0103 0.30 0.196 -3.701 0.0125 0.0117 0.0250 1.0268 1. 12.37 = " -4.36 1.0133 6.190 1.200 0. Ct 0.23 3.19 1.27 1.24 0.0220 10.966 0.99 -0.0143 0.35 1.0097 0.002 0.0154 8.0115 0.12 0.0204 0.217 10.0140 0. = = q " 0.246 12.15 1.0167 8.197 0.88 q cd cd cd 6.506 0.0193 1.74 2.39 9.1175 Ct cd -0.132 = "' Rn = 100600. Ct -3.262 0.809 0.0171 0.0103 2.0115 1.489 0.0187 1.241 13.0097 0.90 0.19 0.28 7.0480 0.272 0.683 0.0783 151300.316 0.0153 1.23 3.82 5.0233 10.21 0.28 5.258 0.362 0.82 0.20 0.0143 7.0208 0.36 10.97 0.0109 0.Ql05 0.19 2.38 11.607 0.86 0.37 12. "' -0.275 0.248 0.397 0. = c.963 0.081 0.17 2. = "' -6.817 0.30 9.965 0. = c.0128 0.97 -3.0233 0.97 Rn -2.0162 0.00 -4.0257 0.0177 9.433 0.20 3.532 0.010 0.86 0.98 -3.94 -2.0296 1.184 -3.198 0.0142 0.189 0.30 9.31 13.0278 -0.322 -0.000 0.34 12.134 0.0202 0.406 0.0236 0.16 1.0220 0.20 4.893 0.483 0.776 0.0223 0.993 0.0126 0.033 0.95 -2.0109 0.256 0.98 -3.0436 151000.167 0.0120 0.852 0.21 4.660 0. 927 0.863 0.0111 0.81 8.86 0.94 0.90 0.18 3.31 11.308 0.120 0.25 6.0127 0.012 0.93 -2.0298 0.19 3.0107 6.0203 0. = c.036 0.0133 0.69 2.0095 0.0180 0.38 -0.917 0.0134 5.20 4.0320 0.31 -0.31 Rn c.0182 0.0227 0.0115 0.95 -2.0690 1.98 -3.78 8.271 0.94 -2.0194 0.0186 0.549 0.136 0.973 1.0151 0. = c.86 -0.184 -0.0292 0.0317 0.0178 0.30 9.058 0.1385 201300.97 -3.29 10.0217 0. " -0.0141 0.225 -0.0090 0.370 0.321 -0.545 0.34 cd -0.954 0.0285 0.278 0. 733 0.114 0.0137 0.18 2.01 -4.0139 0. 78 8.473 0.196 -0.103 0.0184 1.0238 1.310 0.0114 0.023 97800.96 -3.0258 0.81 8.164 0.34 9.21 4.0102 0.0120 0.0498 cd -0.93 -1.0267 0.614 0.26 0.751 0. 12.356 0.0138 1.046 0.184 1.996 0.75 cd -0.0150 0.0156 8.94 0.0144 0.0096 0.0269 0.17 3.42 -0.742 0.30 8.174 0.94 -1.21 5.0139 1.0160 0.98 -3.303 0.0233 0.0167 0.68 5.0127 -0.85 0.161 0.0269 0.10 1.36 cd 0.0114 0.84 11.0105 0.70 5.564 0.0121 0.22 Rn 97600.0188 0.90 -0.166 0. = "' 0.0117 0.29 8.910 0.179 0.279 -0.0285 0.124 1.21 3. "' -6.014 0.0386 0.31 10.144 0.127 0.87 0.89 0.0343 0.911 0.0125 0.0088 0.260 0.0154 1. 733 0.0148 0.84 0.0105 0.248 0.0311 0.134 0.386 -5.410 -6.044 0.0108 0.0111 0.19 3.76 8.32 Rn = "' -4.0162 1.576 0.26 6.01 -4.0302 0.924 0.0120 0.632 0.451 0.87 0. 0.0291 0.98 -4.037 0.0309 0.91 -1.0173 -0.17 3.0186 0.0134 0.034 0.29 9.0195 0.22 4. = q "' -0.00 -4.18 2. q -0. "' -0.0226 0.99 Rn -4.0124 0.0264 -0.355 -0.0173 0.0115 0.093 0.24 Rn 58900.15 2.0141 0.28 7.0082 0.364 0.93 Rn -2.412 0.0303 0.30 8. "' -0.993 0.246 0.041 0.467 0.285 0.14 2.890 0.962 0.681 0.144 -0.43 -0.414 0.31 10.21 4.327 0.31 9.0180 1.0112 0.0481 196500.20 3.125 0.69 2.84 -0.0263 1.0400 0.0146 0.175 -0. "' -0.052 0.0118 0. 0.0155 1.27 6.024 0.34 1.671 0.63 2.98 -3.036 0.824 0. 718 0.16 2.609 0.0080 0.0190 0.0186 1.0113 0.24 0.23 5.36 12.0142 6.90 0.13 1.24 5.14 1.0110 0.18 3.0492 300400.251 0.0203 0.022 0.19 4.0131 0.0404 0.22 5.999 0.012 0.183 cd 0.172 0.89 0.30 Rn = "' -5. 773 0.0163 0.0119 0.92 -1.24 6.038 0.098 0.906 0.649 0. = c. 759 0.190 0.71 5.206 -3. 737 0.82 11.25 6.368 0.20 3.0316 1.96 -2.20 6.0114 0.31 -0.0404 1.0770 c.27 7.27 7.27 7.0426 .0191 0.00 -4.0315 0.843 0.25 Rn 59200.360 0.912 0.0085 0. 773 0.853 0.0098 5.99 -2.979 0.0185 -6.0115 0.0140 0. " -0.268 -0.30 10. c.120 0. cd 0.0625 0.0222 1.953 0.0101 0.006 1.0170 0.835 0.0228 11.062 1.364 Airfoils at Low Speeds -2. -0.0185 0.218 = cd cd 0.28 8.0180 0.0128 7.456 0.35 12.85 c.0183 0.30 Rn E374A-PT 0.48 -0.403 0.30 9.29 8.0308 0.0189 0.578 0.325 0.988 cd 0.290 -0.29 8. c.0401 1.0304 1.30 10.24 6.0146 0.217 0.17 2.0302 0.923 1.203 0.29 0.83 11.60 2.0150 0. 12.68 2.85 Rn = "' -3.573 0.350 -0.0103 0.934 0.0507 1.86 11.0220 -0. 766 cd 0.0149 0.009 -3.183 0.190 0.0239 0.0179 0. 764 0.0317 Rn 199200.94 -2.907 0.125 0.400 -6.90 -0.27 0.0257 0.037 4 0.0309 1.0084 0.673 0.24 5.765 0.0115 0.271 0.23 6.0192 1.0339 0.92 -1.0257 0.083 0.30 10.0118 0.0345 151800.0384 0.484 0.17 3.920 0.171 0.313 0.30 11.83 11.037 0.92 -0.72 5.0437 0. cd -0.32 9.025 0.546 0.93 -2.330 0.0271 0.855 0.0100 0.29 8.0140 0.0132 0.626 0.29 9.15 3.0494 298800.13 1.26 6.0275 0.92 -0.25 6.276 0.29 9.31 12.0141 0.888 1.0164 0.0104 0.39 -0. 703 0.0117 0.94 -1.984 0.41 -0.90 -0.92 -0.957 0.0222 0.31 Rn Fig.0207 0.006 97300.612 0.252 0.44 -0.081 0.20 3.0237 0.0196 0. c.016 1.146 -0. = c.29 9.30 11.0231 -0.0122 0.16 1.34 12.28 6.96 -2.0139 0.25 6.0377 0.86 11.0095 0.96 Rn -2.20 cd -0.0216 0.86 1.0092 4.166 0.24 5.261 0.0135 0.22 6.0180 1.0148 0.27 7.65 2.0118 0.0145 -0.17 1.21 3.020 0.0124 0.0170 E374B-PT Fig.055 0. 12.869 0.415 0.23 5.0495 302500.0264 E214C-PT Fig.994 1.274 0.00 -0.463 0.013 0.0351 1.0147 1.096 0.392 0.28 7.93 -2.90 0.0112 0.369 0.0205 -0.469 0.579 cd 0.88 0.220 0.0145 0.0261 0.0140 0.929 0.0261 1.645 0.28 0.060 -3.0113 0.0159 -0.91 -2.774 0.78 8.232 0.18 3.89 -0.0555 150200.0263 0.70 5.0102 0.85 0.0223 -0.119 0.563 0.0366 0.371 0.254 cd 0.0243 0.31 9.243 0.95 -1.75 5.0145 0.0120 0.16 2.0303 0.576 0.0328 0.80 1.243 1.62 2.0110 0.025 0.034 0.457 0.0138 0.27 7. = "' -3.33 9.446 0.37 1.0197 0.14 1. 18 3.0485 0.0300 0.66 2.30 11.408 0.1085 ad 0.252 0.291 0.0106 0.0319 0.31 12.44 -0.0134 0.0136 0.15 1.74 6.061 0. = a1 " -0.98 Rn -4.0198 6.0094 0.831 0.0125 0.94 Rn 0.562 a.708 0.0112 0.0155 0.907 0.23 6.0102 0.0112 0.0118 0.975 0.738 0.23 5.699 0.688 0.198 -3.0124 0.646 0.81 11.038 0.1307 200700.30 12.0351 0.81 11.987 0.1210 200700.920 0.25 7.982 0.0213 0.0668 0.94 -1.557 0.265 0.237 -0.0348 11.0125 0.0793 308600.41 -0.0236 3.21 4.27 7.31 12.645 0.0339 0.0363 0. 970 0. = a1 " -0.980 0.30 12.41 0. Ct -3. 948 0.0200 0.539 0.062 0.72 6.0113 0.27 7.25 6.0089 0.26 7.0113 0.94 -2.0148 0.97 -0.0147 0.0540 E374B-PT = Fig.94 -2.30 9.0124 0.975 0.79 = a1 "' -0.0110 0.689 0.28 9.802 0.975 0.0164 0.77 9.80 11.069 -2.0122 0.162 0.29 8.25 6.932 0.873 0.270 0.95 0.0097 0.991 0.16 3.815 0.888 0.0142 0.26 6.949 0.0115 0.013 300000.93 Rn -1.43 0.29 10.0097 0.0878 0.519 0.813 0.0154 0.0564 1.820 0. 12. a1 -2.93 Rn ad -0.1964 E374C-PT Fig.0684 201600.058 0.089 0.284 0.349 0.0129 0.124 0.89 0.0108 0.0108 0.20 4.0085 0.30 10.019 0.273 0.92 -0. a1 ad " -0.690 0.14 1.0149 0.0248 0.89 0. 12.29 9.29 0.0153 0.0371 0.0317 E374B-PT Fig.0109 0.0106 0.409 0.069 -2.562 0.0162 0.20 4.453 0.0233 5.142 0.30 10.0133 0. a1 -1.69 5.0149 0.045 0.0285 0.30 9.699 0.0443 0.844 0.22 4.251 0.916 0.18 0.847 0.919 0.75 6.0202 0.0153 0.23 5.984 0. 97 4 0.0123 0.0143 5.079 -2.0118 0.26 Rn 100800.742 0.12 14.981 0.88 0.0119 0.21 4.961 0.71 0.19 3.0139 0.80 14.045 0.32 13. 904 0.0112 0.015 0.27 7.96 -3.0297 0.20 4.Chapter 13: Tabulated Data 11.90 0.26 7.691 0.81 11.79 11.0180 8.72 6.0305 0.411 0.357 0.41 0.0246 0.069 O.168 -3.541 0.0233 9.0155 Rn -1.0227 0.0125 -0.16 3.271 0.953 0.0099 0.66 3.0106 -1.997 0.76 0.17 2.0100 0.16 1.0100 0.26 7.22 4.31 0.0195 0.0144 0.27 8.958 0.822 0.0090 0.0161 = "' = " ad ad 8. Ct -3.0227 0.0327 0.29 10.090 0.0119 0.78 9.0564 0. = C1 " -0.14 1.14 1.0123 0.886 0.25 6.268 0.76 8.043 0.0211 0.96 -0.744 0.64 2.33 = al " -0. 964 0.0102 0.0087 0.87 0. 726 0.0434 151300.71 0.96 Rn -2.0609 ad 0.0435 202400.897 0.813 0.118 0.29 9.893 0.065 0.136 0.0085 0.31 0.0105 0.218 0.0144 0.29 Rn 147300.22 5.394 0.17 0.0148 0.41 0.0506 0.30 0.0196 -2.77 0.0123 -0.220 0.0130 2.829 0.957 0.996 1.31 12.421 0.81 11.67 3.334 -5.0641 ad 0.970 0.366 0.0106 0.0101 0.80 0.535 0.0079 0.0102 0.19 4.0095 0.67 3.0140 3.65 0.89 0.37 0.0231 0.0148 -0.358 0.29 10.149 -3.79 0.222 0.0092 0.088 -2.995 0.703 0.94 -1.0152 0.418 0.0260 0.928 ad 0.914 0.229 0.0110 0.0115 0.651 0.995 " -2.0128 0.80 11.813 0.64 2.005 0.75 ad 0.27 7.42 -0.68 3.27 Rn 103600.409 0.0105 1.0097 0.0120 0.70 5. 12.80 12.0245 0.0466 0.0193 0.29 9.094 0.0101 0.0112 0.19 3.064 0.0121 0.0097 0.009 1.813 0.66 3.0125 -0.914 0.0143 -0.031 0. E374B-PT = = 153400.28 7.978 0.0509 0.0205 0.996 1.272 0.78 9.975 0.0405 0.120 0.70 5.157 -0.23 0.970 Rn = " -2.947 0.0131 0.72 6.287 0. 12.0093 0.91 -0.0118 0.24 5.29 9.23 6.259 0. 101 0.65 2.0091 0.0124 0.0117 0.00 Rn -4.77 9.822 0.96 Rn -2.30 10.0160 0. 938 0.531 0.0171 0.30 12.417 0.547 0.91 0.0091 0.1619 0.81 -0.o126 0.90 -0.14 1. = a1 " -0. 983 0.0224 0.438 0.0324 0.649 0.14 1.28 Rn 100600.90 0.188 -3.0106 0.78 8. = a1 " -0.100 0.929 0.0118 0.551 0.0088 0.29 10.0264 0.0241 0.896 0.984 1.93 -1.0105 0.1213 365 .0101 0.0370 0.0246 0.919 0.675 0.405 0.31 12.555 0.81 0.0539 Rn 196300.071 0.12 1.0136 0.0124 0.0338 0.0143 0.399 0.0204 -0.079 0.80 11.550 0.39 0.0089 0.77 9.93 -1.2177 "' -2.15 0.64 0.557 0.81 -0.0102 0.0112 0.19 3.969 0.0206 2. 952 0.17 2.28 8.43 -0.0122 o.16 2.67 3.72 5.0492 0.30 11.972 0.20 4.902 0.0440 0.455 0.394 0.881 0.193 0. ad 0.29 ad -0.72 6. -0.088 0.31 ad 0.0496 0.0196 0.0141 0.22 0.94 ad 0.681 0.048 0.981 0.89 0.80 0.0997 0.26 6.0115 0.19 4.0156 0.093 0.16 2.953 0.0094 0.29 Fig.548 0.0121 0.822 0.008 0.0153 0.30 9.403 0.0607 305000.79 9.823 0.0127 0.629 0.26 7.88 0.0123 0.0539 0.0538 Rn 149400.82 -0.0331 0.992 304200.272 0.549 0.003 ad 0.139 0.181 -3.867 0.0098 0.0107 0.0230 9.80 a1 = a1 "' -0.983 0.047 0.899 0.0520 0.0161 -2.0354 11.80 11.93 -1.0142 6.0153 -0.D18 0.276 0.28 8.365 -6.97 -3.975 0.0134 0.30 10.973 0. = a1 " -0.0133 0.814 0.23 ad 0.0141 0.979 0.065 0.94 Rn -1.42 -0.29 10.78 8.81 ad -0.24 6.17 3. -0.456 0.0193 0.0150 0.980 1.76 8.0121 0.42 -0.0248 0.18 3.094 0.94 -2.29 0. = a1 " -0.0093 0.0180 0.146 0.96 -2.14 1.923 0.91 -0.0176 0.557 0. 558 0.0173 0.152 -3.18 2.0288 0.0195 0.30 9.0268 0.094 -3.43 -0.120 0.0240 0.89 Rn cd 0.0434 cd 0.519 0.0102 0.18 3.87 -0.169 0.95 -2.0114 0.20 3.030 0.082 0.0153 0.34 11.284 0.096 -3.0349 311800.0176 0.91 -1.22 4.33 10.0089 0.109 -2.0345 149600.202 0.34 -0.94 -2.695 0.95 -2.396 0.33 10.23 4.0194 1.742 0.396 0.0104 0.394 0.0130 0.89 0.443 0.32 12.805 0.914 0.32 9.147 0.893 0.0252 1.398 cd 0.318 0.16 1.19 3.0145 0.0146 0.058 0.32 9. 782 0.0114 0.88 0.134 0.0193 0.188 0.110 -3.26 6.150 1.0324 0.0262 0.34 11.0265 1.24 5.34 11.31 8.33 10.0245 0.0146 0.0158 0.0217 0.0266 1.602 0.131 cd 0.053 1.496 0.91 -0.057 0.23 4.623 0.103 1.15 1.328 0.17 2.73 0.33 10.521 0.92 -1.025 0. Cz "' -0. cd Cz = -0.21 3.30 8.397 0.0091 0.32 Rn = "' Cz cd 0. Cz cd -2.14 1.23 4.120 1.89 -0.420 0.133 1.0100 0.070 1.8.26 6.0368 0.123 cd 0.601 0.0109 0.0128 0.29 Cz -0.24 5.18 2.90 -1.0101 0.055 0.16 1.0146 0.292 0.91 0.29 7.0143 0.125 0.0120 0.20 3.887 0.005 1.0210 0.106 0.95 -2.0325 0.34 10.31 Rn 59700.16 1.0153 0.510 0.212 -0.980 1.93 -1.22 4. 722 0.89 0.100 cd E387A-PT Fig.0204 0.0172 0.599 0.27 7.692 0.0256 1.0233 0.106 0.0382 0.34 11.27 7.109 1.0367 1.81 11.0204 1.35 Cz -0.415 0.89 -0.28 7. 720 0.16 1. = Cz "' 0.221 0.894 0.20 3.27 7.0217 0.0210 0.0187 Rn cd 0.093 1.25 6.0105 0. = "' -3.0324 0. 12.0134 0.0100 0.0135 0.88 -0.0133 0.050 1.26 6.0136 0.19 3.91 0.0151 0.29 9. = Cz "' -0.28 7.0260 1.34 Rn 101000.0179 0.27 6.31 8.31 9. = "' -2.523 0.0157 1.0249 202000.0135 0.0196 1.282 0.909 cd 0.0094 0.110 -0.88 0.0196 1.0111 0.0159 0.20 3.182 0.0141 0.0121 0.913 1.35 10.88 -0. = Cz "' -0.0120 0.0173 0.098 0.0231 0.68 3.0098 0.089 0.0164 0.0108 0. 12.0154 1.322 0.0296 0.0362 l.0086 0.0123 0.20 3.932 0.060 1.0137 0.23 4.302 0.683 0.891 0.005 0. Cz 0.27 6.22 4.33 11.0093 0.0138 0.195 0.77 8.0111 0.0116 0.17 2.896 0.28 7.0146 0.102 0.0128 0.0120 0.0115 0.0204 0.018 0.188 0.059 0.288 0.0147 0.87 0.18 2.32 Rn = "' -2.905 0.211 -0.90 -1.0280 0.0495 0.0106 0.0218 0.24 6.0092 0.973 1.001 0.0128 O.0112 0.133 0.17 2.16 2.95 -1.34 11.32 10.0484 200400.120 302200.75 6.0479 149600. 708 0.31 8.34 Rn = "' -3.0322 1.133 -1.24 5.0118 0.0119 0.0175 0.0122 0.103 0.0130 0.0204 1.139 0.876 0.220 0.183 0. Cz cd "' -0.0120 0.34 11.059 0.003 0.87 0.33 9.0088 0.0080 0.711 0.0139 0.31 8.38 0.0119 0.0085 0.0143 0.34 10.0093 0.0176 0.0246 1.33 Rn 101500.0082 0.19 3.29 8.605 0.0221 0.0101 0.25 5.0150 0.0103 0.32 9.71 5.24 5.605 0.131 0.32 9.0097 0.0088 0.881 0.092 0.973 0.122 0. = "' -3.004 0.0168 1.0203 1.005 0.0273 E387A-PT Fig.297 0.34 -0.783 0.0187 0.102 1.0152 -1.298 0.030 0.042 1.28 7.390 0.015 0.0158 0.29 7.0288 0.0117 0. "' -3.0155 0.16 cd 0.88 0.0165 0.22 4.0322 0.100 0.986 0.93 -1.18 2.63 2.905 0.22 4.95 -2.0309 E387A-PT Fig.984 0.21 4.20 3.29 7.30 8.493 0.24 5.0113 0.044 1.0678 Rn Cz "' -0.95 -2.0183 0.0096 0.67 3.0224 0.985 1.0175 0.31 9.018 0.322 0.14 1.0092 0.14 1.413 0.0188 0.0378 1.33 9.0095 0.0199 0.097 -3.705 0.39 0.994 0.0097 0.13 1.0177 -0.0174 0.31 8.94 -2.0119 0.121 0.33 10.617 0.893 0.0102 0.501 0.216 0.25 6.059 0.0322 0.070 0.87 0.27 7.0586 = 199800.0181 1.126 0.291 0. 702 0.33 9.91 -0.90 -0.29 7.389 0.0102 0.94 -1.18 2.006 0.18 2. 780 0.810 0. 794 0.0154 0.0166 0.89 0.16 1.32 9.624 0.979 0.092 0.103 0.24 5.111 0.0118 0.0138 0.0191 1.699 0.0535 153000.823 0.0162 0.800 0.0125 0.D105 0.21 4. 12.76 6.0268 0.595 0.366 Airfoils at Low Speeds Rn = 300300.800 0.94 -2.30 8.32 Cz -0.17 3.0186 1.25 5.31 8.092 1.500 0.94 -0.llO 0.0083 0.0093 0.90 0.29 8.0113 0.241 0.0125 0.97 -1.34 Rn = 0.0107 0.812 0.0180 .0104 0.487 0.0107 0.988 0.506 0.599 0.601 0.33 11.0258 0.23 5. -3.599 0.058 0.095 0.285 0.0153 0.0185 0.21 4.95 0.115 0.89 -0.0138 0.0224 0.16 1.26 6.115 0.35 0.0120 0.0093 0.0140 0.0449 cd -0.26 6.0306 0.501 0.127 0.91 -1.21 4.24 5.20 3.33 Rn = -0.110 0.33 10.31 8.183 0.23 4.114 1.114 0.0107 0.29 7.0130 0.977 0.95 -1.979 1.0136 0.32 9.34 0.397 0.0085 0.15 1.096 0.690 0.0202 0.129 1.31 9.104 0.27 6.0145 1.0132 0.081 0.35 0.30 8.26 6.0131 0.15 1.064 0.27 6.842 0.595 0.0123 2.18 2.32 10.982 0.0475 303100. 710 0.749 0.445 0.589 0.000 0.417 0.126 0.33 Rn = "' -3.0286 -0. 12. 702 0.92 -2. Cz 0.0141 0.0116 0.88 0.0196 1.0134 0.91 Rn -2.0107 0.999 cd 0.0223 0.0274 -0.286 0.0483 E387A-PT Fig. Cz "' -0.22 5. 797 0.142 99800.32 0.503 0.0114 0.32 Rn 99500.0303 0.0091 0.991 1.88 -0.90 -1.0154 0. 0285 0.18 3.28 6.178 -0.31 9.17 2.32 c.0087 0.19 3.30 8.0225 10.25 5.0316 0.610 0.23 4.280 0.0187 0.0207 1.0131 0.134 0.0142 0.939 1.685 0.186 0.0149 0.29 7.982 0.80 -0.0163 0.0199 1.674 0.0148 0.189 0.324 0.0246 1.0921 199700.30 8.21 6.100 0.32 9.32 8.0107 0. = " -2.0171 0.24 1.34 9.34 11.175 0.0213 -0.19 5.0119 0.30 8.0365 0. " -6.0101 0.054 0.502 0.0118 0.0187 1.111 0. 73 5.515 0.0185 1.0388 1.621 0.014 0.17 2.412 0.031 cd 0.200 0.457 0.001 0.0093 0.33 14. cd 0.0348 0.Chapter 13: Tabulated Data 7.0091 0.184 0. 12.12 2.0167 0. 796 0.0195 1.567 0.25 6.128 0.0158 0.0147 1.614 0.85 0.33 10.26 6. 702 0.0181 0.24 Rn = " c.893 0.92 0.211 0.040 0.42 -0.34 11.89 -0.0397 1.92 -2.90 0.349 0.076 0.086 cd 0.0123 0.015 0. cd -0.0178 0.91 -1.20 6.13 1.85 0.0182 0.19 3.197 0.24 5.282 0.0143 0.0207 0.31 9.0160 0.132 0.91 -1.92 -1.952 0.0290 1.299 0.286 0.622 0.99 -2.0122 O.169 0.32 c.33 10.0301 cd -0.0131 0.635 0.31 9. c.17 2. cd -0.0149 0.89 0.16 4.65 2.0505 1.15 3.0106 0.0442 1.97 -1.16 1.469 0. 0.699 0.0135 0.0168 0. 12.274 0.0112 0.24 5.0106 0.902 0.0147 0.88 0.185 0.829 0.4500 201300.23 5.009 0.30 8.33 11. -0.606 0.0166 0.97 -1.28 7.89 0.759 0.94 -2.0201 1.27 6.604 0. = " 0.0144 0.90 -0.0426 Rn = 301300.28 7.0723 0.0525 1.086 0.0104 0.404 0.012 0.0335 0.028 1.0304 1.0163 0.93 -0.17 2.0491 0.27 2.35 14.0143 0.340 0.96 -1.00 -3. 12. cd -0.28 7.0110 0.405 0.23 4.207 0.0204 -0.15 1.0195 1.0127 0. 721 0.520 0.180 0.667 0.0108 0.169 0.0705 300400.25 c1 0.0420 1.098 1.0105 0.0171 1.337 0.006 1.34 1. 766 0.97 -1. 722 0.417 0.662 0.0158 1.88 -3.25 5.0180 0.100 1.91 -0.285 0.18 1.036 0.168 0.21 4.601 0.441 0.0164 0.0256 1.282 0.0115 0.0250 0.065 0.635 0.31 8.26 6.889 1.116 0. cd -0.897 0.35 11.120 0.0170 0.34 14.0133 0.110 0.27 7. 724 0.0114 0.105 1.22 7.602 0.160 0. c1 cd -3.101 0.055 1.886 0. E387B-PT "' Fig.0263 0.079 0.0116 0.186 0.0223 1.19 5.0228 0.19 3.23 5.606 0.98 -0.18 3.31 9.91 -0.0416 0.560 0.104 0.0144 0.134 0.0129 0.0126 0.500 0.769 0.175 0.83 -1.401 0.0402 1.39 Rn 59600.86 0.24 6.88 -0. = " -3.92 -1.0216 0.32 9.34 12.91 -2.0682 0.0093 0.32 c1 -0.0243 1.0143 0. -0.522 0.96 -2.0116 0.0445 cd -0.0121 0. 12.0114 0.0090 0.109 0.082 0.35 13.0133 0.0153 0.407 0.0544 0.0159 0.0393 298800.379 0.100 0.0335 -0.0165 1.35 12.0246 0.27 7.127 0.0160 0.89 -4.0513 Rn = 203400.0233 1.22 5.0101 0. " -5.93 -1.0137 0. c.0080 0.0139 0.0083 0.14 3.512 0.069 c.94 -0.0162 0. cd -3.20 2.310 0.25 5.99 -2.36 Rn c.33 12.057 0.22 3.0112 0.0238 -0.0112 0.078 0.0181 -0.0166 0.13 2.135 0.0143 0.0149 0.0207 1.800 0.20 6.0183 -1.236 0.40 Rn = 200100.0128 0.18 2.0142 0.0103 0. " -3.26 6.0148 FX63-l37 A-PT Fig.28 cd -0.31 9.30 8.24 4.0402 152300.33 c.100 cd 0.0098 0.0175 0.0159 0.34 Rn = " -4.21 3.0203 -2.34 11.0188 0.052 0.0104 0.0197 0.029 0.0550 Flat Plate-PT Fig. 706 0.16 1.33 13.0287 -0.10 1.34 10.0153 0.0303 1.0123 0.34 12.34 13.98 -2.307 0.95 -0.211 0.0120 0.0143 0.37 Rn = 100100.088 1.16 4.0166 0.0165 1.931 1.050 0.983 0.0121 0.430 0.0161 0.0096 0.275 -0.089 0.0222 1.052 0.21 4.21 4.087 0.0093 0.508 0.0172 0.90 0. -3.0290 -2.0103 0.21 3.296 0.21 3.34 12.852 0.33 10.158 0.048 0.0248 1.0135 0.461 0.403 0.219 0.142 0.0227 0.224 0.131 0.0104 0.541 0.017 0.26 6.0194 0.10 1.190 0.007 1.0600 Rn = 153700.0170 0.295 0.517 0.115 1.506 0.812 0.0268 0.0110 0.35 Rn 1.021 0.959 1.Dl05 0.086 1.22 4.216 0. 713 0.30 7.374 -0. c.246 0.310 0.016 0. " -3.0360 0.0149 0.0494 0.86 0.0201 0. = " -3.0133 0.084 0.32 9.16 1.88 -0.32 " -3.95 -0.077 0.24 4.17 1.415 0.0116 0.0114 0.859 0.0342 0.148 -0.0113 0.23 5.32 10.117 0.206 0.0193 -0.0209 0.35 0.26 6.379 0.804 0.888 0. -0.120 0.0232 0.0487 0.90 -1.0243 0.0425 0.982 cd 0.98 -3.0105 0. 793 0.0107 0.0156 0.177 0.34 11.0358 1.062 0.881 0.31 8.33 10.24 7.18 2.292 0.383 0.34 11.082 0.0307 202600.28 8.0263 1.83 11.18 5.0152 0.0798 -3.0138 0.17 1.0314 -0.33 10.003 0.90 -0.36 Rn = 97000.0177 0.32 11.0189 0.36 Rn = " c.10 1.151 99400.0136 0.79 0.92 -5.0120 0.0097 0.21 3. c.008 0.0095 0.0232 0.87 0.0299 1.107 0.067 0.0231 -0.989 0.16 2.192 0.0103 0.32 9.19 2.0691 0.94 -2.85 -2.92 -1.0335 1.814 0.26 6.0121 0.973 0.0959 FX60-l00 Fig.609 0.0163 367 .91 0.16 4.23 Rn -0.116 1.91 -2.27 7.86 -0.144 0.29 7.33 Rn = " -3.134 0.0309 1.0150 -0.900 0. 701 0.0134 0.507 0.90 -0.138 0.093 0.95 -0.0185 0.95 -2.111 0.085 0.21 4.186 0.225 0.93 0.12 2.30 3.16 1.123 0.109 0. 795 0.0148 0.041 0.15 3.969 0.399 0. 78 8. 31 10.615 0.0095 0.42 1.30 1.210 0.0183 0.551 0.754 0.177 -3.14 1. 0.196 0.0437 HQ2/9A-PT Fig.0148 -0.420 0.10 1.0084 0.0084 0.42 Rn 62000.553 0.24 1.0184 0.385 0.0150 0.556 0.0271 0.156 1. = c.008 0.642 0.0123 0.30 0.992 0.20 3.0275 11.0128 0.0295 FX63-137B·PT Fig.049 -6.0101 0.25 6.31 11.93 -5.652 0.42 10.30 10.27 7.0190 0.597 0.0130 0.28 8.0323 1.0136 0.597 0.045 0.88 0.0146 0.79 0.981 1.0122 0.539 1.0125 0.38 7.893 0.31 9.0371 1.26 2.0238 0.96 -2.28 8.0202 7.19 3.0265 0.85 -1.125 1.338 0.0208 0.0368 Rn = 299500.36 6. 778 0.0109 0. 979 0.0145 0.95 -2.079 0.412 0.0125 0.0144 0.833 0.164 0.589 1.613 0. 0.707 0.339 0.36 1.39 8.90 -4.18 3.0247 1.14 1.94 -5. "' -0.26 7.95 -1.0243 0.79 -0.0204 -0.21 4.0122 0.0281 0.277 0.34 6.0279 0.0203 0.366 0.17 3.0088 0.0276 0.21 4.0117 0.121 0.0130 0.0113 0.0100 0.520 0.0248 0.0094 0.074 0.79 0.43 12.135 0.90 0.0098 0.0182 0.41 9. "' -0.93 -0.44 11.0967 303400.27 3.0155 -0.955 0.39 8.684 0. "' -3.946 0.0364 1.28 2.439 0.080 0.321 0.291 0.0193 6.0162 0.379 1.14 1.28 2.0233 0.0329 0.0219 1.44 1.86 -2.25 6.0150 0.15 2.525 0.084 0.32 1.462 0.79 0.964 1.273 0.186 0.015 0. c.44 c1 c.826 0.457 0.43 0.898 0.0249 10. c.815 0.247 1.0088 0.887 0.92 0.815 0.0443 c.0155 0.017 0.28 3.22 1.0287 0.0075 0.44 c.623 0.94 -1. 0.0212 0.0101 0.024 1.029 0.79 0.0134 0.86 ·2.220 1.0260 0. 79 -0.0072 0.004 0.0108 0.0185 1.33 11.459 0.94 -0.379 0.190 0.607 0.0083 0.738 0.436 0.544 0.43 Rn = 98200.0150 0. -0.0170 -0.373 0. 735 0.31 c1 = c1 "' -0.0225 0.0124 0.40 9.649 0.0386 1.0131 0.251 0.0116 0.0308 0.82 -0.959 0.33 5.83 -1.19 4.36 6. 0.26 2.31 9.93 -0.936 Rn = 98300.0092 0.0313 0.0072 0.222 0.0320 0.0262 1.0175 2.974 0.0254 1. "' -6.0500 0.0395 Rn = 208700. 965 0.0281 1.0128 0.530 0.083 0.042 0.0240 0. 0.571 c. 733 0.27 7.022 0.129 0.411 1.0241 1.092 0.43 11.14 1.94 -1.0519 Rn = 308600.250 0.23 5.647 0. "' -6.94 -1.592 c.90 ·4.862 0.0133 0.0129 0.42 9.31 4.0274 1.32 -0.0141 0.886 0.0144 0.0125 0.821 0.0143 0.609 0.0282 0.31 0.85 -2.0221 1. -0.029 0.227 0.347 0.300 1.96 -2.038 1.93 -1.31 11.30 4.0183 1.44 12.0122 .94 -1. c1 " -0.0359 1.379 0. "' -3.0168 0.0156 0.0175 0.0307 0.452 1.22 5.0163 -0.171 -3.22 5.89 ·4.031 0.0166 -0. 12.0075 0.0522 Rn = 155600.33 5.0164 1.399 0.1163 Rn = 149000.81 -0.106 0.88 -3.0116 0.0125 0.40 9.30 c1 c.25 6.38 1.20 4.25 6.0214 0.061 0.191 0.0139 0.0253 1.0298 1.154 0.599 0.0104 0.84 -2. " -3.0175 1.219 -3.0259 1.40 1.301 0.460 0.83 -1.32 10.560 0.333 1. 734 0.0134 0.331 0.046 0.464 0.16 2.0118 0.83 -1.890 0.30 9.25 1. 0.062 1.226 0.587 0.90 -0.29 c1 -0.42 1.051 0.637 c.0476 c.0210 0.878 0.0332 202100.43 1.320 0.0249 0.13 -0.532 0.207 c. c.236 0.742 0.0219 8.0198 1.15 1.39 8.945 0.0155 0.0144 0.96 Rn c.290 0.0263 1. c1 " -0.0260 0.77 0.27 8.0154 0.25 1.0119 0.30 9.0165 0.20 4.049 0.37 7.23 3.0272 1.210 0.0110 0.0164 1.0326 Rn = 200700.459 0.42 10.123 0.91 0.0131 0.0157 0.538 0.056 0.0291 1.29 9.24 2.229 0.88 0.0353 1.616 0.452 0.35 1.19 3.001 0.076 1.934 0.39 8.88 -3.42 11.0089 0.174 0.0140 0.0106 0.0166 1.626 0.088 0.0124 0.348 0.17 2.0162 0.35 6.0109 0.33 5.512 1.0162 0.32 5.36 7.564 0.23 5.558 0.27 7.19 3.250 0.819 0. c.41 Rn = 100800.89 -3.213 -3.140 0.15 2.0165 0.95 -1.0160 0.0153 0.0184 HQ2/9A·PT Fig.93 -0.92 -0.21 4.056 0.26 7.838 0.0249 1.20 4.0183 0.046 0.573 0.91 -0.23 5.0088 0.43 10.0122 0.010 0.189 -2.826 0.907 0.29 8.0199 1.499 0.575 1.0136 0.0091 0.037 0.0171 0.26 6.119 -0.92 -4.473 0.24 5.97 -2.0185 0.19 4.26 7.0235 9.26 7.368 Airfoils at Low Speeds 4.44 12.842 0.34 6.527 0.538 0.0144 0. = c.870 0.008 0.31 -0. 0.581 0.0139 0.769 0.0086 0.30 3.89 -3. 0. 0.0324 0.0192 0.646 0.23 5.982 0.10 1.18 2.88 0.0154 0.0186 0.44 13.44 c1 -0.21 4.0166 0.0172 0.0137 0.935 1. c1 " -0.0113 0.0121 0.94 -1.0139 0.0143 0.0463 0.390 0.155 0.0148 0.0160 0.1139 Rn = 151600.15 2.0134 0.304 0.0163 1.19 3.0205 0.20 -0.486 0. " -3.38 8.42 9. 736 0.165 0.657 0.0144 0.37 7.31 10. " -5.96 Rn -2.0120 0. 12.0186 4.30 4.443 0.912 3.542 0.0504 1.0172 -0.30 8.28 8.0148 0.0140 0.0220 0. 12.24 6.43 11.139 0.88 0.028 0.0147 0.647 0.23 1.080 0.30 10.92 -0.30 9.532 0.037 0.0107 0.0215 1.635 0.0092 0.345 0.43 11.811 0.32 c.0161 1.0158 1. 0.0183 5.0106 0.36 7.491 1.24 6.42 10.046 0.91 0.0137 -2.28 5.742 0.604 0.79 -0.91 -5.0188 0.32 11.552 1.85 ·1.100 0.96 ·2.209 0.0149 0.17 3.089 0.0196 0.299 0.18 2.122 0.407 0.0532 1.28 9.0123 0.675 0.0152 0.41 10.0158 0. 737 0.0143 0.158 0.0234 1.87 -2.28 8.0205 -0.13 1.0107 0.227 0. -0.0299 0.149 0.0179 0.014 0.18 3.0147 0.0109 0.26 7.004 0.68 3.0201 0.0093 0.16 2.0123 0.95 0.0134 0.96 -2.908 0.0422 0.0107 0.334 0.0132 0.0150 -0.0242 9.45 Rn = 151600.0132 0.933 0.26 7.0293 1.91 -0.158 0. 0096 0.05 Ca -0.Oll6 0.31 1.46 -0. -0.374 0.073 1.65 3.040 0.29 8.31 10.031 0.917 0.0093 0. c1 cd -0.185 -4.25 7.0115 0.119 0.500 0.0196 0.0161 0.0091 0.29 9.0133 0.76 8.74 6.965 0.0134 0.42 0.658 0.0176 0.20 4.295 0.97 -2.30 13.027 0.0092 0.1496 HQ2/9A-PT Fig.952 0.0163 0.756 0.30 9.04 7 0.0236 0.22 5.78 Rn = a -2.74 6.135 0.801 0.250 0.27 8.10 1.43 0.32 12.739 0.0508 0.0249 0.042 0.0287 10.0128 0.975 1.03 1.080 0.31 1.172 0.116 -3.344 0.0239 0.46 -1.331 0.0156 0.31 9.0425 1. 739 0.32 11.1783 152000.397 0.15 2.0086 0. c1 "' -0.0085 0.0177 -0.31 0.825 0.461 0. -0.01 -0.43 0.0256 0.0151 0.38 1.0095 0.747 0.0242 0.970 0.0130 0.25 6.0145 0.62 2.0173 0.332 0.049 0.1588 204700.24 7.863 0.287 0.0736 0.694 0.95 0.549 0.30 10.978 0.93 -1.90 0.858 0.47 Rn = 60800.43 1.051 0.902 0.069 1.32 11.025 0.24 6.014 0.0105 0.0129 0.1970 149300.06 -0.0097 0.460 0.39 1.680 0.31 Rn = a -3.15 2.32 Rn = c1 c1 cd -0.0903 Rn = 305200.008 0.0135 0.0155 0.037 0.77 9.1433 0.0386 0.69 5.0372 308100.04 7 1.19 4.22 4.023 0.0222 0.75 8.48 Rn = 58900.90 0.020 0.94 -1.022 0.0382 11.30 13.0162 0.0151 0.26 7.492 0.29 Ca 0.999 0. 950 cd 0.41 0.051 0.03ll 1.0111 0.12 1.0252 0.0148 0.929 0.37 1.32 11.362 0.15 2.22 6.0677 0.670 0.254 O.73 6.30 9.0134 0.218 -3.0102 0.012 0.0138 0.66 3. = a cd -0.30 Rn = "' -3.550 0.991 1.499 0.180 0.25 6.471 0.0116 0.080 0.432 0.0324 0.966 0. c1 Cz cd c1 cd -0.77 8.13 1.0145 0.0174 8.0201 0.003 0.953 0.68 5.663 0.0092 0.603 0.148 0.018 0. 12.744 0. a c1 -4.41 0.46 -1.19 4.722 0.78 9.23 6.20 4.0074 0.026 0.980 0.829 0. c1 " -0.70 5.382 0.0154 0.064 0.0084 0.0136 0.858 cd 0.49 -0.0124 0.11 1.0249 0.0188 0.900 0.0076 0.116 0.0128 0.94 -1.30 10.020 0.0108 0.842 0.31 10. 740 0.014 0.146 0.82 12.056 0.0102 0.981 1.20 4. c1 c.0093 0.959 0.42 0.0116 0.106 0.143 -2.0378 1.962 0.26 7.Oll2 0.129 0.140 0.0581 0.10 1.070 Ca 0.0160 0.361 0.935 0.0404 1.24 7.29 8.44 Rn = 101600.894 0.98 -2.10 1.0102 0.0138 0.0098 0.31 Rn = "' -3.739 0. c1 -0.814 0.31 12.0218 0.94 -1.906 0.79 9.210 0.0099 0.35 7 cd 0.823 0.960 0.037 1.0130 0.0078 2.864 0.30 10.980 0.0129 0.0707 0.13 1.0103 0.0370 1.31 10.545 0.0104 0.162 0.95 0.0091 0.0107 0.63 2.31 10.92 -0.968 1.96 -0.28 10.67 3.073 1.979 0.215 0.070 0.77 Rn = a -2.0200 0.0136 0.90 -0.0260 1.0116 0.0096 0.0106 0.97 Rn c.21 4.0432 0.0091 0.0074 0.79 12.25 6.0147 0.30 Rn = 0.0195 0.39 0.163 -3.881 0.13 1.0081 0.30 10.0127 0.24 6.79 1.564 0.24 6.744 0.0307 0.226 0.29 10.825 0.32 1.27 7.644 0.027 0.97 -1.0184 0.01 -2. 733 0.043 0.81 12.0102 0.0112 0.1425 0.96 -2.973 0.201 -0.69 5.33 0.0090 0.006 0.78 8.452 0.0189 -0.224 0.15 3.0164 0.604 0.1444 0.866 0.81 Rn = a -2. 12.21 5.0105 0.78 9.74 6. c1 " -0.0214 0.29 9.0088 0.902 0.0174 -0.0166 0.15 2.849 0.0317 1.802 0.0116 0.108 0.28 9.01 -2.0429 HQ2/9A-PT Fig.0546 0.20 5.29 9.0751 100800. 734 0.97 -0.0153 0.0119 0.0228 0.0112 0.013 0.217 0.558 0.29 0.0240 0.323 0.017 Ca 0.0098 0.25 7.182 0.30 10.26 7.0198 0.77 8.185 -3. a -3.0117 0.0270 0.0283 1.80 Rn = c1 c1 " -0.142 0.91 -0.31 Rn -0.76 -2.837 0.19 5.13 1.0157 -0.0157 0.73 6.0081 0.98 -1.0320 HQ2/9B-PT Fig.15 3.2090 202400.18 3.974 0.Chapter 13: Tabulated Data 5.94 -1.70 5.0124 -1.28 8.0176 0.329 0.0097 0.0085 0.0116 0.66 3.80 12.46 -1.0201 -2.888 0.690 0.806 0.0214 0. 12.0136 0.855 0.559 0.0077 0.0208 0.81 11.0098 0.0088 0.75 6. -2.824 0. 0.27 7.0783 311700.16 2.0917 200000.0086 0.13 3.0350 Rn = 303500.0078 369 .30 10.0434 1.45 -1.654 0.77 Ca 0.0282 0.928 0.781 0.13 0.054 0.96 -0.683 0.221 0.23 6.30 13.0132 0.035 0.0168 0.30 9.179 0.0149 0.122 0. -0.17 4.1019 205300.231 0. 12.149 0.91 0.0075 0.132 0.0314 1.262 0.77 9.17 4.0158 0.0541 153300. 01 -2.054 0.24 6.0116 0.15 3.151 0.025 0.049 0.1481 HQ2/9B-PT Fig.31 10.0155 0.979 0.0329 0.0107 0.30 10.391 0.698 0.539 0.19 4.084 0.068 0.0125 0.24 6. c1 " -0.0148 0.576 0. = c1 "' -0.43 -0.0197 -0.07 -3.038 1.071 0.496 0.904 0.538 O.033 1.0241 1.0093 0.15 2.0191 0.71 5.913 0.97 -1.053 1.0176 0.011 c. c1 Ca "' -0.0221 1.77 9.0121 0.445 0.30 9.66 3.31 11.0164 0.0362 1.840 0.0751 100100.0124 0.0150 0.393 0.0136 0.0228 9.27 8.0231 1.561 0.018 0.32 11.73 -0. 0637 99500.0174 0.455 -0. = "' -7.04 71 12.587 0.03i2 200100.0160 0.670 0.0107 0.15 3.002 0.224 -0.0187 -0.0225 0.25 7.548 0.29 13.0242 1.1908 0.036 0.0175 0.683 0.58 -0.683 0.17 3.004 0.0191 0.92 0.0283 0.072 0.011 0.27 7.0160 -0.033 cd 0.27 9.0147 0.354 0.449 0.0086 0.0131 0.19 0.01 -5.339 0.42 0.0111 0.0094 0.0103 0.791 0.76 c1 c1 -1.81 12.052 1.0163 -0.0120 -0.23 8.0285 0.0251 0.02 -6.02 -3.227 0.81 12.004 0.0164 -0.78 9.376 0.401 0. 12. 717 0.0126 -0.880 0.0750 0.97 -1.0128 0.354 0.385 0.94 -1.0206 0.0522 0.376 0.0170 0.66 3.354 0.0215 0.0275 0.12 1.42 0.0100 0.0218 0.28 9.26 Rn = Q -6.01 -2.340 0.793 0.0323 1.0234 0.27 cd 0.0094 0.352 0.745 0.692 0.0350 199700.96 -0.16 3.66 3.0446 0.0263 -0.23 13. c1 -0.16 4.0212 0.329 0.048 0.0119 0.18 5.0114 0.04 -3.169 0.19 5.63 2.551 0.22 7.952 0.09 1.1999 J5012-PT Fig.039 0.002 0.26 10. 0.74 6.1387 0.0147 0.14 2.25 7.947 0.0092 0.19 5.12 2.008 0.26 7.151 0.550 0.0090 0.98 -2.0158 0.93 0.0374 -0.989 0.464 0.1930 HQ2/9B-PT Fig.983 0.124 0.144 0.0149 0.0286 1.218 0.04 -0.0222 0.0096 0.0078 0.21 5.01 -3.0282 0.0286 0.0121 Rn -1.20 4.79 Rn = Q -2.0203 0.0731 100200.0108 0.12 1.24 8.82 c1 c1 -1.77 9.78 9.0080 0.0089 0.230 0.0531 0.370 Airfoils at Low Speeds 1. 12.1300 13.18 Rn = "' -6.20 4.06 c1 cd -0.534 0.81 0.40 0.0215 0.25 9.0089 0.0119 0.0159 -0.80 Rn = Q -2.033 0.645 0.76 0.31 cd -0.23 -2.05 -5.77 9.0265 0.0223 0.391 0.26 9.02 -2.0321 0.0094 0.027 0.210 0.080 -2.530 0.411 0.0119 0.26 7.0597 303100.0126 0.30 1.295 0.0437 -0. = Q cd -0.051 0.22 6.0406 0.0152 0.0084 0.072 0.814 0.130 0.494 0. 760 0.93 0.0128 0.0082 0.115 -0.19 4.0163 0.800 0.13 3.907 0.66 3.26 7.04 71 1.222 -0.81 Rn = Q -2.98 -1.534 0.0084 0.0139 0.66 -4.545 -0.97 -0.0086 0.22 7.0185 0. 739 0.22 7. = Q -3.25 7.97 -1.568 0.24 6.0564 10.0140 -0.22 7.811 0.936 0.0288 0.010 0.802 0.0096 0.26 Rn Rn c1 cd -0.93 0.226 0.11 2.0123 0.0134 0.0276 0.032 0.0165 0.190 0.492 0. = c1 "' -0.194 0.14 2.79 12. = c1 "' -6.46 -0.0280 .173 0.0128 0.0242 0.240 0.16 4.983 1.525 0.0078 0. = "' -6. = " -2.09 1. 742 0.19 4.854 0.74 6.0407 0.518 0.769 0.40 0.0298 0.42 -1.0088 0.0153 0.25 9.0102 0.700 0.92 1.18 3.160 0.98 -3.007 0.0102 0.950 0. c1 Q -6.71 5.0102 0.16 4.603 0.72 6.407 0.342 0.97 -2.0131 0.23 8.44 0.0227 0. c1 -0.22 7.526 0.66 3.1333 HQ2/9B-PT Fig.907 0.458 0.1300 0.0287 0.0224 0.19 4.50 Rn 202400.0097 0.0082 0.345 -0.96 0.210 0.90 0.92 0.97 -0.49 Rn 59000.0519 0.529 0.92 -1.04 -5.09 1.25 6. 723 0.452 0.014 0.531 0.0093 0.0081 0.74 6.0222 0.24 7.27 10.13 3.25 7.79 9.0185 0.0091 0.73 6.96 -2.0206 0.0396 0.0096 0.0108 0.1999 299200.0225 0.0267 0.260 0.112 0.53 Rn 60100.0161 -5.0107 0.764 0.0248 0.092 -0.0140 0.540 0.0122 0.112 0.368 -0.11 2.193 0.152 0.73 6.0366 0.99 -1.0151 0.0105 0.28 10.0107 0.0082 0.31 10.0131 0.0156 -0.0323 1.86 -4.820 0.946 cd 0.0159 0.91 -0.0120 0.915 0.0206 0.038 0.596 -0.26 9.447 -0.0101 0.0135 0.20 6.75 8.528 0.079 0.0081 0.915 0.20 5.0204 0.201 -0.31 10.072 0.629 0.0107 0.12 2.11 1.48 -5.054 0.822 0.97 -1.02 -4.689 0.009 1.00 -2. = Q c1 -0.0158 -0.0151 0.0537 MB253616-PT Fig.0201 0.0189 0.0153 0.96 Rn cd -0.18 5.264 0.72 6.0344 0.0308 -0.870 0.27 -0.09 1.0129 0.122 0.20 4.483 0.0250 0.442 0.0088 0.072 0.855 0.378 0.94 -1.31 10.678 0.82 12.0231 0.540 0.0166 0.485 0.0113 -0.315 0.0115 0.16 3. 12.13 1.378 0.0455 151300.0169 0.43 0. 776 0.107 0.051 1.66 3.18 4.071 -2.65 3.549 0.693 0.01 -4.235 0.19 4.12 1.28 10.605 0.17 4.071 0. 0.41 0.77 9.24 8.231 0.0226 0.385 0.078 0.78 0.215 0.31 10.031 0.843 cd 0.0115 0.94 Rn cd -0.078 0.26 10.933 1.99 -2.18 4.12 1.19 4.767 0.851 0.510 -0.480 0.842 Rn= 97800.0218 0.072 0.631 0.814 0.036 0.29 9.30 13.983 0.0175 0.226 0.802 0.0152 0.78 9.78 9.20 6.0134 -0.010 1.043 0.0122 -0.365 0.0115 0.10 1.289 0.21 6.0159 0. 930 1. 12.068 0.2198 200100.451 0.52 Rn 56500.0327 0.0494 c1 cd -6. c1 -0.285 0.0262 0.911 0.423 0.0113 -0.548 0.95 -1.90 0.29 Rn -2.738 0.165 -0.0194 0.0124 0.080 -2.0579 201900.0196 0.292 0.599 0.15 3.0218 0.936 0.31 10.0181 0.0125 0. 0202 0.67 3.28 -0.331 cd 0.691 0.0132 0.73 6.0158 0.95 -0.0327 0.0564 204700.81 cd cd -0.380 -0.93 -0.215 0.654 0.66 3.348 0.0141 0.13 1.076 0.916 0.824 0.0326 0.0398 0.04 -5.0191 0.0116 -0.26 8.96 -1.0082 0.0086 0.93 -0.13 1.0108 0.286 -0.0324 0.691 0.0149 0.989 cd 0. c1 cd "' -0.03 -4.00 -4.82 = 302500.802 0.632 0.0218 0. = c1 "' -0.0096 0.0096 0. 21 3.14 1.22 4.93 -1.0151 0.0501 254600. = "' -6.03 -3.0238 0.304 0.89 -0.91 -1.0161 -0. 730 0.24 4.849 0.0137 0.0378 0.152 0.0336 232000.13 1.0395 1.914 0.002 0.968 0. Ct -0.25 5.94 5.18 2.0253 0.0127 0. 789 0.0303 0.476 0.500 0.24 5.370 0.070 0.369 0.58 7.0226 0.28 7.24 c.21 3.20 3.146 0.0135 0.13 4.139 0.053 0.971 0.0356 0. 994 0.0135 0.0147 0.26 6.0159 0.0306 0.0109 0.92 -2.25 5.0177 -0.32 10.010 0.90 ·0.16 1.0185 1.0167 -0.23 4.26 6.37 11.14 1.0228 0. " -0. Ct 0.268 0.02 9.28 7.613 0.90 -0.30 0.22 4.18 2.326 0.90 -0.0138 0.220 0.663 0.27 7.517 0.28 6.0298 0.90 -0.91 Rn -1.95 Rn ·3.25 6.0180 -0.17 1.0216 0.354 0.30 8.0228 0.0211 -0.95 0. " -0.26 5.28 6.28 7.32 0.348 0.0235 1.0158 0.247 0.21 3.0169 0.202 0.870 0.19 2. 705 0.0118 0.28 6.998 0.0114 0.0179 1. = c.93 Rn cd -0.Oll7 0.181 0. q cd -0.801 0.040 0.427 0.522 0.38 8.29 Rn = " -7.0342 -3.437 0.978 0.0881 306300.0273 0.28 7.939 1.240 0.991 203600.0138 0.697 0.118 0.92 -1.0218 0.32 9.0433 126700.926 0.170 0.97 -3.0199 0.115 0.0141 0.0103 0.952 0.0125 0.181 0.0160 0.33 11.0156 -0.0161 0.0208 0.265 0.36 10.0151 0.30 4.18 2.17 1.899 0.0448 11.Q108 0.93 -2.004 0.91 -0.20 3.Dl68 0.716 0.0134 0.0272 0.0157 0.0313 0.0136 0.0311 0.0167 0.816 0.94 -2.87 0.33 10. 782 0.31 10.587 0.0344 0.817 0.87 1.20 2.0288 0.0522 Rn 200100.0291 0.391 0.93 -2.0202 1.907 0.0119 0.0252 0.029 O.720 0.949 0.30 9.974 0.0149 0.0121 0.880 0.0325 cd = 149700.77 6.0218 1.17 1.078 0.028 0.92 -0.0258 0.95 -2.89 0.17 2.543 0. cd Ct " -0.88 0. Ct -0.027 0.549 0.0228 0.0329 0.32 9.060 0.36 10.367 0.93 -1.0297 0.0191 ·0.32 0.007 0.0271 0.894 0.0181 0.34 9.0154 0.87 0.0196 1.28 7.291 0.31 14.060 -2.0113 0.39 13.92 -1.761 0.29 8. Ct -0. cd -0.0261 0.463 0.0210 0.043 0.0344 1.91 Rn -1.0143 0.303 0.0224 1.033 0.871 0.0464 cd 0. = c.829 0.166 0.961 0.0179 0.341 0.366 0.32 8.92 Rn ·2.0216 0.032 0.446 0.17 2.89 -0.0284 0.120 0.32 9.0141 0.30 Rn = " -2.0180 0.0371 cd MB263616·PT Fig.141 0.0506 0.0212 1.24 5.21 3.0266 0.23 5.21 3.30 8.0173 0.32 9.0240 1.658 0.88 -0.964 0.23 4.694 0.047 0.0228 0.88 0.0148 0.298 0.975 0.0251 0.0176 1.23 4.0194 1.988 0.947 0. 12.0201 0.0443 250100.952 0.89 0.0118 -0.434 0.0193 1.85 0.30 7.29 8.272 0.31 = " -2.14 1.0391 301600.0144 -0.134 -0.0162 1.26 6.0307 0.24 5. Ct -0.18 3.0201 1.465 O.68 2.27 6.0100 0.193 0.0149 0.668 0.29 8.0180 0.131 0.001 0.460 0.0333 0.86 0.8ll 0.704 0.0168 O.0224 0.0321 0.870 0.860 0. -3.0276 0.0214 1.546 0.30 7.463 0.0251 0.0153 0.41 = " -2.91 ·1. 12.013 0.558 0.19 2.25 5.24 5.ll3 0.61 ll.0300 l.31 13.31 11.0178 M06-13-128-PT Fig.346 0.28 7.863 0.98 Rn -3.31 10.15 cd 0.200 ·4.88 -0.736 0.033 0.266 0.30 0.0209 0.666 0.0227 0. 788 0.25 4.0135 0.31 12.18 2.91 Rn ·1.30 c.972 0.541 0.462 0.28 7.595 0.0254 0.0178 0.028 0.31 -0.0218 1.0414 0.0273 0.56 Rn 201600.0492 0.32 8.0497 0.6ll 0.582 0.29 9.000 0.0534 -0.0193 0.88 -0.0159 0. = c.517 0.0176 0.0111 0.17 1.29 10.0151 0.878 0.0218 1.0126 0.0211 0.0201 0.0147 0.621 0.0423 177600.0965 101300.952 0.26 6.30 9.21 3.29 8. cd "' 0.611 0.20 3.26 6.32 10.21 9. = "' ·2.0177 0.0266 1.0115 0.563 0.0128 0.050 0.31 12.82 10.610 0.0183 1.27 7.0280 0.29 7.26 6.283 0.0150 0.0254 0.0414 0.0116 0.0251 0.982 0.015 1.36 = " .044 0. cd 2.364 0.35 9.922 0.066 O.830 O.31 -1.49 3.0121 0.326 0.22 4.348 0.0103 0.034 0.202 0.0332 cd 10.109 0.37 12.671 0.675 0.98 -4.0180 0.998 0.241 0.20 3.0179 0.800 0.959 0.029 0.388 0.018 O.0143 0.38 = " -4.0459 0. Ct -0.958 0.407 0.0116 0.237 -5.111 0.0144 0.0206 Rn -4.991 0.16 1.955 0.22 3.0112 0.162 0.30 -0.0629 1.0189 0.0129 0.051 1.30 10.961 0.127 0.0219 0.615 0.99 -5.998 0.19 2.0213 0.28 7.0251 0.90 -0.0114 0.0378 0.0101 0.107 0.0390 0.0278 1. 766 0.510 0.54 Rn 59600.Dl68 0.18 2.0110 0.624 0.562 0.396 0.30 8.87 0.83 0.0187 0.206 0.14 1.28 6.96 ·2.24 5.642 0.0172 0.0150 0.789 0.976 0.749 0.0183 0.0270 0.0197 0.890 0.22 4.748 0.0127 0.0270 0.0235 0.30 0.059 0.0129 0.Dl08 0.0130 0.364 0.0714 0.31 9.88 0.090 1.213 0.047 0.23 4.103 0.30 8.0586 0.778 0.309 0.0180 0.955 0.0259 0.90 -0.0122 = " ·2.92 -1.0182 0.0145 1.0302 0.034 0.85 0.920 0.17 2.002 0.25 5.053 0.0165 LOll 0.24 4.034 0.88 0.053 0.24 5.0518 0.0254 0.05 -5.722 0.0105 0.0410 0.353 0.24 5.16 1.09 -6.008 0.26 6.970 0.0324 0.0133 -0.30 11.0250 0.13 1.175 0.19 3.0180 0.0175 cd cd 371 .157 0.97 -1.90 -1.0146 0.Dl08 0.Chapter 13: Tabulated Data 0.650 0.22 4.976 0.30 8.051 0.444 0.38 11.183 0.0267 0.238 0.0174 0.23 4.943 0.97 -3.430 0.30 8. 22 3.044 -0.21 7.26 5.20 2.33 9.0118 0.29 7.001 0.253 0.22 7.23 4.0143 0.0155 0.303 -0.15 4.0213 0.20 2.0558 M06-l3-l28-PT Fig.545 0.36 9.0152 0.353 0.38 12.91 -1.0264 0.0084 0.222 = " -2.95 -0.0084 0.733 0.0223 1.617 0.32 8.331 1.91 -2.307 0.19 2.727 0.0175 1.0165 0.929 0.569 0.507 0. c.0228 0.22 3.19 6.529 0.0140 0.0205 0.933 1.0187 1.0160 0.35 9.0202 1.0174 0.0239 0.808 0.17 1.01 ·4.0201 1.21 3.0350 0.0418 0.313 0.0159 0.37 c.448 -0.22 3.0225 -0.01 -0.98 -1.309 0.430 0.0194 1.0196 0.0184 0.91 -0.85 0.88 -0.293 -0.0305 0. 01 -2. cd Rn = 0. c.906 0.0207 0.0200 1.00 -1.748 9. c.0228 0.10 2. 01 M06-l3-128-PT Fig.422 0.306 0.0528 0.03 -5.32 8. = 01 c.0152 1.38 11.16 4.0133 0.23 8. cd -0.232 1.0117 0.24 8.39 Rn = " c.0233 0. 0.616 0.0106 0.32 8.0100 0.90 -1.03 ·4.28 6. 0.10 2.357 0.215 0.417 0.0122 -0.317 0.346 0.0132 0.523 0.009 1.462 -0.36 0.0156 1.17 1.0244 1.091 1.0081 0.0145 0.2·3 8.504 -0.211 0.354 0.0162 1.544 -0.193 0.28 6. c.0501 0.24 9.0570 0.0462 1.198 0.0172 1.308 0.268 1.10 1. 0.0083 0. 12. 12.0143 0.370 0.0283 0.0094 -0.0192 -0.0125 -0.132 1.068 0.319 0.29 7.510 0.372 Airfoils at Low Speeds 8.627 0.0192 1.167 0.178 1.0099 0.090 0. " -6.206 0.0217 0.913 0.03 -3.04 -5.99 -1.0141 0.0106 -0.127 0.0216 1.0119 1.89 -1.122 0.33 8.281 1.28 6.816 0.18 5.359 0.305 0.0175 0.315 0.23 4.0309 0.13 3.0162 0.09 -0.84 0.37 12.0511 1.0135 0.D205 0.35 10.0148 -0.170 0.0502 NACA 0009-PT Fig.0111 0.000 0. c.006 1.293 " c.435 0.33 8.25 4.540 0.359 1.24 -0.86 0.719 0.0156 0.011 0.679 0.21 6.381 -0.34 12.298 0.29 7.0129 0.34 9.0161 0.646 0.314 0.245 0. 798 0.0198 0.618 0.00 -2.0155 0. Oi -2.596 0.092 0.34 9.26 5.0215 1. = 01 -3.0502 251600.0176 1.0506 0.0170 0.460 0.0433 0.36 Rn 0.19 2.119 0.0165 1.92 0.0149 0.0163 0.15 3.0233 1. 717 0.60 Rn 59700.22 3.0288 -2.87 -0. 772 0.0185 0.0215 1.0163 0.0149 0.508 0.204 0.30 7.0172 0.0162 0.0140 0.31 7.405 0.325 0.626 0.36 10.22 4.709 0.26 10.15 4.28 6.27 5.916 0.90 -0.105 0.0143 0.39 11.224 0.031 1. 709 0.32 8.217 0.0230 -1.352 1.205 0. 0.109 0.0148 0.224 -0.25 5.268 0.321 1.0169 0.25 5.832 0.0157 0.0194 1.0348 0.00 -2.0150 0.0089 0.354 0.58 Rn = 199500.0277 0.37 Rn = " -3.0171 0. c.0175 0. -0.94 .18 2. c.216 0.38 11.0170 0.507 0.031 0.0328 203600.37 10.0124 0.0267 0.347 0.057 0.0718 199000. 01 -3.37 cd 0.0136 0.108 0.23 9.30 6.297 c.29 6.368 0.313 0.326 c.17 2.38 11.332 1.0128 0.512 0.35 12.178 0.366 0.25 4.684 0.0694 200900.18 2.0184 0.008 0.100 0. cd -6.90 0.0085 0.0146 0.0228 0.11 2.998 1.92 0.0087 0.0097 0.104 0.31 7.0358 0.99 -2.0165 0.19 5.34 8.694 0.198 0.34 9.731 0. cd -3.303 0.33 8.580 0.99 -3.90 -0.735 0.22 4.96 -0.0153 0.332 1.0213 0.19 6.0156 0.91 -1.303 0.0336 Rn = 203600.35 c.38 12.25 5.0138 0.0178 0.86 0.669 0.0732 302900.611 0.1067 0.0174 0.88 -0.0163 0.34 10.25 Rn = 01 -6.16 1. = c.57 Rn 200900.211 0.0234 1.0725 204100.0178 1.89 -0.103 1. 720 0.23 7.38 c.309 1.24 4.0150 -0.0142 0.016 0.348 0.914 1. 0.526 0.84 0.830 0. 0.18 1.0168 0.31 7.21 7 .37 12.0118 0.16 1.0394 Rn = 150600.0175 0.0175 0.02 -5.704 0.835 0.311 0. 12.0216 0.0185 0. cd -6.88 0.00 -3.0306 1.194 1.0465 0.37 Rn 1.0438 0.0149 0.37 10.0183 -0.23 4.356 0.283 1.906 0.060 0.0179 0.210 0.611 0.33 8.377 -0.13 3. c.38 11.0209 0.0167 0.04 -5.98 -2. 0. -2.410 0.381 0.412 0.611 0.0189 1.90 -1.0189 1.21 6.440 0.38 11.100 0.388 0.24 8.257 0.036 0.84 0.0149 0.0149 0.0629 .673 0.0185 0.86 0.808 0.0113 -0.28 6.0499 0.913 1.0195 0. 0.439 0.0339 0.0156 0.19 2.0155 0.91 -2.405 0.0201 0.1493 100300.0096 0.100 1.0463 1.38 11.222 0.15 3.19 5.90 -1.0309 0.0433 1.091 0.0156 0.112 0.113 0.309 0.0171 1.93 -2.29 7.267 1.0205 1.523 0.0158 0.460 0.90 -1.37 12.007 0.816 0.27 5.11 2.0189 O.0234 0.231 0.0142 0.0159 0.21 3.0166 0.0337 0.378 0.0216 0.36 12.0161 0.351 1.998 0.86 -0.261 0.35 9.364 0.0351 1.0245 0.0180 0.0148 0.0313 0.17 1.522 0.438 0.190 1.37 10.0181 0.19 5.16 1.0579 0.417 0.521 0.0705 Rn = 307400.37 10.0172 0.473 0.0140 0.96 -0.25 Rn = 01 0.33 13.396 0.465 0.0135 0.32 0.35 9.380 0.0138 0.36 10.0126 0.37 11.601 0. 787 c.26 5.15 1.0118 0.0130 0.104 0.0109 0.17 1.934 0.0148 0.410 0.21 3. -3.09 1.0184 0.004 0.85 0.24 c.24 4.D7 1.007 0.0271 1.34 9.0183 0.326 0.0223 0.0142 0.22 3.027 0.29 6.325 0.022 0.38 11.25 5.28 6.632 0.35 9. -0.19 2.141 -0.0141 0.0162 0.92 -2.816 0.94 O.37 11.0138 0.25 0.0192 0.0141 0.0178 1.0241 -0.29 7.0422 0.0189 1. c.195 0.36 10.90 -1.348 1.86 -0.516 0.21 3.041 0.0238 0.15 4.0179 0.136 0.729 0.35 10.0194 1.91 0.727 0. 01 -4.63 3.439 0.922 0.0122 0.18 1.0156 0.301 1. cd -0.0168 0.0265 0.38 8.08 1.0230 0.972 0.0166 1.206 0.686 0.0068 0.16 2. " -1.23 5.21 3.84 0. 12.19 2.253 0.34 10. c.383 0.040 0.0364 1. -6.25 -0.0127 0.24 6.04 -2.0523 301000.29 0.0242 0. c.537 0.703 0.0233 0.0180 0.413 0.0131 0.0415 Rn = 104200.248 0.11 1.0068 0.0100 -0.873 0.0199 0.25 9. cd 0.411 0.0099 0.015 0.0146 0.978 0.92 0.0391 0.673 0.312 0.0099 0.22 3.0549 373 .0252 0.0479 Rn = 303300.0129 -0.20 7.0206 1.0217 0.0231 1.0235 NACA 6409-PT Fig.0084 0.0272 0.25 c.0276 1.13 2.66 Rn 61400.118 0.0146 -0.0115 0.0101 3.119 0.02 -0.066 0.00 -2.0222 0.168 0.20 4.792 0.35 11. 724 0.184 -0.0110 0.184 0.601 0.18 1.07 1.16 4.0149 0.0213 -0.0240 0.366 0.0179 -0.97 -1. 792 0.002 0.361 0.27 8.11 1.94 -1.0127 0.26 5.0112 0.639 0.0286 0.24 9.0142 0.291 0.537 0.30 10.19 c.887 cd 0.21 4.088 0.280 1. c.18 3.0081 0.572 0.014 0.0158 1.574 0.385 -0.36 Rn = " -0.206 0.559 0.138 0.0094 0.027 0.0180 0.22 c.006 1.15 3.26 Rn cd 0.0140 0.0349 1.35 10. " -1.873 0.0088 0.19 1.84 0.273 0.240 0.95 -0.053 0.0280 0.190 0.274 0.234 0.33 12.30 10.973 0.0124 -0.29 9.0143 -0.0096 -0.301 0.0075 0.0205 1.030 0.0313 0.0120 0.350 0.0124 0.09 1.618 0.301 0.33 11.96 -2.030 0.0125 -2.0413 1.25 c.153 O.29 7.0143 0. 725 0.27 -0. 0.30 11.30 c.08 0.22 5.560 c" 0. -0.22 7.19 5.330 0.673 0.99 0.0080 0.26 7.00 -2.244 0.0365 0.20 6.0091 0.613 0.17 4.140 0.15 3.01 -4.459 0.324 0.385 -0.048 0.0201 0.005 0.15 2.325 0. c" -0. cd 0.24 8.32 0.34 Rn = 0.21 7.0210 0.0200 0.0154 -0.60 1.98 -2.448 0.0203 0.240 0.0121 0.36 0.10 2.29 14.0098 -0.0494 150200.36 13.24 8.012 0.0237 1.90 0.325 1.1167 0.05 c.88 0.95 -0.303 0.493 0.96 -2.369 0.0095 0. -0.484 0.23 8.17 4.32 8.62 2.95 -0.629 0.19 5. cd 0.35 10.35 12.625 0.5411-PT Fig.20 3.0131 0.62 Rn = 58200.063 0.683 0.007 0.0154 0.0159 0.28 8.092 1.16 4.665 0.93 0.996 0.0102 0.670 cd 0.0153 0.0092 0.0088 0.0110 -4.602 0.0333 79400.0431 Rn = 200600.223 0.0170 0.98 -1.438 cd 0.0083 0.0121 0. " -0.26 6.155 c" 0.21 6.20 6.24 7.0080 0.0148 1.96 -0. 765 0.0298 1.0188 0.785 0.0127 0.91 -0.0099 0.01!0 0. cd c. -0. 753 0.0109 0.0265 1.454 0.912 1.11 1.15 1.503 0.0486 0.0193 1.0123 -3.41! 0.20 6.18 5.0308 1.23 7.329 0.29 10.19 5.31 9.0108 0. " -3.541 0.970 0.055 0.373 0.0179 1.365 0.Dl66 0.17 3. cd -0.47 0.0191 -0.362 0.180 -0.370 0.0133 0.294 -0.549 0.21 5.32 9.347 0.0247 1.044 0.0369 0.0120 0.0136 0. -0.244 0.20 6.0376 0.17 2.0163 0.33 11.0221 0.88 0.23 6.24 4.0133 0.D206 0.20 2.64 Rn = 62800.23 8.339 0. cd cd -0.24 5.18 4.0104 0.018 0.0122 0.0334 0.339 1. = " c.676 0.10 1.974 0.0134 0.805 0.26 6.28 7.465 0.481 0.934 0.26 8.347 0.140 -0.093 0.090 0. = " -5.0131 0.0571 Rn = 147300.0309 0.0279 0.18 3.0227 0.271 0.0242 1. " -3.0247 0.0092 0.32 8.926 0.438 0.089 0.265 1.536 0.342 1.0126 NACA 64A010-PT Fig.0374 0.880 0.0114 0.33 13.049 0.0339 0.0190 -0.23 7.0213 0.0370 1.140 0.0276 1.83 0.0084 0.343 0.026 0.0104 0.26 7.31 9.0162 0.35 Rn = " -0.0129 -0.0336 0.95 -0.96 -1.0124 0.0277 1.398 0.0296 0.87 0.01 -3. -0.345 0.029 0.0107 0.227 0.13 1.175 1.12 1.821 0.91 -0.15 4.22 7.333 0.0185 0.21 4.179 0.97 -1.575 0.0093 0.14 1.565 Rn Rn = " -5.328 0.0156 0.25 10.465 0. " -5.0136 0.0174 -0.315 1.98 -1.299 0.91 0.20 5.525 0.155 0.127 0.14 2.320 0.434 0.17 2.16 3.35 13.0152 -0.0271 0.0958 0.730 0. 716 0.463 -0.475 -0.528 0.0223 0.21 2.0073 0.99 -2. 739 0.646 0.30 Rn = " -3.207 0.21 6.596 0.0229 0.0175 -0.24 6.40 0.0155 0.0139 0.0325 0.067 0.266 -0.468 0.98 -1.0124 0.0181 0.0276 1.383 0.0202 1.91 0.36 12.260 " c.94 -1.36 14.384 0.0277 0.0119 0.0165 0.486 0.0203 0.92 -0.0114 0.27 6. 12.191 0.04 -3.0093 0.0073 0.0207 0. = " -0.94 c.0220 1.0314 0.113 0.0290 1.0521 102600.14 3. -5.0100 0.0075 0.0135 -0.368 0.931 0.13 2.631 0.31 11.621 0.130 0.21 5.580 0.0177 O.32 9.874 0.18 2.18 2.0305 0.0093 0.95 -0.525 0.26 7.0138 0.02 -0.1196 1.383 0.23 6.0075 0.0263 1.0187 0.527 0.914 0.404 0.74 c. "' -6.0177 0.0181 -4.91 -0.811 0.19 4. 12.23 4.22 5.28 9.28 7.0200 0.18 3.821 0.21 4.52 -0.353 0.23 5.0303 1.714 0.0293 1. 779 0.0130 0.Dl65 1.0643 101500.851 0.33 12.24 -0.0103 0.0124 -0.0230 5.715 0.23 4.36 13.450 0.90 0. c.4 71 0.0262 0.0180 0.763 0.00 -0.17 4.20 3.0137 0. 70 6.37 11.Chapter 13: Tabulated Data Rn = 302900.0189 0.28 9.441 0.0575 NACA 2.28 8.0382 200100.0366 203300.239 0.652 0.013 0.344 0.0129 0.10 1.0202 -5.0135 0.0112 0.829 0.22 7.31 10.723 0.14 3.327 0.526 O. 700 0.015 0. 90 0. " -0.249 0. -0.368 0.30 9.0149 0.140 -3.28 8.13 1.0625 151600.0112 0.31 10.0101 0.22 5.97 Rn -2.32 11.19 4.214 0.0187 0.92 -0. 1342 201200.721 0.997 0.32 13.0096 0.949 0.0121 0. " -0.31 -0.0095 0.0132 0.19 3.0122 0.12 1. 763 0.112 0.69 5.237 0.981 cd 0.0559 = 202700.29 9.0144 0.125 -0.79 8.0144 -0.30 10.00 Rn cd 0.32 = c.000 1.32 11.0171 0.0693 -0. 730 0.076 0.91 -0.0088 0.13 1.91 0.1993 202300.989 0.21 4.32 0.31 c.0144 0.0115 0.94 -1.022 0.49 Rn -2.0312 0.040 1.27 7.055 0.0194 0.32 11.12 1.965 0.0407 0.924 1.062 0.139 -0.0165 8.24 6.164 -3.0121 0.0093 0.93 -0.70 5.025 0.0109 0.627 0.146 0.30 -0.107 cd 0.90 -0.0402 RGIS-PT Fig.716 0.33 12.017 0.33 c.19 4.884 0.0120 0.691 0.16 2.92 0.30 9.96 -0.78 8.0110 0. 12.0123 0.051 1.0078 0.227 0.131 0.801 0.0293 1.38 1.0246 1.0248 0.16 2.27 7.0154 -0.29 9.0213 0.452 0.38 1.07 44 151900.751 0.229 0.890 cd 0.43 -0.0159 0.69 Rn 102300.0185 RGIS-PT Fig.447 0.086 0.090 cd 0.05 1.14 1.91 0.0096 0.0117 0.92 Rn cd 0.70 5.0144 0.38 1.032 0.15 2.0164 0.0191 9.0096 0.050 0.141 0.0088 0.96 -1.0076 0.24 6.31 10.0087 0.96 -1.027 0.164 -0.32 0.0104 0.24 5.0505 1.0301 1.0353 0.29 9.0083 0.0309 1.1002 0.469 0.23 5.19 3.0124 0.0276 1. 728 0.124 0.27 7.065 0.0083 0.0165 0.214 0.21 4. = c.0075 0.0131 0.27 8.0182 0.0082 0.095 0.500 0.0361 1.0199 0.804 0.31 10.009 1.91 -0.049 0.0103 0.00 -0.0152 0.21 4. c.962 0.032 0.013 0.0100 0.15 2.32 10.0131 0.94 0.32 12.338 0.055 0.0188 0.069 0.0133 0.040 0.526 0.843 0.038 o.566 0.91 0.107 0.30 0.26 7.0100 0.19 3.549 0.15 2. " -0.0195 0.035 0.17 3.0309 11.0101 0.0107 0.0136 0.995 1.0121 0.0108 0.215 0.16 2.97 Rn cd cd 0.856 0.96 -2.0125 0.31 10.31 9. = " -3.0139 O.0116 0.0142 0.04 79 1.026 99600.0346 1.130 0.633 0.25 6.0087 0.998 0.89 0.23 5.058 0.062 0.850 0.13 1.25 6.0108 0.639 0.879 0. 723 0.925 0.0237 1.13 2.20 4.32 10.30 9.92 -0.248 cd 0.509 0.93 -0.97 -0.0264 0.0121 0.449 0.198 0.45 -1.0141 0.21 4.833 0. -0.24 6.33 Rn cd 0.0188 0.17 3.0122 0.0101 0.091 cd 0.0106 0.44 Rn -1.0082 0.97 -1.0233 1.225 0. " -0.747 0.506 0.24 6.320 0.665 0.965 1.917 0.0119 0. 12.17 2.0109 0.0125 0.26 6.860 0.32 11.0086 0.0174 0.0126 0.0097 0.486 0.18 3.0107 0.95 -2.215 -3.084 0.0101 0.0081 .0107 0.0159 0.22 5.68 Rn 59600.065 0.21 4.0779 152800.94 -1.0120 0.0116 0.017 0.0094 0.048 0.0091 0.997 0.0172 -0.043 0.0188 0.0121 0.0136 0.012 0.0219 0.93 -0.622 0.0127 0.0157 0.997 0.23 5.340 0.133 -2.92 0.0226 1.0387 1.430 0.879 0.0135 0.0148 0.0162 -0.911 0.0150 0.93 Rn -1.051 0. cd -3. " -0.78 8.0144 0.20 4.442 0.0129 0.26 6.877 0.221 0.92 -0.0314 0.0134 0.34 7 0.0094 0.148 0.0249 10.0299 1.089 0.39 1.239 0.082 0. "' -0.089 0.0463 Rn = 313500.180 0.0118 0.23 5.0396 1.0087 0. "' c.31 11.0109 0.663 0.0083 0.652 0.28 cd 0.0140 0.316 0.26 6.748 0.30 -2.93 -1.0110 0.24 5.28 8.28 8.04 7 0.0085 0.27 7.110 0.0164 0.99 -2.0098 0.0220 1.32 Rn = " -3.0257 0.97 -1.0086 0.12 1.27 7.804 0.0134 0.103 0.0127 -0.114 -3.189 0.0205 0.125 0.0117 -1.343 0.23 5.0116 0.0362 1.0191 0.049 0. 0.32 0. = c.198 -3.25 6.998 0.30 9.0120 0.0139 O.0190 200900.77 8.90 0. cd -0.440 0.0223 1.29 9.877 0.0114 -1.0128 0.096 1.0401 1.0210 0.907 0.33 1.28 7.0135 -2.0178 0.33 1.647 0.0113 0.995 0.0092 0.541 0. " -0.0210 0.95 -2.93 -0.15 2.987 1.232 -4. " -0.059 0.216 -3.16 2.30 10.392 0. = " -3.25 7.0129 0.198 -3.027 0.034 0.946 0.0289 1.139 -2.461 0.30 cd 0. 749 0.70 Rn 101700.19 3.099 0.0156 0.0130 0. " -0.440 0.075 1.33 11.0097 0.37 4 Airfoils at Low Speeds RGlli-PT Fig.18 3.26 6.91 0.743 0.188 -3.310 0.053 0.78 8.29 8.0186 0.0247 0.0282 1.22 5.543 0.26 6.17 2.929 0. = c.338 0.0469 0.128 0.0097 0.853 0.0166 0.0183 -0.065 0.0125 0.330 0.32 11.14 2.0317 RGIS-PT Fig.0253 0.329 0.09 2.540 0. = " -2.93 -0.31 10.120 0.70 5.137 0. -0. 772 0.0289 0.16 2.508 0.0159 0.69 5.546 0.0069 0.0102 0.31 1.331 0.45 -1.0242 0.0172 -0.31 10.31 9. = c.0106 0.0087 0.92 -0.29 8.29 9.0141 -0.0087 0.97 -2.057 1.12 cd 0.050 0.0105 0.839 0.17 3. 750 0. c.175 0.014 7 0.26 7.0100 0.065 0.25 7.046 0. = c. c.051 0.16 2.098 0.D108 0.0536 146100.30 9. 718 0.24 6. 754 0.93 -1.964 0.092 0. " -0.11 1.92 -0.349 0. 12. 12.28 8.0159 0.544 0.19 3.29 8.D165 0.845 0.24 6.345 0.028 0.0280 305100.39 1. = c. " -0.0090 -0.31 10.29 Rn c.0410 Rn = 302600.33 -1.630 0.564 0.93 0. = c. c.32 11.0097 0.20 4.648 0.076 0.23 6.021 0.549 0.32 11.0084 0.945 0.0113 0.0097 0.71 Rn = 101000.90 0.0117 0.12 1.820 0. 31 9.023 0.21 4.15 2.0074 0.31 12.24 5. 12.247 0.207 0.014 0.10 1.0126 0.919 0.233 0.24 5.996 0.0133 0.29 9.332 0.0092 0.836 0.454 0.91 0.93 -1.0193 0.633 0.1151 153300.430 0.147 0.28 cd -0.632 0.451 0.0119 0.0142 0.934 0. 0.97 -2.048 1.158 -0.75 Rn 102200.31 12.425 0.998 0.18 2.470 0.12 1.0095 0.1301 198700.0189 0. 720 0.93 -1. 727 0.049 0.0080 0.655 0.35 c1 cd -0.91 -0.92 0.97 -0.32 10.23 4.0139 0.044 0.870 0.344 0.89 0.300 0.0092 0. 12.047 0.26 7.19 3.101 0.112 0.20 3.0084 0.986 0.20 3.74 Rn 58800.091 1.0109 0.91 -0.883 0.748 0.16 3.0088 0.0084 0.0144 0.0392 1. 716 0.0072 0.0168 0.0124 0.078 0.834 0.24 5.0119 0. = c1 "' 0.379 2. c1 -0.96 -1.0120 0.0090 0.30 10.0131 0.0116 0.33 11.97 -3.0269 10.638 0.653 0.32 Rn = "' -2.0136 0.747 0.Dl03 0.25 6.0086 0.188 0.0133 0.0146 0.0123 0.29 8.0119 0.25 6.0102 0.0127 0.19 4.089 0.32 11.0073 0.099 0.657 0.30 10.029 0.055 0.0289 1.97 -0.0192 0.094 1.26 6.563 0.0306 0.0136 0.15 1.035 0.854 0.724 0.0097 0.98 0.0190 -0.0153 0.0079 0.333 0.10 1.15 3.32 11.25 7.792 0.11 1.18 4.33 11.0123 0.0113 0.859 0.986 0.0153 0.806 0.053 1.092 0.19 3.27 7.991 0.630 0.0566 1.921 0.30 9.0405 0.449 0.530 cd 0.17 4.89 0. = "' c1 cd -2.0115 375 .30 0.93 -0.18 3.0145 0.15 3.17 2.25 6.0135 0.0157 0.15 2.13 1.136 0.0140 0.086 0.32 1.0117 0.045 0.439 0.0271 0.0091 0.26 cd -0.12 1.080 0.001 0.28 8.0115 0.767 0. 749 0.0138 0.1328 302800.548 0.16 2.26 7.017 0.0287 1.922 0.0228 0.0120 0.32 10.0115 0.021 0.87 0.0180 0.30 9.0098 0.26 6.17 c1 -0.650 0.23 4.0111 0.0138 0.249 0.76 Rn 302800.827 0.079 cd 0.26 6.0325 0.0096 0.27 7.013 0.94 0.0111 0.550 0.540 0.128 0. c1 cd c1 cd -0.836 0.545 0. 718 0.0095 0.91 -0.33 Rn cd -0.25 6.695 0.096 0.048 0.93 -1.0264 1.065 0.96 -1.0402 RGI5-PT Fig.0242 99500.0194 0.0140 0.555 0.0105 0.89 -0.20 3.0175 0.0218 0.0125 0.017 -2.0373 1.0081 0.0097 0. = "' -2.805 0.092 cd 0.646 0.20 3.018 0.93 -0.927 0.042 0.0108 0.18 2.0457 Rn 201300.26 6.268 0.0116 0.943 0.27 Rn c1 = ex -2. = ex -3.142 0.996 1.642 0.0567 300800.641 0.0161 0.277 0.0073 0.0216 9.30 Rn Fig.28 7.20 5.317 0. = "' -2.346 0.002 0.057 0.047 0.809 0.16 3.29 8.24 7.21 5.14 1.087 0.0112 0.87 0.0103 0.0944 307600.108 1.0093 0.071 0.370 0.558 0.0308 200900.21 5.93 -1.0111 0.0240 0.92 -1.28 8. = c1 "' -2.0285 1. c1 ex -4.0231 0.0137 0.28 7.0373 1.345 0.27 7.745 0.34 11.243 0.32 10.304 0.092 0.92 0.14 2.053 0.0323 1.357 0.23 4.870 0. 726 0.446 0.26 6.30 8. = c1 ex -2.30 Rn = "' -2.0103 0. 12.097 0.0242 1.980 0.31 0.0212 0.0416 148800.16 2.92 -0.183 0.25 6.012 0.0194 0. 12.10 2.728 0.800 0.454 0.0096 0.0156 0.30 8.0120 0.31 10.116 0.0157 0.95 -2.0221 0.052 0.066 0.0070 0.0211 1.0129 0.555 0.240 0.0241 0.15 2.0299 1.15 1.800 0.13 1.72 Rn 101000.090 -0.28 8.0111 0.22 4.22 4.27 7. c1 0.0077 0.018 0.19 Rn = 0.654 0.22 5.959 0.23 4. 725 0.29 9.94 -1.23 5.97 -1.546 0.040 0.0133 0.26 7.994 0.906 0.0267 1.220 0.0097 0.0151 0.97 -1.32 -0.30 1.550 0.0882 3.31 10.31 11.0078 0.373 0.0170 -0.0333 0.94 -1.031 0.0583 1.19 3.0157 0.177 0. 790 0.23 5.0245 0.88 0.559 0.457 0.640 0.19 4.32 11.0420 154000.29 8.02 -2.91 0.250 0.054 0.220 0.0070 0.0194 0.874 0.357 0.995 1.0116 0.23 6.24 6.33 12.0125 0.535 0.145 0.31 9.0212 0.0169 -0.245 0.20 cd -0.008 1.922 0.0336 11. c1 cd -0.22 4.881 cd 0.29 9.054 0.0094 0.146 0.0080 0.007 4 0.051 0.0152 0.457 0.976 0.0185 0.17 2.109 cd 0.26 6.0151 0.0171 0.0152 7.0147 0.711 0.27 8.0138 0.0351 1.92 -1.16 1.31 Rn = "' -4.800 0.18 2.950 1.025 0.93 -1.0162 0.442 0.145 0.0102 0.20 5.14 2.0440 1.808 0.013 0.0070 0.0138 0.185 -0.0143 c1 "' -2.19 4.30 Rn c1 = "' -2.89 -0.93 -0.0527 cd 0.0172 0.0072 0.0429 300600.0074 0.082 0.443 0.29 8.0089 0.30 9.32 10.0172 0.923 0.0141 0.88 0.922 0.30 9.0151 -0.152 0.26 6.93 -0.31 10.0113 0.883 0.29 8.836 0.19 3.24 5.747 0.0096 0.0256 0.003 0.13 1.998 0.836 0.0086 0.89 0.0301 0.0105 0.92 -0.94 0.31 9.053 1.92 -0.12 2.539 0.94 -1.023 0.0096 0.341 0.Chapter 13: Tabulated Data 1.0094 0.19 4.24 5.0146 0.0113 0.30 -0.28 0.934 0.0337 1.879 0.105 0.0140 0.0126 0.24 5.89 0.0146 0.0103 0.0332 1.26 6.0162 82048-PT Fig.0084 0.17 3.0074 0.0225 0.24 5.0105 0.0174 82048-PT Fig.0090 0.0149 0.0119 0.634 O.0147 0.11 1.0086 0.0177 8.605 0.0157 0.0129 0.20 4.108 0.0475 1.0146 0.0212 0.33 1.089 -0.0086 0.27 7.0083 0.29 8.25 5.0243 0. = Rn 82048-PT 0.0183 0.31 Rn = ex -2.30 9.0260 0. 954 0. 0.0120 0.0088 0.0200 0.17 1.0211 1. 0126 0.604 0.140 0. = -2.0166 1.925 0.32 c.31 6.0083 0.0261 1.Dl08 0.0231 0.27 6.16 2.136 -0.059 1.87 -1.31 7.96 -2.0240 1. 12.392 0.258 0.0115 0.0147 -0.38 12.013 0.0081 0.0767 -3.80 Rn 60400.344 0.30 1.0134 0.27 5. c.0142 0.35 10.908 0.85 -0.89 -2.366 0.351 0.90 0.0116 0.28 5.0142 0.21 5.94 -1.0204 1.0126 0.36 c.720 0.90 -0.928 cd 0.83 0.70 2.061 0.0174 0.0329 0.888 0.604 0.11 1.329 0.90 -0.0144 0.0085 0.0171 0. 798 0. cd c.128 0.25 6.856 0.252 0.265 0.114 1.235 S209IB-PT Fig.19 1.0091 0.21 2.232 0.572 0.0137 0.091 0.23 3.80 11.619 0.179 0.21 2.34 8. c.29 7.357 0.0141 0.439 0. 722 0.18 3.0101 0.343 0.0087 0.37 Rn c.335 0.321 0.88 0.27 5.0262 = "' S2091A-PT Fig.26 4.17 1.201 0.0340 1.805 0.28 Rn = "' -4.31 0.26 7.0111 0.87 -1.88 -0.0764 cd -0.485 0.0181 0. 789 0.21 4.229 0.914 0.0078 0.10 1.18 3.0193 0.0067 0.0558 301300.22 2.278 0.0089 0.25 7.0151 0. cd -0.0460 0.249 0.92 0.87 0.89 -0.31 10.262 1. = " -3.18 1.37 12.0382 1.350 0.545 0.0106 0.92 -0.0263 0. 12.00 -2.435 0.049 0.0188 0.252 1.380 0.78 Rn 103100.0968 82055-PT Fig.0143 0.0398 1.28 8.25 6.0158 0.0099 0.0075 0.670 0.123 0.0331 1.25 5.246 0.0140 0.0146 0.16 2.31 7.0073 0.37 11.957 0.0459 149300.043 0.0118 0.0212 0.751 0.38 Rn -4.28 c.29 6.0089 O. " -2.27 8.466 O.79 Rn = 60800.006 0.919 0.25 3.002 0.120 0.344 0.0256 1. -0.974 0.35 9.63 2.289 0.19 4.24 6.26 7.95 -1. -0.1021 0.368 0.0141 0.22 2.16 2.36 10.0369 10. " -3.25 5.0399 0.16 3.0106 0.084 0. -0.84 -0.26 6.24 5.23 3.0230 0.0215 1.0088 0.338 0.869 0. "' -2.83 0.19 1.94 -1.0130 0.016 0.250 0.259 0.0104 0.0134 0.0077 0.12 1.30 Rn 0.26 4.0192 0.0113 0.053 0.33 8.0285 0.72 5. 716 0.0350 1.0113 0.36 11.880 0.079 0.36 cd 0.37 10.149 0.0217 1.0226 0. 732 0.0130 0.0209 1.44 -0.0120 0.0209 0.550 0.36 10. cd -0.113 0.225 0.0280 1.0174 0.20 4.0096 0.939 1.0172 0.0068 0.0182 0.011 0.865 0.36 12.0183 0.0077 0.24 6.558 0.84 0.640 0.162 -0.145 0.0137 0.25 4.997 0. "' -4.901 0. = "' -3.0109 0.0317 1.30 6.0143 0.19 1.19 3.134 0.312 0.0131 0.10 1.30 Rn = " -3.0072 0.188 0.808 0.272 0.0151 0.205 0.570 0.16 2.471 0.94 -0.137 0.050 0.981 0.27 5.0331 0.0286 1.73 3.376 Airfoils at Low Speeds -1.540 0.230 cd 0.19 2.145 0.0153 0.830 0. cd -0.84 9.034 0.632 0.22 5.32 8.881 0.0199 1.895 0.85 0.0180 0.0296 0.979 0.0096 0.14 1.90 -2.29 6.29 6.012 0.0078 0.253 0.31 7.0239 0.0122 -0.0073 0.27 8.0128 0.0234 0.795 0.0156 0.0298 1.974 1.0322 0.18 1.0217 -0.0225 1.0113 0.0351 Rn = 306300.0388 198400.21 2.21 5.27 6.0074 0.0108 0.447 0.0535 0.74 5.0335 0. 722 0.352 0.28 9.0322 0.048 0.0214 0.0121 0. cd 0.93 -3.0378 0.0077 0.0084 0.20 3.0254 1.791 0.96 -2.0197 0.89 0.541 0.0214 1.0284 0.844 0.30 9.088 0.0123 0.83 0.28 5.26 7.31 9.91 -2.548 0.349 0.0115 0.437 0.818 0.0073 0.0295 0.29 9.543 0.0165 0.099 0.0081 0.0307 0.16 2. 728 0.450 0.0111 0.04 72 101300.0301 0.1804 Rn = 152700.150 0.827 0.26 4.84 0.24 3.29 8.722 0.99 -1.640 0.38 11.21 4.210 0.127 0.0356 1.0101 0.0328 1. = "' -4.0141 0. -0.87 -1.0110 0.33 Rn = "' -2.650 0.0646 202000.0326 0.90 -3.050 cd 0.635 0.0070 0.464 0.0287 1.180 0.35 0.946 0.225 0.0149 0.0180 0.35 9.282 0. c.13 2.0210 0.0193 0.054 0.126 0.951 0.148 1.0260 0.887 0.0169 0.95 -2.89 -1.24 6.449 0.0120 0.057 0.33 8.0140 0.91 -0.38 Rn -4.0327 0.87 -1.645 0.35 9.019 0.34 8.747 0. 738 0.0149 0.0175 0.0083 0.609 0.0108 0.192 1.88 0. 78 8. 9.18 3.31 7.30 8.95 -1.0238 0.074 0.636 0.29 9.440 0. 0.0140 0.0275 0.453 0.0123 0.090 0.14 1.0313 1.0113 0.0187 1.065 1.012 0.980 0. cd -0.0128 0.079 0.238 0.0206 .0504 1.17 3.0096 0.914 0.89 0.88 0.0073 0.962 0.835 0.962 0.80 8. cd -0.692 0.0261 1.980 0.009 0.13 1.0438 99200.839 0.228 0.23 3.0154 0.07 -2.719 0. 771 0.95 -3.32 7.25 4. = " cd 0.31 9.540 0.0125 0.22 4.97 -1.279 0.0105 0.28 8.36 c.280 0.29 10.31 12.86 -0.79 6.0228 1.0203 1.0132 0.0208 0.758 0.0085 0.588 0.053 0.93 -2.23 3.0077 0.37 Rn " c.37 12.017 0.87 -1.667 0.098 0.19 3.464 0.0125 0.888 0.0683 201500. 12.0096 0.456 0.025 1.865 0.0201 0.0174 303900.0166 0.129 0.31 c.040 0.Oll9 0.37 11.0359 Rn = 304800.0265 1.0189 0.0116 0.0290 0.22 4.0166 0.237 0.0116 0.25 4.630 0. = "' c.35 10.278 1.25 6.239 0.0260 0.0105 0.294 0.90 -0.20 4.227 0.35 9.26 7.85 -0.22 5. cd -0.0074 0.979 1.80 c.469 0.526 0.0111 0.32 10. 792 0.31 7.37 11.330 0.704 0.40 -0.0084 0. c.0149 0.0102 0.091 0.030 1.079 0.154 0.0131 0.0168 0.96 -2. 099 1.29 5.765 0. 79 0.0134 0.866 0.0085 0.32 3.98 -0.122 0.190 0.45 10.29 0.403 0.1727 11.32 4.82 -1.0259 1.0787 1.0117 0.38 7.36 5.090 0.26 2.354 0.0165 0. a C1 Ca -4.324 0.0168 0.31 3.084 0.680 0.26 2.023 0.17 2.657 1.0104 0.145 0.39 6.0308 12.224 1.0101 1.0140 0.921 0.0604 1. = C> -3.0314 1.238 0.225 0.42 0.0244 1.0189 0.37 5.0144 0.574 0.619 1. 707 0.467 0.82 Rn = 59200.2234 Rn 100400.0223 0.0098 0.36 7.014 1.444 0.253 0.31 5.34 7.44 10.45 c. a C1 Ca ·4.0358 1.39 8.31 2.0202 1.055 0.235 0.411 0.373 1.124 0.926 1.523 1. 78 0.0227 5. a ·4.90 13.504 1.589 0. 7 4 -0.52 1.0247 0.744 0.951 0.28 2.438 0.27 3.34 1.0264 1.0240 0. Ca 0.84 ·0.0270 1.816 0.24 0.40 1.0148 0.28 0.o.666 0.15 1.0332 10.0100 0.0245 0.0170 0.517 0.30 1.518 0.567 0.0156 1.373 0.44 11.43 13.33 1.33 5.0315 1.409 0.0123 0.22 1.38 9.0175 1.36 1.334 0.115 0.0133 0. c.0139 0.26 2.34 8.0237 0.30 0. c.27 4.454 0.39 10.0282 0.0217 1.40 9.099 1.81 0.609 0.808 0.0254 0.82 1. 1.Chapter 13: Tabulated Data 9.79 0.42 11.85 Rn = 60900.0358 3.0249 0.896 0.0161 0.325 O.0139 0.90 -1.42 11. Fig.0244 0.0187 0.230 0.620 0.83 0.44 11.40 1.38 13.38 0.537 0.00 ·0.32 6.042 0.886 0.465 0.587 0.0150 0.0297 1.414 1.597 0.0565 1.681 0.0233 0.0218 1.0251 1.890 c" c.0140 0.23 0.040 0.0159 0.560 0.001 0.286 1.90 -3.35 -1.1052 1.81 Rn 99600.618 0.35 6.0371 1.0163 1.452 0.088 0.38 1.0115 0.32 7.0138 0.379 0.90 1.0231 0. 79 ·1.0178 0.0174 0.670 1.42 12.498 0.29 5.0227 10.89 ·3.0110 0.320 1. 77 -1.683 0.0121 0.801 0.0256 0.0285 1.80 ·2.0164 0.83 Rn = 99800.0206 1.0110 0.38 0.073 0.0113 0.456 0. -0.35 0.325 1.21 1.163 0.0238 7.0135 0.0136 0.77 4.058 1.030 1.91 0.0268 4.0171 0.34 6.0128 0. 79 1.0119 0.0157 -0.42 8.0172 1.0148 1.01 ·0.632 0.0164 0.0142 1.33 5.93 0.19 3.82 -1.0366 0.0240 0.330 0.285 0.42 10. a -4.492 1.42 0.76 0.0189 0.40 1. a -2.90 1.80 ·0.88 -1.96 -0.34 7.0375 S2091B-PT Fig.621 0.24 0.85 7.0216 0.0235 0.503 0.764 0.0358 1.079 0.42 11.0186 1.267 0.0483 9.0443 7.43 11.360 0.104 0.43 10.522 0.44 12.0096 0.40 13.468 1.36 1.0112 0.41 11.805 0.38 8.0246 0.39 8.0967 0.0188 2.0132 0.0465 1.0136 0.128 0.892 0.545 0.0289 1.14 0.451 0.79 -0. a C1 Ca -3.395 0.38 11.206 0.992 0.0427 0.416 Ca 0.25 6.031 0.0136 0.0145 0.381 0.85 -0.526 0.014 1. 715 0.32 1.0129 0.0167 0.318 0.0158 0. c" c" 0.003 0.0155 1.013 0.552 0.28 4.559 0.0232 1.24 2.736 0.90 -3.84 ·1.42 12.0553 1.184 0.266 0.25 0.0408 1.378 0.0282 0.632 0.495 0.1330 S2091B-PT a C1 Ca 0.0165 0.0628 Rn 200000.0219 0.40 7.0131 0.091 0.0221 0.997 0.23 1.0418 1.41 0.81 -1.558 Ca 0.180 0.0174 -3.277 0.072 0.87 -2.280 0.0085 0.2719 Rn= 199300.84 1.915 0. Rn = a ·3.0254 1.213 0.27 c.31 4.0479 0.38 9.965 0.23 0.39 1.Q208 1. 721 0.838 0.91 ·2.0213 377 .95 -2.839 0.0295 1.0196 ·0.0411 Rn 100500.0339 0.27 3.553 0.171 0.0259 11.79 0.0172 ·1.29 3.0353 10.22 0.0174 0.491 0.0339 1.468 0.82 ·0.33 8.335 0.0357 9.0280 7.78 ·1.36 4.892 0.81 .256 0.483 1.645 0.0285 1.0463 Rn = 303900.24 5.0210 1.32 3.705 0.0228 0.153 1.863 0.940 0.4 76 0.0346 1.0609 Rn = 199700.88 0.0182 0.131 0.86 0.38 1.560 0.0133 0. c" a -4.40 7.0115 0.0250 1.26 3.0213 1.79 ·0. 793 0.222 0.421 0.0486 302700.0182 0. 0.0236 3.477 0.75 ·1.0158 1.40 9.37 8.0861 S2091B-PT Fig.0256 1.462 0.0232 0.565 1.0237 0.41 8.36 6.20 2. Ca 0.0288 ·2.0259 0.0105 0.78 ·2.370 0.26 4.0231 1.0242 1.0182 0.23 1.0309. -3.86 ·2. 733 0.30 4.0388 11. 12.0130 0.567 0.560 1.0657 1.41 c.35 4.538 0.844 0.0203 0.392 0.54 1.76 ·1.0329 1.2596 Rn = 146200.85 0.38 9.334 1.90 ·2.182 1.142 0.0150 0.0165 0.0233 1.31 5.565 0. 12.382 1. = = a ·4.187 0.0280 6.669 0.0102 0.38 6.43 9.30 1. = a -4.577 0.0257 -3.640 0.31 c.31 6.0227 0.90 10.0234 0.0464 1.925 1.909 0.457 0.31 6.430 0.0988 S3010·PT Fig.586 0.0154 0.42 10.38 C1 0. 12.80 ·2.543 1.0240 0.247 1.258 1.482 1.021 0.23 2.414 0.39 9.26 3.88 -3.29 0.310 0. 797 0.0313 1.89 -0.24 3.851 0.0135 0.22 4.42 12.0198 4.0207 0.0196 0. {)( c.0101 0.0279 8.0190 -1. 789 0. c.0330 0.0316 5.34 5.316 0.966 1.0282 1.516 0.191 1.304 0.89 -2.0244 0. C1 0.594 0.41 10.0125 0.1061 Rn = 306400.688 0.0323 1.35 7.40 14.606 0. 719 0.0180 0.86 ·2. 12.35 9.69 0.84 0.414 0.42 12.130 1.414 1.35 -3.832 0.82 0.474 0. 73 0.956 1. 713 0.0180 1.0150 0.43 9.38 10.0510 13.376 0.329 0.37 7.51 1.1139 298500.0201 0.0223 0.18 0.437 0.378 0.44 11.131 0.640 0.0212 1.468 0.133 0.588 1.0207 0.0138 0.0212 0.0245 0.29 2.39 Rn = c.30 5.0249 0. a ·4.413 0.38 8. 0296 0.432 0.948 0.34 1.185 0.0085 0.693 0. Q -3.019 0.0183 0.24 7.28 6.92 -0.96 -2.972 0.26 5.33 9.447 0.498 0.0181 0.0214 0.0122 0.24 4.148 0.0421 Rn = 152700.0223 0.Oll3 0.0236 0.29 cd 0.0073 0.992 0.948 0.0118 0.17 3.0213 0.0264 0.0429 0.0087 0.0152 0.21 3.144 0.293 0.135 0.015 cd 0.34 11.20 2.665 0.92 -2.18 3.34 10.0227 0.0088 0.0277 0.130 -0.0120 0.84 1.0079 0.20 4.66 2.35 Rn = "' -3.0327 1.0171 0.23 3.774 0.015 0.189 0.34 12.30 7. -0.89 -3.676 0.0124 0.0080 0.18 3.992 1.0080 0.0127 0.551 0.28 6.0120 0. 786 0.17 1.0139 6.293 cd 0. c" -0.35 11.27 5.28 8.190 -0. " -0.354 -4.92 -1.92 -0.0124 0.0133 0.15 2.0237 8.166 0.0163 -0.075 0.525 0.94 -0.31 8.0202 1.365 0.0263 1.85 1. c. 764 0.099 0.29 9.081 1.496 0.40 -0.0193 9.89 -1.086 1.0365 c" 0.35 11.0115 0.396 0.23 6.98 -2.17 1. c" " -0.24 6.938 1.11 1.928 0.Dl05 0.0160 83016-PT 0.94 -1.399 0.27 8.857 203900.0564 14.22 5.0125 0.32 9.99 -3.27 8.98 -2.935 1.0322 0.0601 103300.219 -0.17 c.0566 151500.829 0.0114 O.20 4.0127 0.0153 0.651 0.96 -2.18 3.0107 O.32 1.92 0.480 0.20 4.367 cd 0.0230 1. 777 0.495 0.307 0.394 0.0162 0.95 -1.0253 0.27 8. 0.0124 5.286 0.88 0.0186 . 12.0123 0.0480 Rn = 147400.95 -0.0118 0.24 6.30 7.99 -3.28 Rn = " -3.052 1.0345 12. 0.22 4. c" 0.169 0.051 0.33 Rn = " -0.98 -2.0090 0.29 8.22 5. c.0091 0.196 0.306 -4.073 1.011 0.0109 0.828 0.36 Rn = Q c.0141 0.16 2.393 0.14 1.0227 1.664 0.023 0.160 0.0214 0.0117 0. c.25 6.17 2.34 10.083 0.0397 1.25 7.0143 0.19 2.294 0.22 3.0112 -0.0284 0.31 9.0152 0.32 1.0217 0.0241 0.0339 0.90 0.14 1.27 7. 753 0.626 0.023 1.92 -0.0126 -2.660 0.0098 0.0322 1.90 0.0084 0.104 c" Rn = -3.33 11. 0.34 1.021 1.590 0. " -0.0314 83014-PT Fig.0253 11.0131 0.598 0.470 0.045 0. c.0205 1.0244 0.105 -0.31 8.15 2.512 0.851 0.0257 0.71 0.0219 0.207 -3.120 0.866 0.0099 0.191 0.569 0.0156 1.24 6.26 5.608 0.0158 0.33 c.084 1.14 2.182 1.0111 0.179 0.026 0.071 0.89 -0.36 11.0119 0.26 6.105 0. " -3.22 5.04 77 " -0.381 0.0295 0.0140 0.0208 0.811 0.836 0.73 5.091 0. -0.28 9.390 0.066 0.21 5.0121 0.129 1.0280 0.30 0.0077 0.0359 0.96 -1.0104 0.0139 0.25 7.29 7.24 5.0086 0.0094 0.0174 0.0401 0.866 0.091 1.34 Rn = " -3.0140 0.0302 1.0332 11.20 4.078 -3.013 0.138 0.0193 0.18 3.0237 0.0144 0.0179 0.0232 0.0182 1.208 0.007 1.90 -1.27 6.061 0.96 -3.0115 0.341 -4.118 0.15 2.303 0.750 0.27 8.0096 0.91 -0.67 0.0105 -1.689 o.0081 0. 7 45 0.378 Airfoils at Low Speeds 7.340 0.084 0.23 6.0133 0.159 0.102 -0.0283 0.217 0.0183 0.31 c" Rn = c.481 0.87 Rn = 62000.Ql05 0.554 O.20 0.180 0.29 9.205 0. " -0.0087 -0. 772 0.2484 cd -0.0181 0.0113 0.0216 0.0099 3.925 0. " -0.314 0.19 3.90 0.25 7.819 0.90 0.31 1.837 0.32 11.065 0.0141 0.132 0.0738 Rn = 301000.148 0.89 0.85 11.0230 1.16 2.0309 0.577 0.98 -2.916 0.31 10.0120 0.0099 0.893 0.008 Rn = 59100.245 -0.21 2.0211 0.86 0.99 -3.87 -0.0180 1.0255 0. c.88 -0.0092 0.090 0.94 -1.0143 0.686 0.0220 0.92 0.23 3.224 0.13 1.0107 0. 732 0.29 7.14 1.30 7.22 5.0100 0.96 -2.95 -0.90 Rn = 58800.32 1.197 0. c.116 0.0135 -0.25 6.571 0.0155 0.30 10.0155 0.192 0.0223 0.0238 -0.89 0.32 9. c.34 10.124 304800.81 8.004 0.32 11.242 0.147 0.0088 0.159 0.0132 0.356 0.31 8.0259 0.78 0.126 0.661 0.060 0.0212 0.0171 0.128 -0.91 -0.044 0.0135 0.88 0.010 1.0143 -0.0174 0.0165 0.28 9.28 8.17 1.0157 0.166 0.598 0.0082 0.394 0.0137 0.171 0.26 5.0087 0.301 0.020 0.057 0.0159 8.872 0.0109 0.31 Fig.0175 0.15 2.005 0.524 0.134 cd 0.147 0.859 0.28 6.0196 -0.592 0.0230 0.158 99900.19 2.0176 0.565 0.0181 0.303 -4.0137 0. 12.88 0. " -0. 12.0238 -2.31 9.0155 0.32 10.0211 0.34 Rn = -0.293 0.17 3.158 0.0180 -0. cd -2.687 0.042 0.0130 0.0216 0. c.974 0.13 1.91 -2. 0.85 0.93 -2.0167 0.99 0.0316 0.0281 0.268 0.23 4.34 10.32 -0.923 1.0194 0.0162 0.086 0.0137 0.0185 0.96 -1.493 0.369 0.25 4.0158 0.0282 1.0116 0.479 0.175 c" 0.0180 0.92 -2.0174 0.167 1.777 0.0075 0.127 -0.0193 0.747 0.060 0.650 0.94 -1.168 0.19 4.0382 0.41 0.0434 303700.19 3.167 1.35 0.0189 0.685 0.416 0.277 0. 923 0.0256 9.13 1.34 1.24 4.188 0.235 0.85 14.0493 200800.0601 0.857 0.0098 0.0326 0.90 0.30 9.34 9. c.0137 0.84 0.153 1.947 1.522 0.0193 0.922 0.0115 0.0098 0.33 9.0096 0.19 4.33 -0.069 99500.0314 83021A-PT Fig.31 10.0094 2.950 1.0286 Rn = 208900.923 1.0256 0.19 4.959 1.822 0.28 7.0175 0.25 0. -0.747 0.0115 0.0247 0.93 -0.011 0.0177 0.0102 0. " -0.21 5.033 0.32 8. c.0305 10.27 6.837 0.93 -1.086 0.87 0.0244 1.26 7.148 0.93 -2.0402 cd -3. c.35 10.0108 0. 33 Rn = 0: -3.008 1.338 0. = C< -2.548 0.0116 0. 785 0.34 10.25 6.93 -0.0091 0.90 -0. 730 0.86 0.0262 0.026 0.05 1.0171 1.972 1.147 0.93 0.0362 200600.28 6.0324 1.0126 0.0101 0.0170 0.28 6.458 0.0127 0.0120 0.523 0.17 1.931 0.0106 0.27 7.012 0. c1 0.393 0.21 3.0106 0.0100 0.0149 0.554 0.801 0.725 0.0096 0. 750 0.144 -0.0115 0.97 -3.352 0.36 Rn = " -2.33 10.30 10.0079 0.0132 0.0203 1.27 7.31 7.111 1.0195 0.0090 0.22 5.0204 S302IB-PT Fig.30 7.0209 0.169 1.89 -1.0172 0.100 1.937 1.34 11.193 -4.0177 0.0539 cd 0.0204 0.0115 0.065 0.21 3.896 0.036 0.441 0.039 0.0129 0.0143 0.895 cd 0.94 cd cd 0.0079 0.87 0.0275 0.0115 0.97 -1.26 7.0202 0.001 0.100 1.99 -1.0421 300300.25 7.0132 0.0311 0.0126 0.502 0.29 9.35 11.812 1.131 1.28 7.0222 1.16 1.803 0.26 7.25 5.0174 0.170 0.24 6.86 0.87 0.512 0.886 0.29 8.0186 0.0341 0.152 0.95 -2.946 0.067 1. 12.22 4.22 3.0111 0.93 -0.19 4.019 1.114 0.267 0.90 -1.0336 0.94 -2.947 1.15 1.24 4. c1 c1 " -0.0210 1.0113 0.31 9. c1 0.0170 0.0158 0.0225 1.33 10.376 0.056 S406IA-PT Fig.652 0.099 cd 0.0134 0.001 0.071 0.0109 0.326 0.34 9.279 0.0196 0.30 7.043 0.0143 0.0145 0.Chapter 13: Tabulated Data -1.0291 201300.28 8.08 2.0377 11.698 0.0077 0.240 0.155 0.0207 c1 cd -0.364 0.0195 0.0117 0.0127 0.868 0.131 303700.32 9. <> -3.16 2.234 0.32 9.24 4.29 9.31 8.207 0. 705 0.046 0.154 0. c1 -0.15 1.056 -3.24 5.552 0.0297 304800.0135 0.0265 1.0171 0.27 6.810 0.0250 1.0130 0.0174 0.26 6.124 0.21 4.557 0.92 .93 Rn 60400.24 5.995 1. -0.22 4.27 7.33 10.0264 0.0241 0.0095 0.87 -0.366 0.11 1.15 1.0108 0.0133 1.32 9.0085 0.048 0.89 -0.0128 0. 12.0101 0.20 5.87 0.31 8.142 0.0413 0.609 0.843 0.16 3.0099 0.0095 0.092 0.615 0.872 0.0189 0.104 0.0103 0.35 11.553 0.17 2.0314 0.055 1.0176 0.0324 0.0296 0.983 0.91 0.36 1.0227 0. 12.37 Rn = <> -1.25 5.0124 0.91 -0.080 1.0131 0.236 0.0102 0.23 4.26 5.36 12.92 -0.19 3.177 0.30 9.91 -1.22 5.33 10.0199 0. 700 0.0374 0.34 cd 0.223 1.210 0.0102 0.0095 0.0113 0.0146 0.591 0.104 0.174 0.03 -2.0262 1.0093 0.252 0.17 2.91 -1.076 0.34 0.144 0.20 2.0143 0. 0.15 1.19 4.898 0.32 8.0211 0. = 0: -2.21 4.250 0.846 0.092 0.0171 0.21 3.975 0.0157 0.93 -1.635 0.34 Rn = " ·2.0239 0.0166 1.0377 149300.92 -1. 703 0.0223 0.042 1.0313 0.415 0.818 0.21 4.0121 0.0294 cd 0.296 0.0108 0.291 0.23 5.510 0. 790 0.31 8.34 11.445 0.99 -1.984 1.448 0.23 4.0148 0.0183 0.633 0.0432 200400.87 0.0204 0.259 0.374 0.810 0.24 5.85 0.31 Rn = " -2.603 0.639 0.36 Rn = c1 c1 " -0.16 1.32 10.893 0.05 7 0.22 4.078 0.751 0.35 Rn = <> -3.0083 0.33 11.933 0.069 0.692 0.29 9.0231 2.0095 0.0203 1.0339 0.27 6.17 3.29 8.0144 0.21 5.245 0.19 2.0119 0.426 0.34 10.027 0.0245 0.30 11.0138 0.0149 0.14 2.95 -0.029 0.17 3.460 0.33 9.0256 1.098 c1 S4061B-PT Fig.19 3.0230 0.056 0.0100 0.0163 0.0116 0.270 0.0168 0.0117 0.0268 0.903 0.0222 0.0294 0.0116 0. c1 -0.468 0.362 0.091 0.27 8. 721 0.069 1.056 0.419 0.34 9.0163 0.090 0.18 2.0283 1.0147 1.20 cd -0.452 0.25 6.143 0.318 0.00 -0.0165 0.0233 1.29 8.0186 0. 789 0.34 10.0087 0.624 0.14 3.18 2.0105 0.526 0.264 0.0250 0.0086 0.26 6.0209 0.2.18 3.978 1.28 7.010 1.87 -0.26 5.542 0.98 0.0164 1.0088 0.604 0.0225 1.24 4.0135 0.0284 Rn = 198900.619 0.890 0.0113 0.700 0.35 10.0140 0.457 0.145 0.169 cd 0.0299 0.28 6.911 0.0375 0.33 -0.96 -0.027 0.746 0.0165 0.87 -0.24 cd -0.545 0.31 9.0311 0.96 -1.0297 0. " -3.128 0.33 c1 -0.92 -2.84 0.0163 0. cl = -2.31 c1 -0.222 0.35 l1.0406 150800.0209 0.0157 0.92 Rn 99300.048 0.88 1.17 1.132 0.653 0.0206 1.0209 1.12 1.0253 1.0180 0.008 0.0177 0.0098 0. c1 cd 0.0142 0.0105 0. <> -4.007 0.0218 0.0442 -0.055 0.95 -1.25 6.089 1.0084 0.611 0.269 0.0149 0.0139 0.33 9.082 0.0252 0.0104 0. 726 0.28 9.0197 0.174 0.99 -2.34 Rn = -0.29 7.01b7 0.869 0.788 0.397 0.0191 1.0187 0.25 6.287 0.16 2.88 0.0143 0.31 7.026 0.059 0.356 0.18 1.0198 0.0113 0.348 0.979 0.89 -0.26 5.18 1.92 0.0098 0.0160 0.0195 0.31 8.30 10.0336 1.20 2.30 8.0107 0.0129 0.0212 0.32 10.0109 0.31 10. 722 0.0179 0.24 7 0.008 0.639 0.0171 0.398 0.0222 0.94 Rn 60200.078 0.048 0.0072 0.130 99800.817 0.27 8.20 3.20 3.131 0.194 0.791 0.0245 1.92 -1.135 0.18 2.0458 100000.31 Rn = <> -2.0193 0.944 0.0095 0.27 7.082 0.21 4.0081 0.22 5.500 0.215 0.266 0.16 2.0148 0.253 cd 0.33 8.0195 1.0344 0.0312 1.0127 0.23 6.110 0.176 0.90 -0.142 0.0427 cd Rn = -0.057 1.0310 0.0109 379 .19 3.34 10.0089 0.94 0. 0148 1.0111 0.84 14.234 0.188 0.26 0.90 -1.33 1.0148 0.0114 0.0192 1.0127 0.0141 1.0098 0.999 0.0113 0.28 7.83 9.253 0.0150 0.0117 0.90 -1.682 0.34 10.050 0.229 0.80 9.35 11.233 0. cd 0.0308 1.502 0.25 4.85 14.970 1.0342 Rn = 299700.0288 1.545 0.233 0.90 0.024 0.0257 0.0155 0. -3.058 0.89 -1.810 0.88 -1.992 1.0111 0.0466 Rn = 200500.0134 0.23 3. cd 0.34 9.84 11.32 8.355 0.170 1.73 6.139 0.32 9.0209 0.0252 0.30 6.34 0.21 4.22 4.39 -0.145 1.513 0.0143 0.233 0.35 10.0200 0.706 0. cd 0.25 4.0127 0.88 0.33 8.0209 0.33 Rn = "' -3. -2.41 -0.21 2.26 4.0098 1.529 0.993 0.0108 0.35 10.0108 0.29 6.0114 0.239 0.21 3.86 -0.051 0.144 0.39 0.69 3.86 12.0201 0.93 -2.0146 0.442 0. " -2. c.0137 1.84 -0.362 0.71 3.0293 0.822 0.0151 0.556 0.391 0.201 0.0121 0.0178 0.0170 1.35 Rn c. 12. "' -3.0093 0.204 0.27 5.0227 1.0110 0.233 0.32 8.146 0.945 0.94 -2. " cd 0.0173 1.395 0.183 0.33 9.85 12. 71 5.0096 0.0148 1. 77 6.22 3. 726 0.0172 0.407 0.23 4.0255 0.253 0.88 -0.251 0.131 0.524 0.65 2.948 0.99 Rn = 59300.058 0.380 Airfoils at Low Speeds S4061B-PT Fig. c.29 6. 751 0.0144 0. 79 8.0233 1.0117 0.0191 1.0117 0. cd 0.35 c.33 c.0157 1.0108 0.34 10.34 10.25 6.238 0.18 1.26 5. cd 0. c.34 9.16 1.0122 0.35 10.15 1.0284 150200.0202 0.0437 1.28 7.74 6.31 7.96 Rn 150800.239 0. " -2. c.222 0. c.77 6.78 8.0291 1.050 0.241 0.362 0.34 9.229 0.14 1.28 5.33 8.39 -0.197 0.225 0.0274 0.016 0.0247 0.663 0.0088 0.80 9.0220 0.36 11.80 9.35 Rn = "' -4.87 0.178 0.35 Rn = " c.109 0.38 0.0121 0.86 12.066 0.016 0.039 0.97 Rn = 150200.91 -2.0147 1.241 0.39 0.0202 1.74 5.0447 304900.83 9.0273 0.305 0.199 0.086 0.0198 1.0282 0.0184 1.0415 0.0147 0. = " -1.0226 0.0145 1.197 0.0341 1.15 1.0361 0.355 0. 0.0214 1.0170 0. "' -2.0183 0.15 1.204 0.95 -2.36 12.0148 0.93 0.0338 1.86 -0.27 7.87 0.0466 151000.810 0.71 3.182 0.042 0.009 0.36 Rn = "' -3.34 12.663 0.0108 0.0217 1.20 3.848 -3.39 0.69 3.207 0.0155 0.0143 0.337 0.853 0.501 0.0218 1.0120 0.475 0.031 0.34 c. -0.0516 1. -2.121 0.0425 301300.90 -1.158 1.21 3.235 1.0103 0. 12.0155 0.93 -2. Rn = -0.239 cd 0.0384 150200.0105 0.122 0.90 -2. 79 8.0226 1.0172 0.092 0.0125 1. "' c.71 3.92 -1.227 0.179 0. " c.0104 0.223 0.85 11.30 7.23 4.77 6.0131 0.953 1.16 2. 12.600 0.0300 1.855 0.34 c.0256 0.992 0.0197 0.85 0.0170 1.731 0.970 0.38 0.88 0.0155 1.0191 1.0269 0.0098 0.169 0.807 0.0144 0.0135 0.107 1.13 1.085 1. 12.0186 0.0223 0.019 0.23 3.197 0.0172 1.0281 1.0135 0.86 0. cd 0.0144 0.948 1.34 Rn = c. 761 0.812 0. cd -3.0117 0.243 0.68 3.619 0.95 Rn = 60900.86 -0.0095 0.67 3.0325 1.0270 1.0308 1.865 0.32 10.82 0.0289 0.826 0.36 " c.026 0.0307 1.85 12.345 0.193 0.950 1.100 0.85 12.0086 0.32 10.0136 0.74 6.74 5.36 Rn = " -3.0125 0.430 0.079 0.208 0.94 -2.0146 0.0184 0.86 0.092 0.17 1.0157 0.324 0.25 6.0120 0. 0.0105 0.0172 1. c.27 5. cd -0.0255 cd 0.0292 0.207 0.0274 1.556 0.0451 .205 1.0110 0.0342 S4061B-PT Fig.049 0.31 7.39 0.0084 0.829 0.0624 299700.0359 1.34 9.0172 0.179 0.114 0.0194 0. -2.179 cd 0. cd -0.159 0.79 8.340 0. cd -0.513 0.35 Rn = "' cd 0.0227 1.33 9.829 0.32 9.17 2.650 0.0236 0.147 0.0114 0.34 10.0156 0. " -3.68 2.0134 0.0143 0.072 0.0425 S4061B-PT Fig.217 0.35 11.436 0.36 12.27 7.071 1.0143 0.0198 1.0141 0.0621 201600.168 0.355 0.0140 0.531 0.189 0.0170 1.0192 0.0153 0.853 0.80 9.0214 1.0134 0.27 7.0140 -1. c.654 0.84 0.243 0.0088 0.27 6.051 0.0294 0.335 0.72 5.189 0.916 0.14 0.432 0.19 2.0088 0.0291 0.235 0.83 9.22 4.39 0.19 2.20 3.058 0.810 0.181 = 98900.25 6.650 0.079 1.85 11.0466 Rn = 151000.853 0.0203 0.0206 0.34 10.0134 0.0233 0.663 0.0135 0. cd -2.402 0.0308 12.117 0.28 6.0107 0.852 0.0434 100700.207 0.0099 .0222 0.0185 0.884 0.178 1.65 2.419 0.29 6.25 4.0197 0.0139 0.21 3.66 2.92 -2.74 6.233 0.395 0.233 0.36 Rn = " -2.84 0.0118 0.627 0.048 0.157 0.203 0.079 0.115 1.0228 0.31 7.085 0.27 7.18 1.948 0.0134 0.115 0.0346 0.0150 1.0493 1.0087 0.0313 0.513 0.23 4.090 0.21 4.23 3.0096 0.524 0.095 0.0214 10.21 4.060 0.93 -1.619 0. c.682 0.0259 0.0139 0.0147 0.0344 1.029 -1.33 8.0092 0.31 7.35 10.39 -0.0116 0.246 0.027 1.15 1.84 11.707 0.0603 150200.550 0.28 7.0323 0.32 9.114 0. 753 0.93 -1.0270 0.0134 0.541 0.337 0. 708 0.88 -1.231 0.923 0.0253 1.627 0.0170 0.0141 0.93 -1.0122 1.749 0.989 0.0103 0.707 0.0229 0.241 0.0162 0.86 0.558 0.0270 0.205 0.17 4.91 0.232 84062-PT Fig.21 2.0219 1.85 1.19 1. 94 -0.29 0.0097 0.500 0.977 0.0094 0.845 0.18 2.98 -4.19 3.91 -1.0265 0.23 5.0173 0.0310 0.093 1.869 0.832 0.31 c1 cd -0.025 0.838 0.37 11.98 -0.0226 1.451 0.0404 84233-PT Fig.82 9.111 0.120 0.229 0.0204 0.98 -2.552 0.224 0.17 3.0112 0.87 0.0125 0.306 0.35 11.013 0.0210 1.0126 0.20 4.105 0.0271 1.24 5.639 0.220 0. = "' -2.0139 0.234 0.0156 0.17 2.0349 1.86 0. 12.0212 0.92 -1.11 1.0191 0.32 c1 -0.32 9.102 Rn 60200.88 -0.0111 6.0162 0.0162 1.0122 0.16 1.0241 0.170 -0.049 0.35 c1 cd -0.33 10.0253 -0.27 7.34 9.064 0.815 0.0121 0.734 0.36 -0.27 7. 724 0.226 0.37 c1 cd -0.342 0.076 0.0102 0. a -5.0166 0.921 0.34 10.163 0.0167 1.27 6.0123 0.056 0.23 5.213 0.0277 0.028 0.0125 0.33 11.88 1.136 0.403 0.595 0.116 0.35 10.101 Rn = 60300.90 -0.137 8D2030-PT Fig.202 0.629 0.0206 1.170 84233-PT Fig.220 0.0171 0.0113 0.0162 0.0204 0. 77 6.0139 0.162 1.128 0.308 0.0197 1.0242 1.409 0.Chapter 13: Tabulated Data 5.824 0.596 0.0193 0.225 0.431 0.233 0.0121 0.411 0.0182 -0.099 0.0418 0.0123 0.0221 1.162 cd 0.0182 0.23 5.84 12.0211 -0.33 9.0115 0.27 6.0258 298900.0183 1.92 -2.16 c1 cd 0.92 -1.111 0.35 9.0213 1.0114 0.964 0.528 0.917 1.18 2.18 2.97 -3.0149 0.021 0.613 0.159 0.945 0.100 Rn 106400.32 9.138 0.0114 0.253 0.0198 -0.0127 0.0292 0.217 0.0104 0.0166 -0.313 0.423 0.0155 -0.0097 0. c1 -0.0336 0.124 0.047 0.0217 1.92 -1.32 11.22 4.25 5.30 7.97 -3.0160 0.265 0.19 3.0100 0.0121 0.017 1.20 4.0128 0.0109 0.34 10.16 3.0135 0.0112 0.142 1.149 0.43 -2.0133 4.158 0.16 1.25 5.635 0. Rn = a c1 -5.140 0.15 1.0281 305100.0173 1.20 4.0189 0.0300 0.094 0.0127 0.0121 0.32 10.520 0.313 0.20 3.25 6.967 1.0181 0.98 -1.90 -0. "' -3.32 9.0123 0.92 -1. 730 0.0263 0.075 0.0147 0.175 0.011 0.0195 1.0153 0.13 1.23 5. "' -4.0137 0.93 -2.292 0.0106 0.89 -0.538 0.057 0.0294 0.94 -3.0139 0.21 3.622 0.065 0.35 10.0183 0.284 0.202 0.0264 0.0260 0.0134 84180-PT Fig.0165 0.36 0.30 7.0212 0.27 6.29 7.354 0.214 0.0408 1.935 0.849 0.30 8.962 0.255 0.35 13.289 -0.34 11. 12.033 0.20 3.939 0.0368 1.15 2.024 0.920 0.94 -2.36 10.0143 0.0162 0.Ql3 0.640 0. 12.0311 0.0139 0.0109 0.1460 201800.833 1.31 1. c1 -0.0117 0.22 4.24 4.011 0.119 0.25 6.35 10.862 0.124 0.14 1.074 0.0178 0.116 0.0118 0.0149 0.32 9.30 7.18 3.0148 381 .234 0.37 12.0151 0.27 7.0292 0.0136 0.0197 1.0123 0.0210 0.0223 1.36 11.474 0.97 -0.27 6.0190 0.0118 0.94 -2.0249 -0.713 0.18 3.99 -2.93 -0.0154 1.269 -0.189 0.88 -0.92 -2.0130 0.88 0.24 5. 754 0.0328 1.056 1.32 8.148 0.91 -0.006 cd 0.175 1.31 8.418 0.0152 0.97 -4.0381 Rn = 101300.17 1.28 7.663 0.34 9.691 0.912 1.144 0.33 c1 cd -0.0293 0.0177 0.94 -0.014 0.15 1.34 12.042 0.199 0.011 0.172 0.16 1. a -2.0109 0.005 0.318 0.020 0.93 -2.16 2.0116 0.428 0.86 1.0111 0.98 0. 766 0.35 11.32 8.0418 c1 -0.0147 0.0121 0.36 Rn = a 1.23 5.0262 0.552 0.019 0.30 8.28 7.257 0.89 0.0172 0.643 0.0162 0.171 0.186 cd 0.86 0.196 1. c1 0.0266 0. c1 cd -5.41 0.0173 1.89 -0.0485 101700.745 0.0188 1.92 -1.968 0.28 8.28 7.05 -3.0375 0.0180 0.415 0.0199 1.97 -1.005 0.0092 0.080 0.0188 0.32 Rn = "' -2.0199 150200.0114 0.23 4.0156 0.23 4.0282 0.97 -4.098 0.0975 = 304700.0647 0.0192 -0.00 -4.398 0.30 8.0213 0.219 0.220 0. " c1 cd 0.929 1.0256 1.176 0.20 3. "' -2.0137 0.18 3.0130 0.651 0.0172 0.90 0.201 0.321 0.91 0.91 -1.0205 0.0242 1.0301 1.17 2.91 -1.0112 0.0290 c1 = -4. 789 0.26 7.021 1.175 1.0164 -0.25 5. 787 0.14 1.146 1.0267 1.221 0.0112 0.28 6.94 -3.222 0.35 -0.32 9.29 7.306 0.0543 1.221 0.0167 0.36 Rn = " -6. a Rn -0.0284 0.234 0.30 8.373 0.197 0.0152 0.826 0.0144 0.36 Rn = " -3.29 8.513 0.26 6.0248 1.0124 0.064 0.0452 1.174 -0.68 3.33 10.21 3.173 0.0290 0.024 0.0203 0.0154 1.0263 1.101 0.145 0.0182 0.0629 cd -2.0162 0.0154 0.290 0.31 8.0108 0.17 2.752 0.0182 0.152 0.26 6.0264 0.526 0.191 1.881 0.0212 8.057 0.24 5.87 0.076 0.0292 0.34 9.0164 0.110 0.460 0.624 0.30 9.0122 0.35 10.95 -2.108 0.88 -0.0257 0.156 1.0160 0.0398 152100.0299 0.331 0.0179 0.0180 0.170 1.189 0.0370 1.0267 200200.055 0.616 0.23 4.33 9.34 12.100 1.741 0. 770 0.0161 0.87 0.0193 0.15 1.538 0.35 10.542 0.0119 0.13 1.35 Rn c1 cd O. = " -3.87 0.002 0.062 1.17 2.433 0. 12.0185 0.234 0.95 -1.95 -3.0316 1.33 Rn cd cd -6. = a -5.30 8.0210 0.30 7.019 0.22 cd -0.90 -0.0125 0.0155 0.0464 Rn = 202200.0149 0.0100 0.144 0.92 -0.27 6.077 0.0097 0.0185 1.104 Rn = 101000.25 5.92 -1.21 4.18 2.117 0.0132 0.34 10.0162 1.88 0.0298 0.0356 Rn = 127300. 26 7.94 -1.0121 0.27 7.273 0.137 0.036 0.813 0.248 0.0098 0.0232 0. = "' -6.013 1.0289 9.0075 0.095 0.31 Rn = c.0429 300700.089 -2.20 4.93 -0.90 0.21 3.742 0.29 8.31 9.0237 0.0312 0.0289 0.19 2.0135 0.325 O.291 0.0089 0. -0.405 0. "' -3.0141 0.25 6. " c.0176 6.049 0. -3.265 0.0124 0.95 -0.057 0.0162 SD5060-PT c.611 0. cd 0.92 -0.092 0.26 7.0112 0.004 99800.0272 1.956 0.20 3.15 2.39 0.29 8.593 0.29 0.25 7.0154 0.603 0.0151 0.34 1.0095 0.0069 0.20 2.0249 0.821 cd 0.996 0.19 4.0193 -1.23 3.28 7.92 -0.148 0.0374 Rn =·155700.Ql03 0.22 0.399 0.475 0.929 0.191 0.009 0.0327 1.0457 0.25 6.846 0.34 1.484 0.0217 0.081 0.679 0.29 8.85 0.24 5.0212 0.95 -2.128 -2.566 0.538 0.89 0.20 4.27 6.34 1.0089 0.20 2.429 0.008 1.426 0. 704 0.0232 0.867 0.26 8.21 3.0258 0.074 0.19 4.0331 0.72 6.694 0.27 7.217 0.0268 1.0155 0.061 0.0141 0. " c.0123 0.165 0.143 0.0163 0.0154 -0.0088 0. -0.1826 198100.18 2.471 0.96 -2.30 9.23 4. " -0.95 -0.22 5.28 10.801 0.892 0.0124 0.991 0.0242 0.92 -1.0145 0.66 3.124 0.0151 0.267 1.17 0.499 0.0352 0. = 0.90 0.0186 0.0223 -0.666 0.002 Rn= 200900.30 cd 0.0196 0.0092 0.863 0.0100 0.0199 0.89 -0.91 -0.0212 0.23 5.003 0.0106 0.0092 0. "' -0.26 6.0134 0.0207 0.385 0.0095 0.269 -0.929 1.0206 .960 0.76 Rn 0. 12.99 -0.22 3.18 3.140 -0.96 -1.088 0.0143 0.395 0.142 -0.0344 1.0072 0.845 0.0211 0.0218 0.0215 0.0245 Rn c.089 0.821 0. 0.0125 0. = c.0086 0.98 -0.27 7.07 -5.921 0.007 1.0160 0.0145 -2.99 -0.91 0.89 0.841 0.0104 0.935 0.0431 202500.0182 0.0111 0.32 10.157 0.764 0. "' -0.90 -1.211 0.0156 -0.93 -0.0536 303800.255 0.0136 0.0215 Fig.0388 0.96 -2.0113 0.835 0.0283 0.19 0.0258 0.21 4.0247 1.876 0.0123 0.619 0.27 6.144 0.0170 0.0384 1.14 1.16 0. 721 O. cd 0.371 0.97 -3. 767 0.98 -3.209 0.0627 cd 0.403 0.26 5.599 0.640 0.127 0.0081 0. " -0.30 10.0132 0.29 8.32 0.23 6.0119 0. 792 0.0090 0.23 5.247 -0.20 4.536 0.884 0.751 O.176 -3.146 0.071 0.88 0.315 0. 726 0.94 -0.718 o.176 0.28 Rn = c.0129 0.0228 0.0330 1.31 9.0172 0.129 0.758 0. 795 0.08 Rn -4.26 7.15 1. 794 0.151 0.81 12.0605 Rn 304100.28 8. cd "' -0.161 -3.93 Rn cd -2.0170 5.31 10.849 0.83 0.0400 12.66 3.0231 0.0183 -3.30 8.372 0.32 Rn 0.0167 0.32 10.91 -1.0122 0. cd -0.0093 0.0207 -2.0273 150200.0142 0.931 1.88 0.187 0.0253 0.0228 0.27 0.0116 0.15 1.23 5.983 0.0075 0.31 10. 12.040 0.0097 0.31 = c.928 0.505 0.92 -2. 755 0.00 -4.918 0.009 0.91 -1.908 0.0204 0.371 -0.0150 0.0133 0.25 6.123 0.006 -3.95 0.474 0.0270 0.0144 -0.31 9.74 6.0098 0.0083 0.07 2.0097 0. = c.0115 0.572 0.635 0.0160 0.0179 0. c.0239 1.0076 0.22 4.32 10.12 1.30 9.0163 0.26 5.876 0.521 0.844 0.32 11.0780 0.885 0. -4.0147 0.0186 0.788 0.0244 0.91 -2.98 -1.0196 0.11 1.27 0.26 0. c. c.303 0.31 0.0237 -4. = c.0138 -0.273 0.27 7.0211 0.400 O.85 0.0143 0.15 1.519 0.25 6.0154 "' -0.96 Rn -2.93 0.29 7.388 0. 721 0. cd -5.0296 11.0116 o.13 1.0160 0. 12.22 4.0132 0.27 0.0176 0.579 0.035 0.249 0.30 9.974 0. ex c.563 0.32 11. -5.0152 0.30 9.0109 0.0142 -4.322 0.17 3.010 0.21 0.16 2. "' -0.936 0.17 1.494 0.376 0. 968 0.077 0.142 0.874 0.0206 8.0135 0.0187 0.060 0.15 2.0109 0.0166 0.004 0.0314 SD2083-PT Fig.993 1.107 Rn= 60000.19 3.0094 0.27 6.0073 0.0092 0.393 0.942 0.0180 4.106 Rn 63300.395 0.0155 0.94 -1.030 0.0140 0.341 0.0113 0.24 6.231 O.436 0.661 0.13 1.026 0.Ql08 0.87 -0.17 1.373 0.0182 0. " -0.0096 0.0097 0. "' -0.866 0. = c.28 8.945 1.0323 1.0288 -0.0108 0.30 c.0125 0.693 0.0108 0. 761 0.053 cd 0.33 = -0.96 -3.0101 0.90 -1.635 0. 10.28 9.29 8.0403 150400.0206 0.15 -0.0177 2.14 1.042 0.201 0.021 0.19 2.16 3.593 0.90 -0.20 3.479 0.30 9.109 Rn 59100.0189 SD6060-PT Fig.022 0.145 0.Ql03 0.0162 0.0121 0.11 1.0460 101200.0187 0.322 0.806 0.25 0.29 0.066 0. = "' -6.0100 0.0145 -0.0169 0.914 0.0121 0.21 5.259 0.749 0.081 -3.92 -0.23 5.0136 0.18 1.24 5.0240 0.0195 0.0109 0.0277 1.0180 3.96 -0.24 4.90 -0.0261 0.0492 0.743 0. 711 0.18 2.94 -1.0132 -1.983 1.20 3.D203 0.021 0.0345 0.0148 0.062 0.87 0. c.0178 0.88 -0.265 -5.0156 0.906 0.277 0.0229 0.78 9.165 0.0156 0.23 4.958 0.045 0.Ql71 0.0115 0.07 -2.331 0. 0.24 4.0460 0.30 9.524 0.0104 0.572 0.25 6.0662 100600.0080 0.05 -4.01 -2.125 0.18 2.056 0.0138 0.95 -2.30 8.0163 .24 4.26 7.0107 0.19 3.28 8.320 -0.0138 0.670 0.23 4.0197 0.92 -0.645 0.424 0.0169 1.17 2.658 0.867 0.161 0.825 0.23 5.92 -1.30 0.25 5.007 cd 0.302 0.490 0.382 Airfoils at Low Speeds 0.28 7.31 Rn = c.93 -0.053 0.90 -1.0158 -0.0237 0.628 0.0893 0.072 c.0150 0.0113 0.184 0.0152 0.27 7.0174 7. 491 0.0092 0.171· 0.033 0.25 7.20 4.90 -2.224 0.30 10.93 -1.33 8.0094 0.950 0. "' -0.512 0.975 205500.30 11.161 0.0220 0.838 0.468 0.0134 0.30 7.225 1.083 0.66 3.29 10.0957 154200.0131 0.0097 0.0126 0.287 0.006 0.19 4.0167 0.30 9.0327 0.829 0.30 11.0192 1.407 0.0398 0.0339 1.113 -0.112 0.0117 0.0466 0.30 11.16 2.501 0.349 0.679 0.0334 0.11 1. cd "' 0.0128 1.27 6.0119 0.0198 0.0151 0.0132 0.667 0.0094 0. "' -0.942 0.177 0.0189 0.934 0.30 11.0394 0.305 0.634 0.0107 0.89 0.731 0.937 0.113 Rn 59400.221 0.397 0.29 10.26 5.21 3.130 0. 742 0.28 8.0103 0.0119 0.610 0.0239 0.197 0.93 -0.0090 0.0102 0.20 4.0115 0.0154 0.0094 0.288 0.21 4.67 3. -3.20 4.17 2.0129 0.0159 -0.801 0.114 -2.0499 301100.0291 1.33 12.37 10.31 cd 0.992 0.0184 0.27 6.0150 0.0180 0.34 9.01P2 a.0246 -1.32 9.0139 0.001 0.090 1.0123 0.91 0.0101 0.0611 0.30 Rn 0.89 0.27 6.36 11. -0.25 6.0114 0.14 1.24 6.0267 1.0091 0.0077 0.27 7.129 0.510 0.871 0.22 5.0150 0.0114 0.15 2.011 0.0134 0.12 1.0118 0.94 Rn -1.21 2.24 7.0109 0.27 7.0274 0. 705 0.935 0.0141 1. " 0.30 0.977 0.915 0.28 9.0114 0.455 0.860 0.96 -2.27 5.110 -2.24 6.83 0.28 8.0295 0.91 -0.797 0.95 -2.353 0.0711 309100.24 6.0284 SD6060-PT Fig.0201 1.29 6.96 -1.0106 0.15 2. -2.42 0.213 0.26 7.0095 0.0117 0. 12.41 0.0121 -0.235 0.89 0.601 0.41 0.0096 0.0132 0. = "' -3.0106 0.24 4.536 0.0310 0.208 0.32 8. "' -0.200 0.695 0.23 4. 980 0.0106 0.144 0.94 -1.84 0.0336 0.0239 0.19 2.972 0.0106 0.020 0.0252 0.0087 0.166 1.21 4.0112 0.0078 0.15 2.022 0.94 -1.017 0.067 0.22 5.160 0. -0.0094 0.971 0.0198 0.17 3.404 0.123 -2.0334 0.29 9. = c.0331 0.200 0.0199 0.25 4. "' -0. = " cd 0. = c.15 2.0203 0.0420 307400.0251 0.0315 0.0179 0.0157 1.824 0.102 0.0212 0.078 0.0133 0.873 0. 796 0.27 8.30 9.28 9.0324 0.0101 0.499 0.0382 1.0235 1.0211 -0.167 0.598 0.94 -0.0082 0.0437 cd -2.0123 0.510 0.89 0.93 -1. 729 0.0258 1.28 8.89 0.28 9.0128 0.0098 0.0185 0.34 = c.605 0.884 0.12 1.21 3.28 8.101 0.983 SD6080-PT Fig.978 cd 0.0215 0. 731 0.0204 0.0233 0.0310 0.213 0.897 0.987 0.012 0.0370 0.0143 0.23 5.35 10.28 9.200 0. "' -0.0175 1.0080 0.30 cd -0.0167 0.0095 0.0669 313800.954 0.93 -1.19 4.0167 0.105 -0.108 -2.0328 0.D15 0.90 -0.26 7.0548 100400.0334 0.35 0.96 -2. c. 12. 701 0.93 Rn cd 0.29 9.609 0.0133 Rn 0.27 8.350 0.0092 0.087 0.22 -0.0564 0.137 0.0160 1.115 0.34 11.30 8.0209 0.0131 0.93 -1.13 1.0230 0.92 -0.18 3.608 0.0256 0.35 Rn 0.923 0.981 0.31 10.563 0. 794 0.0209 0.0112 0.0098 0.0266 0.0114 0. = c.Chapter 13: Tabulated Data -4.398 0.211 0.30 cd 0.0168 5.006 0.34 11.29 9. = c.0115 0. 965 0.24 4.507 0. cd 0.16 3.33 10.846 0.0283 0.707 0.0136 0.0143 0.916 0.13 1.88 -0.180 -3.18 2.521 0.003 1.879 0.805 0.23 3.0268 0.26 7.0214 0.381 0.0465 199700.0077 0.66 3.27 7.88 -0.985 0.67 3.0224 0.41 0.0131 0.942 0.0414 152000.965 0.28 8.0159 0. c.27 8.0170 0.663 0.099 -0. " -0.0150 1.0143 0.399 0.154 -3.0153 0.31 Rn = "' -0.0103 0.17 3. "' -0.062 0.31 7.23 3.17 1.&73 0.15 2.0281 0.0249 0.0095 0.30 Rn = c.90 0.25 7.981 cd 0.0107 0.13 1.804 0.0134 0.0358 0.504 0.30 10.D78 0.30 -0.930 0.191c.0107 0.443 0.28 = c.94 -2.0297 0.99 -2.0145 0.661 0.0081 0.0137 0.19 4.334 0.119 0.0098 O.877 0.24 6.0157 0.0123 0.0103 0.975 0.317 0.29 Rn = Rn cd 0.27 8.97 Rn -2.110 Rn 100900.74 6.29 10.0098 0.26 5. "' -0. "' -0.0450 0.30 9.0179 0.0119 0.1261 200900.016 0.88 0.90 0.0122 0.954 0.640 0.012 O.22 5.0121 0.91 0.300 0.731 0.12 1.200 0.627 0.0415 SD6060-PT Fig.97 -3.32 8.21 4.18 3.914 1.633 0. c.978 0.31 10.0100 0.30 7. 709 0.74 6.12 1. = c. 794 0.0133 0.0109 0.0141 0.431 0.804 0.20 4.14 1.0102 -0. 707 0.001 0.0199 0.212 -3.0150 0.0451 = 149700.0098 0.0166 0.223 0.0099 0.0125 0. c. cd 0. 12.829 0.0178 0.20 4.0196 383 .061 0.92 -0.0192 0.0455 0.31 10.0097 0.18 3.0094 -0.213 0.0241 0.0121 0. 784 0.232 1.0318 0.582 cd 0.0109 0.15 1.30 11. 771 0.142 0.417 0. 770 0.35 9.22 5.0121 0.0117 -0.18 3.026 0.91 -0. = c.13 1.0122 0.0112 -0.215 cd 0. 944 o.081 -4.26 8.209 -3.967 0.0087 0.35 Rn = "' -3.26 7.0113 0.0257 200500.94 -1.480 0.859 0.91 -0.352 0.937 0.94 -1.315 0.846 0.499 0.89 -1.177 0.879 0.23 6.24 6.28 8.023 0.0217 0.302 0.076 0.0098 0.240 -1.103 0.297 0.191 0.75 6.0110 0.88 -1.092 0.923 0.0138 0.35 9.29 -0. 983 0.98 Rn cd 0.111 Rn 103200.892 0.609 0. 977 0.398 0.036 -3.96 -2.0398 1.328 0.30 9.0109 0.26 7.30 10.0123 -0.0107 0.29 7.923 0.18 1. c.186 0.94 -1.16 2.22 5.299 0.0102 0.73 6.094 0.961 0.0653 -0. 973 0.89 -1.69 3.280 0.0086 0.12 1.34 10.166 cd 0.0158 0.0102 0.23 4.33 9.0124 0.0163 1.0247 1. 12.0222 1.0202 0. " -2.27 6.210 cd 0. 0.595 0.434 0.26 8.0082 0.0150 0.629 0.0091 0.85 Rn = " -2.024 1. 72 3.0308 1.34 9.210 0.429 0.0099 0.88 -1.204 0.35 cd c.0179 1.31 8.0210 0.056 0.003 1.28 7.34 9.437 0.34 c.627 0.21 2.0103 0.0072 0.90 -1.0169 0.698 0.34 Rn = "' -3.32 8.0099 0.890 0.0221 1.34 10.181 0.095 0. 0. "' -2.94 -2.421 0.0121 0.39 0. " c.581 0.0860 Rn = 299800.0217 1.30 7.91 -0.26 5.0138 1.590 0.34 cd cd 0.426 0.0825 Rn = 308500.0123 0.169 0.88 -1.92 -2.0225 1.24 4.028 0.161 0. 709 0. 724 0.0274 0.0136 0.35 0.88 -1.210 0.528 0.32 8.28 7.0494 1.30 7.0108 0. " -3. cd 0.33 9.36 1.028 cd 0.431 0.35 c.92 -2.190 0.87 -0.0179 0.230 0.0165 0. " -2.35 c.0144 1.0479 0.31 0.889 1.19 1. 0.0130 0.730 0.121 0.24 4.0179 1.35 10. 0.207 1.20 2.431 0.036 0.215 0.0501 Rn = 153800.0206 0. 0.891 1.0150 0.188 1.0143 0.085 1.30 7.133 0.176 1.64 3.219 0.0102 0.0092 0.0230 cd 0.406 0.0718 Rn = 204800.24 4.116 Rn = 99100. 70 3.0149 0.0098 0.34 9. " -2.0143 0.35 c.586 0.0164 0. c.0134 0.18 l.354 0. " -2.78 6.381 0.0133 0.34 10.125 0.0118 0.0098 0.0149 0.0149 0.213 0.32 c.026 1.34 10.527 0.0080 0.003 1.691 0.34 9.0107 0.29 6.0147 0.127 0.35 11.114 Rn = 100100.35 9.208 cd 0.0115 0.0167 0.0119 0.743 0.0161 0.186 cd 0.34 11.0097 0.88 -1.0108 0.0443 SD6080-PT Fig.042 0.34 10.323 0.22 4.31 7.1122 151900.90 -1.114 0.0080 0.19 6.118 Rn = 60300.37 0.0118 0.77 6.36 0.13 5.0186 0.672 0.990 1.87 -0.0165 0.36 0.0153 0.36 1.911 1.741 0.074 0.93 -1.522 0. 72 3.0406 1.0128 0.105 0.0225 9.21 3.0118 0.835 0. 0.35 10.212 0. 0.91 -1. " -2.19 4.0285 10.25 4.35 9.0461 Rn = 197800.17 1.521 0.35 10.110 0.188 0.0150 0.188 1. " -2.31 7.0192 1.367 0. = " c.86 12.0125 0. 75 6.0220 0.35 Rn = "' -3.739 0. " -2.27 5.025 1.0112 0.593 0.0201 1. 0. 12.36 10.78 6.0148 0. cd 0. cd .29 7.417 0.172 0.32 9.34 9.110 1.0094 0.37 0.0115 0.267 0.0243 0.0140 0.113 1.13 1.0264 1.0204 0.583 0.182 1.230 0.0180 0.0208 1.16 3.16 4.89 -1.17 1.180 0.981 0.18 1.32 8.983 0.0297 0.0135 0.30 7.209 0.032 0.116 1.438 0.32 7.180 0. cd 0.0111 0.189 0.0090 0.89 -1.285 0. cd c.0125 0.0137 0.87 0.0164 0.79 6.0101 0.112 1.25 4.0169 0.88 -1.14 1.35 10.893 0.0119 0.0157 1.92 -2.0106 0.0169 0.19 1.0150 0. 77 6.160 1.38 0.287 0.0202 0.23 4.320 0.98 -1.0145 0.0144 0.419 0.27 7. 78 6.014 0.0134 0.0405 Rn = 304200.212 1.0201 0.0474 151700.23 4.74 6.0177 1.17 1.17 1. 739 0.33 Rn -2.0163 0.161 1.0153 1.0124 0.086 1. cd 0.32 11.0321 1.384 Airfoils at Low Speeds -3.0186 1.0215 1.670 0.164 0.14 1.106 0.113 0.69 3.0215 1.742 0.0093 0.16 1.0089 0.0314 1.39 0. 0.279 0.19 2.0098 0.81 9.910 1.621 0.000 1.874 0.34 c.358 0.087 0.0246 0.23 3.0133 0.87 0.210 "' c.24 4.024 1.0145 0.0122 0.121 0.35 10.0175 SD6080-PT Fig.208 0.205 0.125 0.427 0.16 2.34 10.622 0.0094 0.0174 0.67 3. 72 3.218 1.35 c.0292 1.835 0.591 0.0125 0.24 5.181 0.680 0.826 0.76 6.88 -1. 72 3. 72 3.214 0.17 1.0102 0.88 -0.40 0.16 1.34 c.0141 0.190 1.0116 0.215 0.17 1.34 11.889 0.36 c. -2.977 0.267 0.0168 1.29 7.0116 0.013 0.115 Rn = 100500.0444 Rn = 149300.225 0.197 0.0162 0.0221 0.743 0.0154 1.31 7.818 0.0340 1.274 0.32 8.32 9.38 0.0506 1. 12.174 0.0109 0.0241 200000.0226 1.0190 1.211 0.0157 0.0241 1. 70 3.0286 0.510 0.897 0.011 0.91 -1.0231 1.807 0.0337 0. 70 3.117 Rn 98100.820 0.36 0.152 0.108 0.35 1.885 1.0105 0.0264 0.0417 1.109 0.77 6.79 9.211 0.270 0. cd 0.0244 1.28 6.212 0.199 0.80 9.0086 0.0434 11.0130 0.0795 199400.35 0.108 0.35 10.077 1.0158 0.329 0.85 0.885 1.075 0.182 1.0222 0.0097 0.90 -1.133 0.0148 0.280 0.0110 0.85 0.32 0.0123 0.136 0.0199 0.934 0.0261 1.32 8.23 3.719 0.0096 0. cd 0.0110 SD'T003-PT Fig.203 0.0482 1.16 1.0117 0.416 0.25 4.210 0.0131 0. c.0139 0. 0.0205 0. 71 6.0112 0. 725 0.90 -1.33 8.026 0.127 0.0134 0.558 0.0104 0.0103 0.29 7.24 4.214 0.243 0.510 0.0443 SD6080-PT Fig.0130 0.32 8.739 4.0440 1.0114 0.33 11.23 4.582 0.0091 0.113 0.811 0.0109 0.136 0.31 8.048 0.0171 0.0117 0. 12.0099 0.844 0.0775 SD6080-PT Fig.0143 0.30 7.82 9.302 0. "' -2.0165 0.32 8.0155 1.216 0.519 0.27 7.0118 0.0709 Rn = 300300. 71 3.095 1.0101 0.33 8.28 7.32 8.0107 0. 78 6.057 0.0178 0.0148 0.150 0. 12.0173 1.073 0.39 0. = "' -2.910 1.24 c.803 0.24 4.0119 0.77 6.140 0.528 0. 20 4.0112 0.26 7.66 3.0143 6.20 0.0109 c.28 0.80 9.0086 0.29 0.950 0.79 0.510 0.11 0.0104 0.0101 -0.507 O.816 0.0092 0.0115 0.41 0.0223 0.0114 0. 12.1169 0.0128 0.943 1.0167 0. Ct -2.19 4. 779 0.66 3.0091 -1.195 0.201 0.31 9.516 0.0245 10. 799 0.40 0.353 0.0079 0.0210 9.0225 1.0126 5.0119 4.138 0.0208 0. Ct -2.0160 7.925 0.122 Rn 101700.0110 0. = c.14 1.513 0.0084 0. Ct -2. Ct -0.0185 7.31 1.400 0.0215 9. Ct -2.0086 0.409 0.029 0.15 1.0153 0.0079 0.0949 200800.696 0.053 0.31 10.0093 0.784 0.67 3.14 1.099 0.350 0.Q111 4.20 0.92 -1.0231 8.014 0. = 0.0099 -1.501 0.26 7.0471 300500.0086 3.0194 8.66 0.0374 = " c.0138 0.983 1.978 0.79 c.992 0.28 8.927 0.1814 12.92 -0.354 0.024 0.19 4.099 1.17 0. Ct " -0.055 0.31 1.42 0.83 1.0254 0.12 1.12 0.32 10.26 0.32 1. Ct " -0.874 0.66 0.04 7 0.72 0.18 3.24 0.302 0.25 0.0138 0.888 0.184 0.507 0.79 9.28 0.17 3.0159 385 .995 0.0284 1. = "' c1 cd -3.95 -1. 0.74 6.82 1.936 0.40 0.98 -2.413 0.0450 11.0168 0.40 0.0080 0.108 0.0416 11.199 0.904 0.12 1.0092 -1.664 0.0142 0.83 11.0455 11.053 1.917 0.0183 0.1398 205000.659 0.518 0.33 Rn = " -3.0144 0.79 0.16 2.13 0.0150 0.000 0.0325 0.809 0.1858 300300.40 0.0100 0.20 4.296 0.0114 0.926 0.1601 150600.67 0.0236 1.41 0.0076 1.0293 10. 12.598 0.14 2.077 0.666 0.024 0.0109 0.195 0.936 0.79 Rn = c1 " -0.057 0. SD7003-PT Fig.0091 0.0391 Rn = " -2.0095 0.1200 151800.0253 0.78 9.992 0.0165 0.74 0. 0.0186 9.25 0.119 0.501 0.672 0.94 -1.0417 0.0078 0.0077 0.0101 -1.978 1.142 0.19 0.054 0.933 0.0267 0.000 0.31 10.19 0.0131 6.94 -0.0177 0.0151 7.120 Rn 199500.0376 11.181 0.601 0. 953 0. 5.995 0.0302 10.651 0.113 -2.92 -0.41 0.301 0.905 0. 0.94 0.024 0.0093 0.402 0.099 -2.954 0.899 0.038 0.0093 3.34 11. Ct -0.0083 3.808 0.0088 1.34 11.0480 11.13 1.0100 0.0082 0.0074 0.865 0.0100 0.14 1.80 0.32 10.30 Rn = " -2.30 0.18 3.78 9.67 3.96 0.16 0.78 0.67 3. 719 0.0112 0.191 0. c1 -0.23 0.0079 1.112 0.95 -1.805 0.0095 0.96 -0.0214 1.906 0.0124 2.66 3. SD7003-PT Fig.82 Rn c.0252 10.22 5.070 0.0108 1.12 0.286 0.73 0.119 -2.0505 SD7003-PT Fig.0133 6.100 0.31 10.0188 0.0073 0. c.0098 0. 0.0222 1.0158 7.658 0.14 0.0086 0.0290 9.982 0.25 6.0089 3.92 -0.32 10.25 0.0080 0.789 0.79 9. SD7003-PT = 9.110 0.234 0.41 0.28 8.206 0.0102 0.20 4.19 4.0121 0.0103 3.18 4.31 10.809 0.034 0.74 6.921 0.21 c.0104 -0.0110 0. Ct -0.73 6.0125 0.78 9. 12.92 -1.133 -2.039 0.40 0.929 0.0084 0.28 7.035 0.82 0.094 0.103 0.40 0.0227 1.645 0.0096 -0.21 0.126 -2.121 Rn 201200.0091 0.797 0.30 0.078 0.0129 4. -2.31 10.699 0.0076 0.0207 0.95 -1.1563 Rn 153300.632 0.94 -1.0187 = = "' SD7003-PT Fig.32 10.0166 6.131 0.23 5.172 0.66 3.73 0.25 0.0258 0.0121 0.0278 0. c1 " -0.517 0.507 0.67 3.0308 0.108 c. c.26 7.926 0.203 0.511 0.82 Rn = " -2.73 6.83 1.73 6.496 0.620 0.0081 0.26 7.0157 0.89 0.0122 -0.098 c.27 7.78 9.Chapter 13: Tabulated Data -2.82 Rn = " -2.0139 7.212 0.036 0.662 0. 793 0.20 4.119 Rn 99600.511 0.27 7.209 0.20 4.0108 -1.30 9.0079 1.0120 0.808 0.93 -0.31 1.80 0.0145 5.0325 0.19 0.0143 0.196 0.93 -1.0123 0.115 0.28 0.015 0.80 0.026 0.0093 -0.24 6.31 10.15 0.91 0.005 0.049 1.79 9. 791 0.112 0.2202 c.40 0.0327 10.871 0.66 0.0127 0.359 0.942 0.366 0.878 0.939 0.0337 0.0130 0.207 0.0110 0.076 0.90 0.214 0.353 0.29 0.013 0.030 0.176 0.109 0.048 0.511 0.27 7.13 1.921 0.0111 -0.0118 0.0193 0.32 c.106 0.0072 1.19 4.32 1.0251 9.93 -1. Ct " -0.0095 0.0353 303700.103 0.0085 0.94 -0.31 9.31 1.0146 0.502 0.19 0.14 1.0121 0.072 0.0158 0.21 0.33 0.036 0.0081 0.26 7.901 0.0298 1.80 0. Ct " -0.0248 0.12 1.362 0.504 0.037 0.0072 0.12 1.189 0.91 -0.21 4.74 6. 12.0096 0.878 0. 783 0.0111 4.0081 0.801 0. " = c1 " -0.19 0.90 0.32 1.664 0. c. " = c.0123 0.0356 0.611 0. 12.0167 0.939 0.0146 0.82 1.806 0.0101 0.058 0.664 0.0099 4.0097 0.13 0.008 0.042 0.814 0.91 -0.27 7.192 0.0082 0.903 0.0109 0.27 7.0399 0. Ct -0.596 0.354 0.0131 0.362 0.79 0.111 -2.0116 -0.0127 3.2182 Rn 298000.0117 6.92 -0.81 Rn = -0.0091 0.1646 Rn 201500.0095 0.657 0.1799 12.022 0.91 0.93 -1.0490 = Fig.711 0.0079 0.0346 0.0736 102000.79 0.27 0.0205 0.1899 Rn 301600. c.052 1.0096 2.0158 0.29 8.494 0.0094 0.358 0.0100 4.0115 0.123 Rn 99700.362 0.318 0.74 6.692 0.101 1.046 0.74 6.017 4 0.23 6. c.392 0.864 0.15 2.0129 -1.007 4 0.0141 7.040 0.0088 0.94 -1.370 0.0376 10.0114 6. 860 0.11 c.0125 0.0089 0.109 0.888 0.0491 0.213 0.37 1.267 0.675 0.31 1.0311 0.53 -0.35 1.0161 0.878 0. 12.383 0.18 3.18 1.20 2. c.105 0.692 0.0175 0.43 5.0166 0.72 Rn = "' -3.65 0.0176 0.35 9.09 9.747 0.0157 6.161 0.863 0.647 0.0102 2.0121 -0.0134 0.559 0.25 4.0127 0. 12.09 8.291 0.72 11.161 0.633 0.0123 0.93 -2.0254 1.0316 11.126 Rn 64200.210 0.16 1.87 8.0078 0.74 13.297 0.49 -0.0131 0.928 0.0341 1.0148 0.51 10.091 0.655 0.89 -1.410 0.404 0.098 0.28 0.845 0.124 Rn = 100600.0098 0.0139 0.87 Rn = "' -3.0098 0.0080 0.197 0.08 8.538 0.035 0.0100 -0.95 -2.81 1.839 0.659 0.0261 0.214 0.010 0.0219 0.0223 0.28 Rn = "' -3.0128 3.85 7.0132 4.0101 0.29 8. c.076 0.0112 0.816 0.0353 1.0398 1.386 0.31 9.525 0.51 10.756 0.0143 0.404 0.30 0.049 0.831 0.86 7.267 0.0120 4.31 9.63 5.0090 0.282 0. -0.18 3.222 0.163 0.86 7.0242 1.661 0.0130 0.066 0.642 0.079 0.002 0.29 8. "' -3.84 0.512 0.27 5.71 0.0121 0. -3.96 2.96 -2. -3.0178 0.109 0.0100 0.97 3.956 0.36 1.360 0.0195 0.33 10.24 0. "' -3.38 Rn = "' -3.101 0.038 0.028 0.0161 0.0273 0.0221 0.29 6.301 0.0094 0.891 0.233 0. O.40 4.0164 0.71 Rn = ex -3.0199 0.0253 1.96 -2.0253 -0.834 0.0093 0.530 0.0132 -1.199 -0.21 4.0205 0. 779 0. -0.415 0.0112 0.0094 0.0105 0.0097 0.52 -0.29 9.19 0. cd 0.942 0.24 3.90 -1.0252 0.21 2.0437 301600.0200 SD7032B-PT Fig.98 13.73 -1.065 0.0635 13.117 0.0231 0.0208 0.057 1.0108 0.0197 0.863 0.0078 0. 729 0.89 0. c.34 11.0108 2.541 0.0168 -2.29 9.1817 0.50 10. -0.0148 0.0376 0.36 10.92 c.41 4.286 0.83 0.151 0. 768 0.0266 1.0179 0.0202 0.053 0.0098 1.293 0. 764 0.087 0.431 0.2246 SD7003-PT SD7032A-PT Fig.0097 0.0328 1.30 0.702 0.94 2.553 0.37 11.0140 0.0200 0.882 0.74 1.0248 0.206 0.620 0.243 0.654 0.090 0.046 0. c.0333 1.050 0.96 -0.259 0.87 -0.632 0.73 Rn = ex -0.903 0.610 0.0257 10.0360 0.25 4.775 0.0192 0.404 0.524 0.97 -2.0376 1.0144 0.167 0. Ca -0.27 0.0110 0.109 0.89 -0.74 -1.85 7.0123 -1.457 0.29 7.041 0.594 0.0137 0. "' C1 Ca -3.341 0.27 6.24 0.0295 1.2163 104600.315 0.868 c.0091 0.34 9.0129 -0.34 11.0096 1.131 0.73 11.196 0.91 -2.965 1.525 0.32 8.96 -1.284 0.106 0.023 0.472 0.207 0.022 0.26 7.D108 0.0219 9.94 -2.076 0.74 -1.0272 1.64 6.0145 0.32 10.409 0.86 -0.162 0.0188 0.122 0.782 0.0215 1.28 6.0366 0.36 1.744 0.0271 1.0143 0.0119 0.0214 8.53 10.0134 0.078 0.18 3.0178 6.0496 11.84 0.73 11.954 0.872 cd 0. c.271 0.0178 0.0081 0.25 0.909 0.88 0.0077 0.003 0.94 2.0138 0.0108 0.0145 5.25 5.0150 0.983 1.859 0.40 4.63 5.0108 0.0311 1.0084 0.17 3.29 9.483 0.52 11.30 9.242 0.85 0.28 0.0588 14.499 0.869 0.0116 -0.0079 0.26 5.929 0.414 0.0477 Rn = 101300.501 0.93 -2.0154 0.0265 0.92 -2.0328 1.286 0.17 3.85 0.873 0.91 -0.21 0.076 0.154 0.0386 12. cd -0.35 -0.0121 -0.0343 1.0186 0.0111 0.22 4.39 4.642 0.0123 0.0108 3.043 0.40 4.217 0.0113 0.969 0.0274 1.51 -0.997 0.256 0.26 6.28 7. c.30 0.17 1.0966 Rn = 301700. -0.0144 0.272 0. "' C1 Ca -3.0115 2.517 0.039 0.955 0.29 0.24 0.73 -1.09 8.881 0.067 0.27 0.0214 0.37 1.87 0.939 0. 769 0.851 0.63 5.18 0.24 5.28 0. "' -3.0130 0.29 0. 763 0.340 0.23 3.32 1.087 0.49 0.30 7.0139 .52 -0.0162 0.20 2.34 1.33 13. c.07 8.1591 Rn = 300700.319 0.0084 0.140 0.0217 0.0196 0.0310 1.0139 0. c.0100 0.013 0.804 0.0274 0.0203 0.446 0.0164 0.0224 0.977 1.093 0.23 3. 12.90 -1.202 -0.0145 0.51 10.87 0. -0.0131 -0.045 0.40 4.52 10.045 1.23 4.813 0.29 0.039 0.095 0. Fig.64 5.0215 0.0175 7.25 4.18 3.078 0.201 0. "' c.0163 0.0089 0.92 -2.30 9.0329 1.91 0.86 7.21 3.88 0.35 10.43 -0.0219 1.0228 0.30 9.0153 -2.72 -0.0256 0.64 5.0156 0.86 7.0494 0.0242 1.0416 Rn = 203400.995 0.0127 0.94 2.1357 151800.998 0.679 0.73 -1.92 -1.0090 0.19 1.26 5.0236 1.357 0.045 0.19 2.63 5.67 -0.0106 -0. cd -2.0167 0.415 0.07 c.0161 0.065 1.151 0.22 2.779 0.86 -0.34 1.25 0.0095 0.012 0.07 8.0103 0.0209 1.0081 0.0141 0.31 7.0324 9.927 0.0121 0. 712 0. c.0090 0.33 10.0134 0.0102 0.0120 3.0221 1.20 0.1724 0.0202 0.386 Airfoils at Low Speeds -2.152 0.37 11.981 0. cd "' c.51 -0.1808 Rn = 150200.0090 0.95 2.0196 8.127 Rn = 61200.27 6.744 0.0461 202300.21 3.0208 0.0093 0.2264 = <> -3.0424 1.0114 0.123 0.0220 0.157 0.180 0.31 0.84 0.08 8.952 0.83 0.961 0.0114 0.96 2.106 0.915 0.603 0.0424 0.069 0.533 0.149 0.1911 201400.34 12.582 0.19 0.286 0. c.0105 -0.0122 0.32 8.254 0.37 -0.19 1.192 0.0093 0. 78 0.89 -1.689 0.73 -1.232 0. c.567 0.0153 -2.115 0.0110 0.31 8.0152 0.0079 0.037 0.0152 5.0407 200100. c.973 0.116 0.951 0.0179 0.064 0. 37 c.0101 0.0180 0.17 2.0217 1. 01 -4.31 5.0292 1.18 1.D368 9.0140 0.0083 0.26 4.20 2.0334 1.0254 0.670 0.87 -0.86 -4.39 9.830 0.0115 0.0335 Rn = 102100.418 0.592 0.0117 0.26 2. = "' -3.689 0.0215 0.977 0.21 4.0169 0.0216 0.0301 0.399 0.22 1.0090 0.24 4.0118 0.0221 0.0314 0.786 0.218 0.0183 0.0275 1.28 5.36 11.0112 0.0258 1. 0.0150 0.0116 0.385 0.108 0.0121 0.294 0.0382 Rn = 151200.686 0.300 0.269 0.952 0.0218 0.0285 0.26 4.0225 0.0186 0.411 0.90 -1.722 0.34 8.0155 1.304 1.31 8.88 -2.0316 1.0180 SD7032C-PT Fig.646 0.067 0.22 1. -2.0121 0.221 0.88 -1.24 3.35 10.35 10.0521 Rn = 302500.84 0.18 1.177 0.18 1.405 0. c.127 0.0333 1.879 0.313 0.0125 0.149 0.955 1.27 2.29 5.0088 0. "' -2.961 1.072 c" 0.369 0.330 0.0503 Rn = 200500.383 c" 0.31 6.0152 0.36 Rn = 01 -4.0143 0.31 5.041 1.046 0.22 3.831 0.0248 0.31 8.32 7.271 0. 0.84 -0.263 -3.0157 0.317 0.906 0.82 0.24 3.78 -0.172 0.368 0.185 0. 12.011 0.274 0.844 O.23 2.32 c.023 0.0166 0.303 0.367 0.542 0.335 0.35 9.575 0.257 0.0424 1.20 2.86 -0.34 7.25 4.006 1.163 0.0190 0. -3.85 -0.39 Rn = c" c.28 3.0110 0.553 0.0243 0.0226 0.30 6.81 0.38 11.0201 1.497 0.353 0. "' -4.78 0.0131 387 .0114 0.29 4.303 0.23 2.413 0.0334 1.33 6.663 0.0252 0.048 0.31 7.989 0.36 10.0093 0.0248 0.391 0.0213 1.053 0. 753 0.871 0.177 0.0156 0.499 0.963 0.0132 1.35 9.127 0.31 4.506 0.395 0.0163 1.33 8.120 1. 0.23 2.0094 0.0107 0.596 0.0206 0.282 0.962 0.26 4.419 0.397 0.247 1.0280 1.32 7.30 5.213 0.125 1.87 -0.0239 0.0214 0.0099 0.0097 0.0339 1.0090 0.091 0.479 0.0219 0.26 3.0263 1.37 11.0241 1.0103 0.0189 0.0181 1.22 2.939 0.Dl08 0.92 -3.23 3.412 0.38 9.30 5.0150 0.0231 0.88 -2.309 1.0194 0.486 0.0191 1.27 4.124 0.281 0.34 8.982 1.0130 0.0129 0.35 c.224 0.0104 0.33 8.0189 1.0141 0.83 0.30 7.779 0.0131 0.0097 0.35 6.30 6.0345 1.0213 1.26 4.28 c.257 1.240 0.337 0.84 -0.30 7.120 0.0412 11.987 0.88 -0.38 8.387 0.85 -1.688 0.684 0.0189 0.0117 0.040 0.0188 1.117 0.312 0.871 0.0109 0.048 0.0106 0.0172 1.Dl76 0.0146 1.258 0. 0.34 9.0255 0.749 0.927 0.27 3.0277 1.0148 0.0250 1.0176 0.22 2.38 0.92 -1.34 10.21 2.0118 0.34 9.24 5.686 0.042 1.Chapter 13: Tabulated Data 5.0161 0.0082 0.493 0.238 0.588 0.126 0.22 3. 12.0331 Rn = 302000.38 12.027 0.12 1.88 -1.0142 0.38 11.35 9.0403 1.0172 0.0121 0.197 0.0086 0.0092 0.0112 0. -2.0226 1.27 6.416 0.825 0.160 1.86 -1.005 0.645 0.0104 0.0139 0. 12.91 -2.79 0.896 0.0212 0.27 5. 01 -3.129 Rn 60100.193 0.201 0.485 0.869 0.0276 0.33 8.199 1.291 0.260 0.857 0.0434 1.20 1.130 Rn 62200.29 6.0081 0.0164 0.244 0.0236 1.25 4.482 0.0288 0.0309 0.26 5.86 -1.83 0.39 10.0431 1.86 0.0125 0.27 9.0178 1.31 7.20 3.0166 1.0132 0.0152 c.37 0.38 11.37 11.31 7.85 -0.33 8.0269 0.34 9.91 -3.368 1.88 -2.30 6.0213 1.32 c.87 -1.194 0.0198 0.179 0.970 1.0413 Rn = 202700.23 3.19 1.052 1.594 0.402 0.911 0.87 -0.0194 0.201 0.909 0.92 -2.344 c" 0.0362 1.34 9.20 1.131 0. 770 0.695 0.0150 0.20 1.22 3.0292 1.38 11.30 8.0136 0.0148 0. 0.926 1.35 10.89 -1.0102 0.86 -0.82 -0.0244 1.328 0.0133 0.0187 1. 0.304 0.0260 0.30 6.25 5.35 9.85 0.0270 1.33 1.0333 10.832 0.305 0.0114 0.111 0.182 0.0290 0.26 3.123 1.250 0.317 0.39 1.30 6.100 0.0177 1.290 0. 0.0088 0.0201 0.0156 1.0112 0.85 0. c" 0. "' c.0528 Rn = 200300.0366 1.89 -1.947 c" O. 735 0.105 0.0088 0.38 c.621 0.27 5.0246 1.307 0.777 0.0196 1.37 0.89 -1.0101 0.0081 0.0159 0.30 7.36 10.362 0.23 1.32 8.0091 0.387 0.579 0. "' -5.850 0.0340 1.080 1.257 c" 0.89 -2. 715 0.196 0.0182 0.051 0.29 3.33 0.0162 0.26 5.0319 1.31 4.139 0.19 1.0235 1. c" "' c.24 3.473 0.35 9.261 0.38 0.0246 0.300 0.84 0.85 0. 775 0.570 0.499 0.37 11.229 0.058 0.0137 0. 780 0.36 10.0306 SD7032C-PT Fig.38 12.0171 0.959 0.26 4.0167 1.82 -1.128 Rn =59400.37 0.0450 1.863 0.23 1.0206 0.0092 0.0123 0.230 1.0234 1.85 -0.0238 0.213 0.86 -1.0112 0.30 6.508 0.0325 1.20 2.0172 0.546 c" 0.0149 1.39 1.25 2. 779 0.30 5.84 0.0247 0.84 -3.0312 Rn = 305400.0271 1.0302 O.0380 1.0159 0.84 -0.668 0. -0.0152 0.592 0.26 4.0128 0.318 0.0345 303200. c" -2.0313 0.1061 Rn = 101400.0192 1.383 0.28 7.30 6.0232 0.608 0.31 7.392 0.0195 1.0260 SD7032C-PT Fig. c" "' c.107 0.37 0.0261 0.82 -2.0148 1. 764 0.0299 0.31 7.0104 0. "' -2.87 -1.764 0.36 10. = "' -2.344 c" 0.343 0.0107 0.0186 1.0251 0.835 0.35 8.324 0.192 0.075 0.83 0.868 0.0211 0. c" 0.0167 1.31 6.188 1.22 2.196 0.27 6.90 c" 0.0141 0.34 10. "' c.265 0.32 7.146 0.17 c.688 0.0283 0.36 10.0148 0.197 1. "' -2.37 11.305 0.312 1.28 8.28 5.0124 0.0467 100500. 0142 0.0209 Rn 301600.0142 0.131 0.90 0.565 0.20 3.0772 0.0333 0.0216 -0.730 0.364 0.0130 0.91 -0.17 3.0118 0.20 2.29 8.239 1.28 8.29 6.91 -0. 12.638 0.33 1.34 11.33 9.013 0.077 0.95 -3.0259 0.0146 0.0403 101500.0128 0.915 0.0114 0.0266 0.0114 0.20 4.25 6.33 8.083 0.26 0.32 Ca 0.073 0.30 9.0138 0.427 0.0292 0.488 0.0145 1.793 0.975 1. 714 0.114 -2.291 0.35 1.180 -0.38 12.97 -2.156 0. 784 0.22 5.35 1.993 1.0189 0.19 4.063 1.0325 1.0277 SD7032C-PT Fig.0263 1.521 0.581 0.28 0.206 0.85 0.0173 0.0261 9.92 -0.24 7.0103 0. 794 0.0174 3.207 0.0206 0.326 0.26 7.048 0.30 7.425 0.0287 0.0225 1.146 1.22 5.0228 0.089 0.147 0.052 0.131 1. 1.243 0.31 10.18 4.91 -0.0336 1.330 1.729 0.183 0. c.125 0.362 0.0480 302600.13 2.12 1.115 -0.323 0.0201 0.0277 0.034 0.27 8.12 1.35 1.285 0.33 10.29 9.90 0. -0.0084 0.839 0.180 0.0224 8.917 1.88 0.19 3.17 3.0250 1.35 1.215 0.436 0.35 11.0239 0.27 8. 755 0.0150 0.31 10.0091 0.0208 0.377 0.924 0.88 0. 777 0.195 0.94 -1.91 -1.0524 SD7032C-PT Fig.823 0.536 0.37 12.272 0.37 Rn -0.0132 0.090 0.691 0.28 8.98 -3.34 c1 -0.0224 0.478 0.877 0.0182 10.23 3.900 0.0236 1.223 0.0112 0.894 0.073 0.23 3.027 0.632 0.108 1.0140 0.96 -2.0134 0.168 cd 0.25 0.0114 0.307 0.0090 0.0271 0.0151 0.35 -0.153 0.89 -1.91 0.0155 0.147 0.0255 -0.0188 4.0174 2.34 8.0290 300700.22 5.891 0.16 2.0126 0.980 1.92 -2.624 0.0394 1.579 0.145 -6.97 -4.060 cd 0.0166 0.297 1.0200 0.290 0.90 0.141 0.122 -2.243 0.281 0.0449 0.0171 0.34 1.20 c1 -0.29 9.0140 0.0095 0.0200 0.108 0.93 Rn -1.86 0.0291 1.0263 0.278 O.025 0.127 -0.15 1.33 8.042 0.18 3. = c1 "' -0.21 c1 -0.0211 0.0115 0.581 0.31 7.0165 0.28 8.0387 11.162 1.98 -3.0207 0.14 2.18 4.28 6.325 1.96 -3.0150 0.0576 = 199100.004 0.230 0. = c.225 0.002 0.23 6.0536 Rn = 201500.33 9.0193 0.0177 0.052 0.35 10.0127 0.0104 0.0320 1.35 13.333 0.93 Rn -1.20 3.307 0.0086 0.0845 Ca 0.30 10.35 0.813 cd 0.0273 Rn -4.619 0.31 1.083 1.17 1.444 0.237 cd 0.683 0.0105 0.0283 0.0239 0.0290 0.30 10.0182 0.0323 0.0346 0.0144 0.0131 0.153 0.22 5.34 Rn = "' -2.033 0.068 0.0317.33 11.91 -2.0104 0.0127 0.34 13.0185 0.0214 0.268 cd 0.613 0.046 0.34 10.35 11.433 0.00 0.620 0. = "' -3.061 0.30 7. 707 0.27 5.0136 0.168 0. 720 0.0158 8.0188 0.27 5.94 -1.17 2.0096 0.0220 7.0087 0.680 0.225 0.0385 0.29 9.24 6.0182 0.19 1.0159 1.0180 1.692 0.191 -2.25 7.671 0.0101 0.0306 10.21 5.20 5.0133 0.0080 0.131 Rn 62600.292 1.0086 0.16 1.128 -0.073 1.25 7.0269 0.231 -0.0356 0.275 1.0178 0.90 -0.31 10.34 9. " -0.260 0.34 11.0161 0.0117 0.221 1.93 -1.0095 0.882 0.0145 0.30 10.0182 0.0280 0.92 -2.35 12.34 11.13 1.22 5.01 -6.28 6.067 0.067 1.713 0.34 1.0158 0.0330 0.30 7.0319 0. = "' -5.21 0.442 0.0322 0.94 -1.229 1.0149 0.0157 0.89 0.525 0.0225 0.0096 0.0160 0.370 0.33 11.0165 1.96 -3.26 7.0096 0.26 5.0105 0.34 10.196 0.0202 0.36 11.18 3. 794 cd 0.0113 0.34 Rn = "' -4.36 11.94 -1.21 4.0102 0.0096 0.167 0.175 0.691 0.0107 0. cd c1 "' -0.192 1.275 0.02 -4.524 0.022 0.082 0.0134 0.269 0.0173 0.D106 0.31 c1 cd -0.271 0.90 0.20 4.0178 5.984 1.0163 0. 12.91 0.216 0. c.568 0.37 13.226 -4.89 -0.0122 0.93 -3.072 0.995 0.D103 .0175 0.0369 0.153 0.0152 0.0237 1.0165 0.30 1.34 0. = "' -5.36 Rn = "' -7.140 99800.0138 0.123 0.140 0.883 0.0173 0.33 12.0160 0.352 -0.0119 0.29 9.87 -1.17 2.0201 1.22 6. = c1 "' -0.999 0.544 0.0531 0.0212 0.0121 0.24 6.984 0.025 1.32 9.523 0.14 1.0250 0.20 4.0237 0.869 0.906 0.298 1.0469 Rn 147800.347 0.0093 O.141 1.054 7 0.17 3.961 0.045 0.0204 0.24 6.0254 0.G106 0.0141 0.0291 0.0280 0.31 0.18 1.0225 0.0205 6. 728 0.474 0.058 0.342 0.23 0. 12.0237 0.33 12.17 3.24 4.85 0.23 3.90 -1.808 0.110 0.91 -0.135 -0.473 0.89 -1.0246 0.0102 0.0123 0.0388 102000.310 0.0205 0. -0.0279 0.0115 0.0193 1.904 0.067 0.25 7.0169 0.0174 0.049 0.16 2.0443 200800.0295 0.87 -0.062 0.0335 1.13 1.20 2.817 0.523 0.21 4.0216 0.13 1.266 0.28 9.000 1.486 cd 0.179 1.17 2.24 4.13 2.0164 -0.0141 0.88 0.132 Rn 59900.25 7.34 12.94 -4.056 -2.89 -2.20 2.0120 0.383 0.91 -0.15 1.808 0.0124 0.0136 0.17 2.774 0. = c1 " -0.0151 0.851 0.0102 0.451 0.413 0.0167 0.27 8.0173 1.078 0.0094 O.30 9.441 -0.0203 0.22 6.605 0.586 0.955 0.334 0.0322 0.0110 0.0155 0.658 0.806 0.0146 0. = c1 "' -0.198 1.95 -1.0262 0.89 -0.0227 0.0196 0.137 0.91 -0.28 0.426 0.140 1.87 -0.90 -2.388 Airfoils at Low Speeds 6.964 0.369 0. "' -2.134 Rn 59200.0147 7.049 0. c1 -0.0086 0.0199 0.0156 0.0098 0.22 5.32 10.32 8.0140 0.021 0.92 -0.0124 0.0167 9.23 6.0176 SD7032D-PT Fig.0111 0.26 cd cd 0.26 4.0140 0. 155 0.77 5.87 -0.339 0. = c. " 0.609 0.011 0.907 0. 0.0237 0.0101 0.844 0.676 0.0124 0. 703 0.0115 0.95 -1.23 3.144 0.89 0.36 12.23 5.528 0.313 0.106 1.243 -3.139 -3.16 3. cd -0.968 0.0246 0.313 0.72 5.228 -0.70 2.339 0.214 1.95 Rn -2.29 6.126 -1.43 -0.0161 0.28 8.0402 204200. = c.37 -0.111 0.0282 11.89 Rn -2.735 0.389 0.0140 0.0154 0.0210 0.16 2.0228 10.0092 0.23 6.163 0.83 0. " 0.26 6.219 0.28 6.68 2.38 -0.0196 0.0471 1.054 98700.001 0.91 -0.0234 0.30 1. 708 0.054 0. = c.072 -3.0214 0.068 1.23 3.0159 0.0161 1.98 -2.0142 SD7037-PT Fig. c.83 0.0067 0.83 = c. " -2.0092 0.22 4.16 2.0218 0.64 2.059 -3.236 0.358 0.29 6. 0.83 1.0189 1.195 0.18 3.380 0.18 3.951 0.33 1.32 12.0128 0.0085 0.0169 0.0243 0.35 1.0230 0.31 7.441 0.0206 0.86 11.0491 SD7037-PT Fig.90 cd 0.0100 0.0180 1.31 c.0135 0.0103 0.92 0.275 0.117 0.31 -1.0127 6. " 0.089 0.0280 0.83 0.0422 302900.284 0.33 1.0150 0.303 0.0097 0.17 3.22 4.110 0.0126 0.514 0.97 -1.406 0.42 -0.233 0.944 0.0194 1.19 3.89 -0.699 0.021 0.0193 0.21 3.0216 -0.837 0.069 0.0187 0.0128 0.95 -2.0208 0.18 4.28 1.0169 -0.297 0.0182 1.35 11. 792 0.0142 0.479 0.0214 1.278 0.0145 0.2166 SD7043-PT Fig.537 0.683 0. = c.0075 0.471 0.0253 1.0387 1.18 3.936 0.053 0.247 0.0105 0.076 1.674 0.0196 0.486 0.37 -0.310 0. 12.019 0. " -0.0385 1.30 9.0090 0.93 -0.0117 0.666 0.0129 0.22 5.33 8.0218 0. 779 0.24 6.0689 201200.0144 0.0432 Rn 149600.010 0.059 0.25 6.22 cd 0.0609 151500.0104 0.365 0.0399 12.30 6.0259 0.0312 1.0396 0.0643 13.34 12.0167 0.28 8.874 0.988 0.17 2.19 3.0176 0.0321 0.24 5.982 1.31 9.0091 0.26 6.84 0.35 12.0338 11.0168 0.997 0. cd -0.98 -2.26 6.0292 cd 0.271 0.0178 0.367 0.26 6.977 0.125 0.0127 0.0135 0.0331 1.0135 0.283 0.24 3.0185 cd -0.0246 0.29 8.0145 0.78 8. " -0.78 8.77 8.18 3.119 1.0147 0.81 cd 0.352 0.472 0.29 5.095 0.811 0.23 4. = c.96 Rn cd 0.041 0.667 0.324 0.28 9.0174 8.81 8.72 5.0185 0.26 7.0114 5. " -0.0144 0.416 0.19 3.0170 0.35 9.30 -1.85 0.0232 0.0116 0.223 0.0243 0.0261 -0.098 cd 0.079 0.0110 0.31 10.053 0.0116 0.0258 0. " -0.91 Rn -2.0567 153400.0100 0.90 0. 70 5.0273 1.0289 0.20 5.32 9.100 0.32 -2.28 8.0182 0.533 0.0264 1.0163 0.0117 0.82 11.159 0.049 -3.0382 1.299 0.579 0.31 10.0163 0.21 4.90 -1.0102 0.712 0.27 8.571 0.0144 0.73 5. 0.0321 1.28 9. 77 5.0239 1.058 -2.85 -2.0101 0.96 -2. 12. = c.834 0.613 0.91 0.0145 7.837 0.22 4.0736 0.0213 0. 798 0.983 0. " -0.93 0.046 0. = c.974 0.0153 0.24 6.26 7.0186 9.321 0.0120 0.26 7.0198 1.153 0.058 0.0407 0.0137 0.127 0.32 1.891 0.0185 389 . = Rn ·a -3.513 0.0121 0.0090 0.286 0.904 0.Chapter 13: Tabulated Data 4.279 0.88 0.98 Rn Rn Rn cd 0.696 0.462 0.33 9.917 0.138 0.0455 1.0226 1.85 -2.0190 0.85 1.833 0.0240 0.219 0.068 0.156 0.0373 100700.0117 0.138 Rn 61600.72 5.0208 -0.0301 0.27 6.0095 0.31 = c.485 0.35 1.18 1.0099 0.214 0.876 0.30 10.0148 0.0216 1.0090 0.0489 SD7032D-PT Fig.87 -0.128 -2.32 = 0.132 0.85 0.83 11.65 2.94 -2.29 0.054 0.35 10.84 14.0121 0.279 0.885 0.784 = " -2.91 0.0108 0.01 2.521 0.67 2.92 -0.0088 0.27 7.82 0.17 1.0190 0.120 0.988 0.924 0.0147 0.0192 0.15 2.35 9.119 0.259 0.0118 0.90 0.0227 0.224 0.217 0.0163 0.0099 0.037 0.0129 0.22 cd 0. " 0.858 0.09 1.008 0.81 -0.86 11.14 1.0222 1.0186 0.30 9. 757 0.24 0.611 0.945 0.259 0.0343 1.0339 1.0178 0.0105 0.208 0.32 cd 0.37 12.0149 0.25 6.34 11.93 Rn -2.866 0.95 Rn 6.0127 0.25 7.309 0.0206 0.23 5.0317 1.287 cd 0.18 3.0140 0.35 1.039 0.64 2.0130 0.0167 0.0270 199800.89 -1.100 0.0098 0.19 3.0084 0.0134 0.194 0.828 0.0137 0.0193 0.65 2. 0.0275 1.138 0.92 -1.758 0.0210 1.135 Rn 60500.0685 0.694 0.0171 0.0137 0.0157 0.78 8.685 0. = c.617 cd 0.968 0.0198 0.109 0.0255 103900.13 1. " -0.173 0.27 8.206 0.30 10.015 1.0157 0.13 1. = " -3.757 0.606 0.945 cd 0.14 1.42 -0.96 c.318 0.84 11.136 Rn 62600.588 0.22 5.0086 0.35 9.046 -3.80 8.981 1.846 0.986 1.0350 c. " -0.829 0.0158 8.13 1.74 5.0179 0.29 0.008 -3.24 0.004 -3.074 0.93 -0.887 0.198 0.0100 0.0132 0.204 0.0320 1.22 2.308 1. = c.061 -3.0265 0.0147 0.86 0.415 0.16 2.45 -0.0221 0.0187 1.063 0.276 0.0087 0. = c.0131 0.27 7.30 9.79 8.597 0.92 -0.28 9.0116 0.261 0.36 1.0109 0.27 4.0219 9.89 0. 12.519 0.0101 0.225 -3.0246 0.223 0.0448 302000.0179 0.105 0.368 0. 71 2.618 0.0127 0.38 -0.947 1.137 Rn 57500.093 0.287 1.30 9.91 Rn cd 0.972 1. " -0.508 0.1752 103800.212 0.0113 0.0197 0. 12.0323 1.0166 0. 753 0.615 0.008 0.433 0.0138 0. 93 -4.886 0.464 0.0097 0.0281 1. 12. c1 cd -0.33 8.0108 0.472 0.109 0.995 0.30 6.0123 0.0254 1.0367 1.0214 0.712 0.89 -2.OllO 0.191 0.20 3.34 8.Dl06 0.34 8.0245 0. 771 0.0182 0.33 8.0151 1.34 9.364 0.068 0.932 0. 12.408 0.25 3.27 4.163 0.0144 0.94 -4.569 0.86 -1.160 0.0277 1.0179 0.0293 1.0248 l.0164 1.371 1.20 1.0138 0.323 1.37 c1 cd c1 cd -0.0115 0.198 0.40 Rn = " -6.0120 0.262 0. 727 0.0267 .272 0.39 11.82 0.34 9.22 2.914 0.32 7.20 1.31 6.0231 0.38 0.0201 0.133 1. " -7.0133 0.0270 SD7062-PT Fig.38 13.0264 0.0216 -0.388 1.91 -3.0150 1.35 8.812 0.24 4.0161 0.975 0.0133 0.0141 0.30 6.0297 1.26 4.0147 0.284 0.33 8.0161 0.Dl10 0.372 0.OllO O.39 11.398 0.187 0.502 0.0238 0.38 12.953 1.166 0.32 8.0135 0.0305 1.0148 0.0390 1.28 5.0370 1.906 0.94 -5.0125 0.0239 -0.20 2.0152 1.255 0.0167 1.87 -0.31 7.218 1.941 0.477 0.430 0.0410 204500.91 -3.0110 0.0308 1.740 0.39 -0.85 -0.0251 1.374 0. 795 0.0107 0.40 12.0173 1.88 -1.139 0.92 -1.0126 0.26 4.0131 0.31 7.237 0.0118 0.0176 0.864 0.24 3.0226 1.0512 1.082 0.25 3.34 9.0127 0.0201 1.25 4.0645 303400.115 -0.27 5.227 0.0191 0.375 0.208 0.0375 1.89 -2.38 11.0167 0.0171 0.927 0.334 0. 767 0.195 0.26 Rn = " c1 cd -0.33 8.34 8.379 0.39 Rn = " -4.841 0.91 -3. cl cd 0.0352 1.21 2.0124 0.300 0.92 -3.892 0.0151 0.30 7.0084 0.0115 0.31 7.243 0.25 4.35 9.95 -5.590 0.301 0.0173 0.0162 0.168 1. " -6.31 6. 735 0.533 0.37 10.95 -4.2ll8 0.034 1.21 1.0163 0.0185 1.0147 0.0183 0.0206 0.0102 0.276 0.0141 0.ll8 0.0106 0.88 -2.21 1.20 1.37 11.483 0.020 0.302 0.032 0.330 0.0180 0.522 0.85 -0.2228 149700.23 2.0127 0.331 0.30 6.321 1.0125 0.0162 0.0177 0.086 0.0175 1.534 0.28 5.0186 1.0212 1.0136 0.0095 0.0290 1.0137 0.419 0.437 0.88 -1.075 0.28 4.38 10. 765 O.83 0.83 0.0271 1.909 0.359 0.0114 O.099 0.35 8.605 0.0150 0.0253 1.88 -2.006 0.552 0.015 0.257 1. 0.31 6.0087 0.252 0.0215 1.483 0.164 0.37 10.94 -5.0101 0.20 1.0164 0. 770 0.0111 0.0183 1.24 3.27 5.31 6.0212 0.26 4.950 0.0144 0.0155 0.41 14.175 0.0112 0.0126 0.0147 0.0266 1. = " -6.83 0.748 0.93 -4.187 1.84 0. c1 cd -6.0122 0.037 0. 732 0.0159 0.139 Rn = 202900.0181 0.20 1.371 0.0140 1.014 0.36 Rn = " -3.0339 1.0160 0.19 1.88 -1.35 9.35 9.84 -0.0190 1.985 0.171 0.170 0.376 0.84 0.38 10.0372 1.84 -0.0460 154900.85 -0.350 0. " -3.0407 Rn = 204000.284 0.88 -1.695 0.24 3.203 0.29 5.90 -2.431 0.86 -1.631 0.0125 0.246 0.0225 1.0128 0.29 5.28 5.353 0.818 0.022 0.0197 -0.26 5.0101 0.141 Rn = 102600.0092.431 0.0355 SD7062-PT Fig.83 0.0126 0.25 4.33 c1 cd 0.0206 1.0099 0.022 1.570 0.36 9.317 0.0252 1.337 0.636 0.90 -2.37 10.111 1.049 0.33 8. -6.442 c.0216 0.0110 0.0221 1.0130 0.457 0.97 -5.0588 Rn = 305800.0206 0.28 6.0124 0.20 1.35 10.0101 0.37 10.0228 1.92 -4.089 0.864 0.716 0.316 0.336 0.572 0.90 -2.21 2.0100 0.256 0.29 6.113 1.30 6.0250 0.367 0.20 1.0139 0.33 9.0429 0.451 0.0115 O.89 -3.124 0.0131 O.29 5.0162 1.165 0.188 0.86 -1.476 0.40 11.364 0.25 3.84 -0.37 0.28 12.667 0.023 0.0206 0.88 -1.0343 1.590 0.081 0.30 7.485 0.107 0.30 6.22 2.0176 0.0197 1.34 8.22 2.0176 0.0170 0. 0.0141 0.623 O.0308 1.37 10.40 c1 -0.0133 0.0191 0.191 cd 0.40 12.0102 0.0141 0.35 10.18 1.35 9.91 -3.83 0.0176 -0.26 4.0253 -0.142 Rn 100800.057 0.00 -4.Dl68 0.084 1.36 11.28 5.0128 0.0095 0.0128 0.201 -0.0179 0.0128 0.0199 0.190 0.397 c.0113 0.497 0.659 0.097 0.23 2.0315 1.0184 1.0204 0.0145 0. c1 0.0366 Rn = 151700.00 -6.23 5.314 0.Oll5 0.91 -2.0261 1.0189 0.84 -0.87 -1.0152 0.40 12.933 1.37 11.33 7.0140 0.23 3.020 0.26 13. " -2. 12.075 0.0340 -0.219 0.013 0.465 0.0136 0.85 -0.0210 1.0252 0.0130 1.009 0.0305 1.82 -0.924 0.997 1.0181 SD7043-PT Fig.376 " c1 cd 0.307 0.066 0.0277 1.259 0.23 3.90 -3.357 0.0114 0.0183 0.694 0.83 0.25 3.37 c1 0. " cd c1 -2.28 5.0159 0.27 5.078 0.26 7.0159 0.0197 0.850 0.949 1.41 Rn = -0.0151 0.090 0.832 0.23 2.33 7.0252 1.390 Airfoils at Low Speeds 2.23 3.671 0.0189 1.0349 Rn = 303700.094 0.86 -1.82 0.204 -0. " -5.668 0.229 0.076 0.98 -3.38 11. 0.26 4.108 0.010 0.36 10. 790 0.859 0.383 0.30 6.0143 0.708 0.90 -2.539 0.328 0.117 0.117 -0.25 3.237 0.0155 0.094 0.0099 0.0291 1.299 1.0086 0.30 7.37 9.0207 1.159 0.0299 0.27 4.86 -1.045 1.359 0.0222 0.22 2.843 0.41 13.183 0.22 2.277 0.0103 0.32 7.0253 1.34 10.35 10.0120 0.42 -0.0241 1.0173 0.27 4.195 0.441 0.0231 0.369 0.38 11.04 75 1.644 0.288 0.33 7.35 9.458 0.1586 0.0139 0. 735 0.26 3.577 0.192 0.0223 0.249 1.870 0.93 -4.87 -2.95 -5.0215 1.87 -0.82 0.242 0.173 0.293 0.0218 0.284 0.83 0.0219 0.85 -0.0131 0.007 0.835 0.0154 0.147 0.678 0.0144 0.19 1.22 2.100 0.0470 210500.32 7.36 9.25 6.827 0. 072 0.0075 0.22 0.0443 SD7080-PT = c.27 7.29 6.36 8.89 0.28 8.0246 1.89 -2.031 0.218 1.38 1.0128 0.0106 0.26 4.0158 5.91 -0.0153 0.359 0.98 -0.538 0.131 0.116 -0.42 Rn 0.0206 0.999 0.813 0.29 0.91 0.12 0.0298 1.32 Rn = cd 0.677 0.481 0.86 0.216 0.465 0.31 10.964 0. c.20 1. c.97 -1.0222 -0.92 0.0115 -2.0098 0.90 -2.94 -0.85 -0.17 2.045 1.204 -6.33 Rn 0.864 0.41 1.95 c" 0. " -0.0691 1.0177 0.34 8.91 0.0283 0.0216 0.240 0.167 0.24 0.0634 0. = = c.93 -3.625 0.92 -3.570 0.0160 1.977 0.885 0.97 -2.82 15.0147 0.0144 0.36 10.046 0.439 1.607 0.25 6. c" " -0.0138 0.053 0.0508 1.14 1.0195 -0.02 -0.009 0.26 0.074 0.83 0.0155 0.0087 0.242 0.084 0.30 9.15 1.545 0.23 3.629 0.28 5.0130 0.30 9.0506 1.677 0.151 0.0114 0.0628 149800.31 10. = c. 705 0.0404 12.668 0.0182 1.35 9.180 -0.37 1.Q208 0.648 0.488 1.0250 -0.24 5.056 0.046 0.049 0.0602 204900.0133 0.21 2.146 Rn 100900.13 1.047 0.034 0.244 0.40 12.932 0.454 0. c.320 0.953 1.0127 0.91 -0.98 -3.0121 0.064 0.82 12.32 11.91 -3.0109 0.0141 -3.39 1.770 0.19 0.88 -1.26 7.555 0.188 0.997 0.41 12.43 c.376 1.27 5.587 = = c.171 0. -4.055 0. "' -0.0147 0.0158 -4.417 0.31 1.0130 0.0117 0.31 7.0069 0.29 6.228 0.30 9.389 0.24 5.388 0.37 10. 12.0210 0.37 9.285 -0.41 -0.0165 1.07 44 301800.0294 0.871 0.0139 1.0141 -2.0179 0. 776 0.23 6.435 0.95 -1.0130 2.0356 204500.0171 0.27 0.93 -1.0222 1.20 1.24 5.Chapter 13: Tabulated Data 11.003 0.470 0.0112 -1.0126 3.93 -0.951 0.470 0.0323 1.21 4.25 6.110 0.28 8.0348 0.0198 -2.94 -1. SD7084-PT c.0205 9.0267 0.023 0.0192 0.0113 0.00 -4.461 O.802 0.0125 0.535 0.18 3.95 -4.0097 0.973 0.305 0.0291 0.0197 0.0123 0.85 0.94 -4. 12.0114 0.0260 -0. 737 0.944 0.740 0.93 -0.0115 0.28 8.994 0.692 0.0411 SD7062-PT Fig.0348 1.372 0.0148 4.0298 0.274 0.092 -0.0146 0.0145 0.197 0.0181 -0.0230 1.Dl05 0.488 0.0176 1.440 0.391 0.0308 303000. 799 0.0192 0.27 7.95 -0.056 0.96 -2.00 -4.90 0.177 0.31 11.0128 0.0073 0.268 0.0191 0.552 0.31 10.089 0.88 0.36 9.0148 0.0129 1.41 11.22 0.11 0.0525 1.292 0.0179 -1.090 0.0296 12.17 2.33 13.299 1.33 1.Dl06 0.90 0.92 -0.0156 6.448 cd 0.868 O.0183 -5.25 6.89 0.961 0.0245 0.032 0.0175 -0.036 1.0191 8.0086 0.94 -5.98 -3.0144 -0.31 1.576 0.186 0.0303 9.129 0.133 1.31 10.88 -1.22 2.004 0.420 0.0673 Fig.19 4. 2.079 0.0081 0. 714 0.0197 1. a c.20 0. c" -0.026 0.0402 Rn 103500.143 0.30 9.0125 0.41 13. 721 0. " -0.647 0.30 -0.0194 5.143 Rn = 207600.106 0.0155 1.340 0.088 0.0407 1.0125 0.384 0.27 7.516 c" 0.144 Rn 62600.0238 6.20 0.159 0.0122 0. 935 0.029 0.0126 0.267 O.278 0.980 0.15 1.317 0.444 0.0083 0.0273 8.94 -1.14 1.817 0. a cd -3.173 0.29 0.527 0.0237 0.111 0.34 1.876 0.552 0.97 cd 0.29 8.243 0.115 -0.95 -2. cd -2.482 0. "' -0.0222 10.0168 6.0115 0.078 0.40 0.0181 1.0205 9.034 0.25 0.0136 -1.013 0.0121 0.0117 0.0135 0.0164 0.0304 1.31 6.0323 10.0153 0.0265 7.0138 3.268 0.196 0.0254 1.21 0. 12.0203 1.0185 3.924 0.86 -0.88 -2.222 0.96 -4.18 3.041 0.26 0.0384 0.200 -6.0126 0.007 0. -5.187 0.413 0.0475 0.21 4.24 3.96 -1.42 -0. 726 0.0101 0.046 0.84 0.94 -0. a cd -6.88 0.0111 0.0185 1.32 7.28 0.0095 391 .20 3.857 0.29 5.0140 0.31 13.40 12.0113 0.997 1.31 0.17 2.0094 0. cd "' -0.691 0.0128 0.32 13.0228 1.22 5.054 0.393 -6.0171 1. 726 0.2306 197500.0170 0. c.608 o.0244 11.0349 -6.058 0.31 Rn -0.0227 1.0154 1.0199 4.829 0.645 0.272 0.27 4.0102 0.24 6.14 1.0233 1.31 10.109 0.504 0.0140 0.0179 0.4 70 0.091 0.31 10.311 0.33 8.0129 0.0143 0.0129 0.353 0.830 0.0175 -0.0392 -0.32 13.91 -0.36 1.0123 0.17 0.0360 0.081 0.0112 0.32 12.044 0.0112 0.27 Rn = a -3.0137 0.0151 0.0500 Rn 152400.022 0.0400 1.96 -2.24 0.41 1.0252 10.0141 0.0145 0.0196 -1.93 -0.0182 3.0443 Rn = 302700. 790 0.0116 0.0159 4.0107 0.0113 0.0091 = c.409 0.21 0.036 1.14 2.39 10.155 0.1087 205700. " -0.39 11.252 -3.480 0.68 0.028 0. " -6.881 0.801 0.0184 2.30 9.042 0.0130 0.497 0.95 -2.076 0.515 = c.0102 0.78 9.048 1.803 0.32 1.0095 0.22 0.42 14.0099 0.32 12.74 6.0222 Fig.0298 1.34 7.0261 1.31 Rn = 4.300 0.0084 0.058 0.0136 0.0379 1.014 0.515 0.19 1.86 -1.41 13.20 4.42 13.18 Rn -0.231 0.156 -0.0182 0.0171 0.160 0. cd a -6.98 -0.95 -0.68 3.0139 -3.98 -0.570 0.280 0.79 0.351 0.371 0.079 0.533 0.0292 1.99 -3.0139 0.39 11.0171 0.0126 0.0191 1.822 0.371 0.28 8.0168 0.394 0.0191 7.0179 7.40 12.21 4.0143 0.17 3.917 0.0144 0.0158 0.361 0.226 0.16 2.84 0.0223 0.158 0.0122 0.31 11.73 0. cd -0.32 12.22 5.0348 1.26 7.917 0.0158 0.0128 -0.0207 1.0159 0.85 -0.381 0.034 0.149 -2.007 0.920 0.012 0.900 0.25 4.0115 0.439 0.0521 1.164 0.26 7. -5.82 1.23 3.0176 0. 98 -2.621 0.915 0.31 9.206 0.0073 0.044 0.585 0.984 1.187 0.984 1.0079 0.0259 1.667 0.25 7.00 cd 0.0113 0.095 0.15 1.88 0.70 5.046 1.0159 0.0077 0.429 0.136 -4.0088 0.0383 1.072 0.373 -5.0078 0.039 0.0228 1.0135 .Q118 0.357 0.30 9.30 0.0158 0.0112 0.91 -0.0085 0.093 0.0115 0.98 Rn -4.671 0.0109 0.25 6.0335 1.0163 -0.22 5.92 0.948 0.0298 1. " -0.171 0.92 0.98 -4.18 3.0145 0.91 -0. 0.11 1.0374 SD7090-PT Fig.26 7.045 0.19 4.0290 0.105 304000. " -0.28 8.915 0.0163 0.116 0. = c.30 0.064 0.0221 -4.0139 0.892 0.97 Rn -1.0135 -0.201 0.627 0.0162 0.0215 0. = c.876 0.0121 0.0280 0.0239 -0.24 5.577 0.99 -4.519 0.30 9.0152 0.0074 0.0103 0.576 0.321 0.30 10.12 1.140 0.31 10.487 0.0089 0.576 0.97 -1.150 Rn 302500.16 2.24 5.0265 0.92 0.0079 0.92 -0.0085 0.270 0.848 0.0092 0.0260 0.0114 0. " -0.0504 0.841 0.31 11.0205 0.951 0.0432 0.038 0.0249 0.017 0.0133 0.136 -0.0147 0.007 0.931 0.0293 Rn 304600.197 0.28 7.0173 0. c.0180 0.0116 0.02 -2.24 6.18 2.274 0.0106 0.354 -5.148 Rn 59600.0187 0.26 6.22 4.0092 0.13 1.91 -0.343 -5.191 -3.47 -0.0252 0.090 0.21 4.96 -0.0102 0.0126 0.0083 0.640 0.33 11.059 1.0207 0.26 7.994 0. 758 0.01 Rn -2.377 0.95 -0.283 0.18 2.916 0.0102 0.037 0.11 1.0094 0.478 0.27 8.28 8.28 8. 718 0.138 -0. 731 0.29 12.184 -3.0096 0.392 0.478 0.31 c" -0.0151 0.360 0.94 -1.576 0.669 0.0404 SD7090-PT Fig.480 0.0180 0.27 7.270 0.25 6.0091 0.272 0.31 11.0133 -4.91 -0.0118 -2.0113 0.0086 0.0194 0.92 -1.816 0. -0.247 -0.93 -1.130 -0.0101 0.072 1.0212 0.0139 0.0157 0.0101 -1.0128 0.064 0.083 0.478 0.24 5.18 3.0076 0.0158 -2.0162 0.0223 0.24 6.26 6.02 -4. -0.092 0.27 8.24 6.0240 0.91 0.0111 0.0081 0.23 6.88 0.16 2.32 10.93 -0.88 0.19 3.092 0.27 7.32 11.94 -0.32 10.20 4.96 -0. 772 0.0269 0.95 -2.0276 1.0453 0.280 O.685 0.1747 199100.0145 0.29 9.0094 0.95 -1.038 0.0103 0.0093 0.33 0.842 0.0102 0.211 0.21 5.00 c.107 0.30 10. 0.0097 0.186 -3.280 0.15 1.895 0.31 1.805 0.003 0.256 0.23 5.15 2.0085 0.31 9.651 0.29 9.004 0.161 0.140 0.30 9.0210 0.210 -4.004 0.148 0.359 -0.097 0.002 1.0115 0.22 4.200 -4.0086 0.0154 0. 758 0. 957 1.778 99100. 721 0.377 0.883 0. 757 0.043 0.31 1.23 6.0083 0.0084 0.27 7.0130 0.0078 0.95 -2.069 0.22 4.055 0.0145 0.082 0.94 -2.269 0.27 7.322 0.20 4.0078 0.126 0.670 0.96 -3.0248 1.77 8.0076 0.17 2.094 1.17 3.20 3.20 3.20 4.0125 0.840 0. " -0.0204 0.457 0.15 1.13 1.055 0. = c.381 0.34 11.00 Rn cd 0.87 0.361 -6.021 0. -0.0301 0.0118 0.94 -0.0240 10.381 0.Q157 -0.17 3.0171 0.0240 SD7090-PT Fig.1565 149000.17 2.310 0.29 cd 0.884 8.025 0.29 -0.27 7.067 0.624 0. " -0.13 2.0227 0.0297 1. -0.17 2.18 3.0107 0.16 1. 798 0.0146 -0.98 Rn -2.670 0.96 -3.0160 0.842 0.32 10.0104 0.1435 0.0182 0.0127 0.0165 0.0191 -0. = " -3.33 Rn = " -6.32 11.018 cd 0. 12.91 -0.16 2.530 0.94 -0.0108 0.284 0.804 0.0168 0.0097 0.17 3.31 10.759 0.957 0.0213 0.200 0. 717 0.756 0.22 5.96 -1.0084 0.960 1.PT Fig.0126 0.0171 0.0314 298400.0137 O.13 1.0096 0.0166 0.28 8.28 -0.138 -0.045 1.19 3.93 -1.29 9.590 0.98 -1.0210 1.88 0. = c. 722 0.004 0.31 10.061 0.0342 1.20 4.711 0.627 0.20 4.0126 0.00 -2.31 9.26 7.954 1. 12.297 0.0130 0.32 10.580 0.987 0. 12. = c. " -0.0104 0.25 6. = c.0077 0.272.33 cd 0.054 1.29 9.301 -0.31 10.0143 0.802 0.0092 0.13 1.0212 1.977 0.893 0.0118 0.984 0.95 -1.96 -3.88 0.0197 -0.0093 0.91 -0.33 Rn -2.23 5.91 0.152 Rn 62500.012 0.923 1. " -0.111 0.954 0.0072 0.99 -1.0125 0.15 1.32 0.418 0.185 0.24 6.178 0.457 0.029 0.0105 0.82 11.0096 0.399 0.0083 0.527 0.0085 0.093 0.362 0. = c.30 9.0252 1.0094 0.065 -0.556 0.95 -2.861 0. " -0.0110 0.98 Rn -2.0111 0.0118 0. " -0.0087 0. = c.0144 0.0075 0.489 0.0287 0.099 1.0134 0.0207 9.842 0.429 0.0182 0.0126 0.15 1.0442 298600.0113 0.0088 0.24 6.0224 0.0243 0.076 0.0195 0.326 0.0101 0.32 c" -0.18 3.549 0.15 2.132 0.18 3.26 7.20 c.0175 0.0164 SDSOOO.392 Airfoils at Low Speeds -0.0135 0.0099 0.32 = c.0096 0.15 2.0314 cd 0.092 0.29 8.33 12.0360 c" 0.0194 0.019 1.0128 0.049 0.821 0.228 -0.0138 0.0103 0.27 cd 0.203 0.29 9.090 0.29 8.530 0.0076 0.0312 1.092 0.0385 0.064 0.0165 0.26 8.0081 0.746 0.0156 0.22 5.28 8.149 Rn 304000.91 0.0147 0.0135 0.884 0.0121 0. " -0.0111 0.20 3.0911 0.17 2.22 4.103 300000.27 8.65 2.20 4.21 5. 12.096 -0.32 11.0126 -0.0099 0.106 0.23 5.116 0.116 cd 0.427 0.88 0.1960 cd 0.26 6.0085 0.90 0.0370 0.31 11. = c.0120 0.953 0.887 0.838 0.508 c" 0.0376 305900.90 -0.23 5.915 0.0139 0.0171 -3.91 0.0138 0. = " -4.0158 0.0243.0136 -0.25 7.432 0.0076 0.0084 -0. 0097 0.82 11.30 8.984 0.785 0.23 4.892 0.95 -1.0137 0.498 0.065 0.91 -1.0093 0.30 8.0354 302300.25 6.66 2.94 -2.88 0.885 0.0099 0.641 0. "' -0.0110 0. c.0264 1.021 0.227 0.915 0.30 8. c.30 9.0125 0.156 0.19 3.490 0.68 2. Rn = -3.21 4.0109 0.0099 0.21 3.0200 0.980 0.0444 0.160 0.91 -2.0097 0.33 Q c.0294 0.149 0.28 7.917 1. cd -0.126 0.0113 -0.0166 0.715 0.2198 200800.0129 0.0119 0.16 3.24 6.83 11.0268 1.81 11.889 0.097 0.0111 0.806 0.0112 0.17 2.532 0.1990 393 .0122 0.681 0.0095 0.0087 0.43 -0.0186 153800.0159 0.93 -1.076 0. c.0116 0.0231 1.0327 0.069 cd 0.599 0.34 0.140 0.0260 1.220 0.007 4 0.94 -1.25 5.0127 0.21 3.0106 0.546 0.34 Rn = " SDSOOO-PT cd 0.0132 0.104 0.0115 0.318 0.773 0.0600 0.93 -1.2233 0.31 Rn = " -3.076 0.0280 1.692 0.634 0.29 9.23 4.0116 0.98 -0.712 0.907 0. 971 0. 737 0.18 2.0187 0.0209 0.0076 0.92 -2.0105 0.15 2. c.22 4.744 0.0094 0.289 0.0125 0.888 0.0132 0.16 1.79 Rn = "' -2.0186 0.88 0.054 0.322 0.532 0.0147 0.24 5.32 Rn = c.0144 O.803 0.039 0.0073 0.69 .0107 0.18 2.995 202100.0127 0.147 -3.31 9.394 0.2499 306100.629 0.32 10.055 0.Ql03 0.905 0.27 6.197 0.089 0.0523 102500.25 6.035 1.0261 1.0148 0.682 0.078 0.0422 0.219 0.0171 0.34 Rn = "' -3.154 Rn= 99900.082 0.0341 0.13 1.0241 0.94 0.0326 1.0104 0.0113 -2.0104 0.0148 0.32 9.0586 SDSOOO-PT Fig.0131 0.89 -0.92 -2.074 -3.0093 0. 753 0.0095 0.21 3.0114 0.382 0.0130 0.0205 0.138 0.0133 0.946 0.27 7.28 Rn = Rn 0.69 3.24 5. " -0.497 0.0139 0.33 10. " -0.0175 0.056 0.736 0.94 -2. cd -0.90 -0. 0.85 0.823 0.89 -0.78 11.920 1.0192 0.79 13.669 0.18 2. 75 6. 12.791 0.18 2.23 4.0129 0.Ql06 0.16 1.73 5.0098 0.96 -2.23 5.959 1.640 0.67 2.0278 0.26 6.0376 150500. Gd -3.638 0.053 0.0081 0.17 1.204 0.26 Rn = "' -3.31 cd 0.0258 1.0108 0.0087 0.063 0. cd -0.0103 0.624 0.0302 1.153 Rn 104900.424 0. cd -0.23 5.28 7.0334 1.2004 0.101 0.26 5. 12.021 0.0123 0.0137 0.0086 0. 730 0.21 3.0144 0.23 4.074 -3.829 0.0169 0.924 0.0125 0.916 0.529 O.25 6.342 0.948 0.552 0.0442 311100.0169 0.87 0.0136 -0.091 0.72 5.0119 0.142 -3.0095 0.003 0.107 0.0181 1.0088 0.0147 O.140 0.0458 200000.339 0. cd -0.643 0.115 0.93 -2.34 Rn = "' -3.0149 0.31 9.28 7.536 0.100 0.386 0.1512 cd 0.685 0. = c.84 11.42 -0.0185 0.127 0.650 0.87 0.78 8.81 11.127 0.0539 154500.85 0.548 0.450 0.87 0.213 0.021 0.0104 0.0084 0. 12.958 0.14 1.0083 0.0163 0.0093 0.72 5.376 0.75 12.40 -0.992 1.062 0.20 3. 736 0.677 0.15 1.033 0. c.0157 0.69 3.507 0.070 1.90 -0.86 0.0369 1.0496 0.Chapter 13: Tabulated Data 3.0083 0.104 0.642 0.171 0.0143 -0. " -0.0150 0.26 6.088 -3.0213 1.94 -1.0387 206500.0109 0.31 0.21 4.0095 0.0324 1. = c.28 7.992 0.80 9. 771 0.0125 O.870 0.91 -1.0216 1.0323 1.37 0.155 Rn 101300.491 0.73 5.74 6.066 0.0090 0.23 4.87 0.13 1.93 -2.892 0.0126 0.30 8.0096 0.26 6.1473 301300.28 8.22 4.33 10.97 O.0084 0. -0.226 0.0071 1.26 6.344 cd 0.0126 0.424 0.454 0.0124 0.020 -2.25 6. " 0.0231 9.105 0.38 0.94 -2.0364 1.121 1.057 0.12 1.28 7.32 10.126 0.0101 Fig.0186 0.909 0.0126 0.26 6. 3.071 0.896 0.15 0.053 0.0104 0.990 0.0207 0.26 7.18 3.15 1.80 9.18 3.76 12.24 = cd 0.0087 0.0123 0.93 0.540 0.901 0.0121 0.0084 0.036 0.2143 0.Dl78 0.0395 0.046 0.26 7.0133 0.085 0.99 -1.001 0.823 0.30 8.75 6. 0.0208 0.0116 0.0120 0.89 -0.27 7.17 2.0078 0.133 0.27 7.188 0.809 0.313 0.31 9.385 0.0243 1.0218 1. c.21 4.79 8.0140 0.0258 1.0104 0.25 5.080 0.0093 0.058 0. "' -0.0094 0.0106 0. c.42 -0.0153 0.0541 0.0187 1.0141 0.0109 -1.525 0. " -0.828 0.0096 0.0089 0.25 5. cd 0. c.936 0.31 9.66 2.81 11. " 0.448 0.32 9.826 0.432 0.0098 0.824 0.0166 0.81 cd 0.058 0.0079 0.531 0. c.93 -2.001 1.93 -1.Dl8 0.334 0.30 8.29 7.620 0.122 0.546 O.0278 0.33 Rn = "' -3.0208 1.18 3.17 2.0236 1.0138 0.0234 SDSOOO-PT Fig.0149 -0.31 9.0132 0.32 10.048 -2.0256 0.0286 -0.429 0.30 8.87 0.0110 0.0158 0.20 3.21 4.009 cd 0.80 9.251 0. 768 0.0193 1.78 Rn = Gd 0.133 0.155 1.79 8.93 -1.818 0.30 Rn = c.40 0.28 7.1782 152200.86 0.0095 0.0140 0.053 0.74 6.91 -0.88 -0.Dl72 0.0128 0.638 0. 765 0.76 8.004 0.33 1.071 0.0326 1.0229 0.32 9.909 0.041 -2.849 0. c.035 1.833 0.0088 0.931 0.39 0.0162 0.0091 0.16 1.69 3.0111 0.0095 0.0097 0. 906 0.250 0.211 0.0163 0.0086 0.934 1.Dl3 0.23 4.96 -2.27 6.043 0.33 10.350 0.81 9.0180 0.193 0.0132 0.0104 0.583 0.116 0.078 0.975 0.022 0.296 0.26 6.16 1.0135 0.0110 0.21 4.91 -1.25 5.20 3.354 0. c.308 0. 94 0.569 0.21 8.0085 0.162 Rn 61300.0227 0. " -0.0163 0.07 0.681 0.0122 0.0124 0.0158 -0.689 5.542 0.0436 11.0077 0. Rn cd -0.22 7.87 0.0354 ·0.098 0.0142 SD8020·PT Fig.269 0. 77 4 0.0163 9.81 11.23 10. SPICA·PT Fig.090 0.0221 = " = "' c.88 0.24 0.0208 1.241 0.99 ·1.73 6.0087 0.22 0.0629 204900.334 0.0204 1.0240 0.98 -1.26 6.080 0.329 0.97 -0.0387 -1.558 0. 926 0.499 0.0334 2.91 -0.0192 0.0179 -0.02 -4.11 2.0139 6.575 0.0340 1.1224 0.02 -4.0347 0.640 0.96 -0.0141 0.068 0.30 8.0282 1.199 0.0134 -0.278 0.157 Rn 59800.15 0.1609 151200.069 0.064 0.1373 0.793 0.0080 0.0126 -0.349 0.518 0.63 3.0593 Rn 202300.0360 3.72 0.19 0.25 Rn = "' -6.450 0.0397 Rn 102300.0209 4.0081 0.1579 100500. Ct -0.299 -3.24 = " -6.15 0.18 6.0303 0.42 0.88 0.27 10. Ct -0.23 12.448 0.0167 0.808 8.093 0.803 6. 12. Ct ·2.0119 4.0108 0.0257 -0.0073 0.761 0.062 0.28 0.95 -0.0355 1.006 0.0272 -0.836 7.0156 3.0157 6.12 1.0172 0.048 0.0299 1.23 8.244 0.0288 0.0260 0.959 0.159 Rn 62100.0115 0.229 0.239 0.437 0.547 0.0340 0.340 0.0111 0.88 0.067 0.143 0.369 0.011 0.537 0.0101 0.117 0. Ct c. Ct -0.0189 4.18 0.219 0.073 0.0384 1.317 0.93 0.16 4.884 0.0318 0.352 0.24 0.289 0.28 10.0293 1.0132 0.0127 0.0422 0.218 0.0258 0.0237 9.18 5.655 0.0136 -0.0122 0.0175 0.21 0.81 11.688 0.666 0.34 10.0179 3.10 2.39 0.582 0.0123 -0.0100 0.0154 4. 730 0.32 9.0584 301800.69 6.188 0.085 0.81 1.26 0. Ct -2.19 5.l7 0.79 -0.0274 1.0196 0.39 14.0098 0.219 0.428 0.26 7.0207 13. c.929 0.566 0.0147 -0.94 0.590 0.93 0. 12.184 0.426 0.27 0.0114 3.30 0.03 ·4.20 7.0098 0.Oll8 -0.0140 0.l9 0.28 10.73 0.0299 10.0083 -0.20 0.98 -0.077 0.0165 ·1.501 0.082 0.27 0.0135 Rn -1.0217 -1.214 O.0267 1.0107 -0.123 0.0153 8.754 0.12 3.821 0.021 0.631 0. WB135/35-PT Fig.071 0.0167 -0.75 9.0247 0. Ct c.518 3.96 ·1.326 0. 0.394 0.143 0.008 0.0196 0.816 0.412 0.23 0.094 -0.17 5.22 0. 12.809 0.26 10.15 4.658 O.0117 3.1627 0.839 0.38 12.96 -0.0113 0.489 0.09 2.0095 0.34 1. 723 0.564 0.61 1.0144 0.044 0.19 0.39 13.0241 0.579 0.32 1.056 0.0252 0.0239 0.242 0.028 0.187 -2.0159 -0.356 0.183 0.583 0.047 ·1.485 0.22 7.0097 0.284 0.51 -2.0244 7.817 0.249 0.0245 1.0196 3.27 10.26 7.0186 0.0212 2. 799 0.30 10.14 4.18 0.19 4.26 11.539 0.26 0.19 6.0285 10.0112 0.29 0.1304 306700.0923 c.25 c.0186 11.585 0.09 1.0128 -0.0344 0.825 0. c.0150 0.0166 -0.922 0.0385 Rn 301500.024 0.0321 -0.23 0.28 11.0137 0.23 8.01 -2.0132 -0.94 0.0313 0.0331 0.0185 1.009 0.21 7.22 0.615 0.0255 1.23 0.0090 0.0114 -0.0204 0. Ct -0.0119 0.413 0.0266 0.24 9.05 Rn -5.840 0.666 0.636 0.21 7.0199 0.0096 0.245 0.0168 0.599 0.90 0.30 1.66 3.015 0.468 0.134 0. Ct -2.684 0.227 0.813 0.063 0.684 0.0269 0.0122 0.393 0.99 -1. "' -0. "' = "' c.682 0.98 -1.0101 -0.0096 0.42 -0.0407 1.616 0.16 0.042 0.0171 0.0093 0. 734 0.96 Rn ·1.23 0.24 11.41 1.394 0.412 0.0123 -0.1359 c.160 Rn 59900.533 0.01 -2.16 0.611 4.66 0. 734 0.40 0.0176 10.0139 1.410 0.0108 0.0142 0.0403 -1.0194 12.Gl05 5.81 0.0191 7.1469 cd c.0100 4.075 0.0145 0.39 1.21 4.09 1. Ct -0.250 0.410 0.513 0.0117 0.692 0.40 0.0126 0.920 0.0262 9.740 0.20 6.0545 1. Ct -1.37 1.0112 = c.007 4 0.341 0. = "' -6.0679 = "' = "' = 150300.16 4.373 -4.0258 0.08 1. 12.658 0.14 3.0577 -2.27 7.27 0.01 -2.85 0.011 0.0090 -0.126 0.03 -5.89 0.443 0.22 0.61 0.20 6.78 9.015 0.00 -2.40 0.362 0.0328 0.496 0.150 0.0437 4.0619 0.0102 -0.329 l.388 0.82 = " -2.16 5.0103 -0.67 3.35 1.0155 0. Ct -5.14 1.0415 0.926 0.12 2. 5.0101 0.80 12.13 0.0085 0.058 0.23 = "' -6.413 0.0197 -0.0380 1.0128 0.81 1.0166 0.006 0.90 0.02 -4.426 2.541 0.27 11. Ct -3.0268 0.20 0.0209 1.13 3.107 0.0073 0.555 0.003 0.643 0.13 1.420 0.137 0.78 0.22 0.736 = "' c.85 0.87 0. Ct ·1.93 -0.0151 0.24 9.25 7.74 6.649 0.335 0.41 15.0142 0.319 0.25 0.05 ·4.22 9.155 0.536 0.837 0.426 0.73 6.0296 0.1121 .0281 0.0268 1.394 Airfoils at Low Speeds = 201100.167 0.442 0.738 0.0120 0. c.16 0.224 0.0149 2.63 3.0128 0.071 0. c.67 0.0128 5.77 0. 793 0.0327 1.16 0.087 0.605 ·6.06 Rn ·5.44 0.81 11.29 10.91 -0.0188 6. SD8040-PT Fig.79 9.21 0.0231 1.0129 1.98 -0.664 0.815 0.19 4.0237 14.27 0.09 1.40 1.99 -1.176 0.660 0.0080 -0.96 -0.95 -0.35 11.81 Rn = "' -2.479 0.0105 0.18 0.26 0.82 1.278 0.0468 0.052 0.172 0.402 0.25 9.014 0.0204 -0.0112 0.249 0.0825 0.12 3.0128 2.077 0.256 0.0143 7.419 0.15 0.0163 l.513 0.24 8.15 4.065 0. -5.20 0.0729 Rn 99600.407 0.79 9.96 -0.936 0.0238 0.0327 1.0626 6.047 -2.0407 0.0521 5.96 -0.771 0.747 0.0113 0.352 0.0575 0.312 0. 95 -4.212 0.259 0.0108 0.29 6.226 0.514 0.104 0.90 -1.98 -1.805 0.35 9.1035 14.043 0. " -5.94 -3.85 0.24 4.17 1.22 4.0165 0.24 4.117 0.155 0.24 5.18 0.085 1.0129 0.22 3.0178 1.0182 0.19 2.0533 204000.87 0.0116 0.154 0.31 8.86 -0. "' c.0224 1.0462 Rn = 200500.30 -0.95 -0.0174 -0.0125 0.0166 0.36 10.0157 0.0249 0.0113 -0.0099 0.0278 0.91 0.730 0.0157 0.163 -0.19 2.334 0.97 -3. a c.894 0.36 Rn = " -9.0158 0.29 7.0297 0.057 0.335 0.228 0.020 0.0312 1.800 0.0281 1.0196 1.35 1.047 0.0135 0.0180 1.88 0. c.0772 Rn = 155700.973 1. c" -0.18 0.0114 0.076 1.32 12.410 0.019 0.114 0.32 8.231 0.29 6.0106 3.895 0.85 0.34 9.648 0.0128 0.32 -0.0204 1.116 0.536 0.27 5.33 8.0228 1.896 0.97 -2.127 0.35 10.26 5.178 0.638 0.28 6.069 0.30 7.151 0.0270 0.0135 0.711 0.0143 0.92 -3.303 0.822 0.21 3.001 0.0576 1.30 7.438 0.0140 0.011 0.0131 0.133 0.511 0.107 -0.18 2.0143 0.0135 0.383 0.434 0.021 0.23 8.28 7.26 5.89 -2.0314 1.0687 1.0118 -3.0210 395 .230 0.0326 1.23 3.89 -0.1316 Rn = 100500.0112 -1.85 0.35 c.228 0.0227 0.110 0.096 0.91 -1.102 0.94 -0.95 -2.029 0.190 0.088 0.27 0.33 10.32 11.0253 1.0123 0.0196 0.0138 0.37 12.0154 0.19 1.0122 0.330 0. 728 0.125 0.0162 0.99 -6.23 3.28 6.0226 0.92 -3.416 0.27 6.34 1.816 0.0106 2.603 0.85 0.429 0.36 c.89 -0.0122 0.0170 -4.0274 -0. 738 0.193 0.0194 -0.196 -0.24 5.274 0.0280 -0.26 7.06 -6.189 0.0120 0.91 -2.0173 0.0443 1.0132 0.0281 0.287 -0.0212 0.016 0.34 11.07 -8.168 1.0252 1.512 0.23 4.88 0.0551 1.0146 0.0480 1. 702 0.90 -2.814 0.0148 0.930 0.93 -1.0116 0.510 -0.0180 0.095 0.15 1.095 0.35 1.0169 0.90 -2.17 1.083 0.051 0.Chapter 13: Tabulated Data 9.29 9.32 10.0224 1.0177 0.636 0.079 0.96 -4.17 1.31 7.0308 0.0216 1.20 3.1266 Rn 302700.608 0.0152 0.271 0.078 0.164 Rn = 98500.34 9.604 0. "' -5.31 8.0625 153900.0116 -2.0177 1.157 0.34 10.0231 -0.155 0.239 c" 0.0623 12.0113 0.403 0.0114 0.094 0.22 4.0220 0.388 0.984 0. -6.01 -4. c" a c.33 9. -0.0160 0. 12.076 0.0114 = c.88 -0.499 0.34 9.20 2.908 0.87 -0.21 2.0111 0.406 -0.0170 0.0141 0.0131 0.0356 11.33 8.84 0.0153 0.987 0.0109 0.0123 0.0109 0.22 0.219 0.20 3.31 Rn = -5.34 11.21 2.28 6.627 0.128 0.0287 0. cd -6.02 -0.0182 0.063 0.37 13.0154 1.0289 -0.18 2.031 1.24 4.0169 1.549 0.97 -4.0100 0.0200 0.0144 0.0183 0.012 0.85 0.0426 1.089 0.0263 1.0168 0.0140 0.0143 1.0256 1.24 4.34 10.87 -0. Fig.23 3.0186 0.052 0.114 0.97 -3.89 -1.91 -0.125 0.526 0.18 1.0140 0.0290 0.175 0.201 0.235 0.070 0.0207 0.0122 0.35 1.332 0.0142 0.0227 1.0136 0.0149 0.287 0.0238 -0.0104 0. c" -0.839 0.607 0.152 0.0195 0.0653 Rn = 307600.298 0.00 -4.33 8.0180 0.505 0.24 5.0148 -3. c" " c. cd -5.0138 0.923 0.19 0.94 -1. 746 0.0109 1.0228 1. 725 0. c" WB140/35/FB-PT.211 0.33 -0.198 0.982 0.39 11.0807 13.228 1.224 0. 700 0. a -4.94 -2.182 0.0280 0.32 12.210 0. 799 0.0237 1.26 6.0264 0.95 -4.093 0.0131 0. -0.0104 0.097 0.95 -3.29 7.17 1.137 0.27 5.0130 0.34 13.0163 0.0289 1.132 0.999 1.32 9.088 0.900 0.305 0.85 0. 719 0.89 -1.293 0.0144 0. 396 Airfoils at Low Speeds . I I I I I I I I I I I I . 398 Airfoils at Low Speeds .