XFOIL Version 6.96 Calculated polar for: Eiffel 10 (Wright) - 1903 Wright Flyer airfoil 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.500 -0.3856 0.10595 0.10211 -0.0090 1.0000 0.0136 -8.250 -0.3774 0.10316 0.09935 -0.0101 1.0000 0.0137 -8.000 -0.3692 0.10050 0.09671 -0.0112 1.0000 0.0138 -7.750 -0.3605 0.09786 0.09412 -0.0126 1.0000 0.0139 -7.500 -0.3492 0.09510 0.09139 -0.0146 1.0000 0.0140 -7.250 -0.3370 0.09233 0.08866 -0.0167 1.0000 0.0142 -7.000 -0.3241 0.08957 0.08593 -0.0191 1.0000 0.0143 -6.750 -0.3107 0.08684 0.08324 -0.0214 1.0000 0.0144 -6.500 -0.2973 0.08418 0.08062 -0.0237 1.0000 0.0146 -6.250 -0.2852 0.08168 0.07818 -0.0255 1.0000 0.0147 -6.000 -0.2672 0.07885 0.07538 -0.0288 0.9971 0.0149 -5.750 -0.2302 0.07481 0.07131 -0.0366 0.9863 0.0152 -5.500 -0.1884 0.07067 0.06713 -0.0452 0.9669 0.0156 -5.250 -0.1454 0.06678 0.06314 -0.0537 0.9330 0.0158 -5.000 -0.1071 0.06351 0.05977 -0.0609 0.9081 0.0159 -4.750 -0.0741 0.06033 0.05651 -0.0663 0.8881 0.0161 -4.500 -0.0573 0.05740 0.05350 -0.0665 0.8587 0.0164 -4.250 -0.0330 0.05489 0.05077 -0.0687 0.8098 0.0170 -4.000 -0.0036 0.05232 0.04795 -0.0720 0.7616 0.0176 -3.750 0.0294 0.04969 0.04509 -0.0760 0.7188 0.0181 -3.500 0.0641 0.04747 0.04230 -0.0799 0.6104 0.0185 -3.250 0.0974 0.04749 0.04026 -0.0845 0.0356 0.0187 -2.750 0.1642 0.04234 0.03499 -0.0910 0.0313 0.0193 -2.500 0.1992 0.04010 0.03269 -0.0942 0.0303 0.0199 -2.250 0.2371 0.03793 0.03043 -0.0976 0.0280 0.0208 -2.000 0.2811 0.03578 0.02812 -0.1016 0.0258 0.0217 -1.750 0.3170 0.03370 0.02597 -0.1042 0.0251 0.0222 -1.500 0.3497 0.03206 0.02428 -0.1062 0.0246 0.0234 -1.250 0.3932 0.03018 0.02213 -0.1088 0.0243 0.0253 -1.000 0.4220 0.02894 0.02091 -0.1102 0.0241 0.0278 -0.750 0.4588 0.02761 0.01937 -0.1118 0.0240 0.0339 -0.500 0.4905 0.02669 0.01832 -0.1129 0.0240 0.0395 -0.250 0.5217 0.02594 0.01742 -0.1137 0.0242 0.0461 0.250 0.5818 0.02576 0.01684 -0.1142 0.0248 0.0700 1.000 0.6757 0.02325 0.01349 -0.1139 0.0260 0.0323 1.250 0.7042 0.02333 0.01330 -0.1134 0.0266 0.0267 1.500 0.7326 0.02345 0.01320 -0.1128 0.0274 0.0242 1.750 0.7612 0.02348 0.01324 -0.1121 0.0283 0.0257 2.000 0.7880 0.02447 0.01399 -0.1117 0.0272 0.0259 2.250 0.8155 0.02489 0.01451 -0.1110 0.0261 0.0252 2.500 0.8445 0.02556 0.01536 -0.1101 0.0298 0.0251 2.750 0.8740 0.02736 0.01723 -0.1093 0.0372 0.0253 4.250 1.0292 0.03468 0.02691 -0.1040 0.0366 1.0000 5.250 1.0736 0.03127 0.02459 -0.0923 0.0287 1.0000 5.500 1.0912 0.03393 0.02735 -0.0912 0.0280 1.0000 5.750 1.1071 0.03708 0.03054 -0.0903 0.0274 1.0000 6.250 1.1338 0.04414 0.03832 -0.0863 0.0259 1.0000 6.500 1.1453 0.04769 0.04214 -0.0843 0.0252 1.0000 6.750 1.1543 0.05149 0.04616 -0.0824 0.0247 1.0000 7.000 1.1606 0.05545 0.05033 -0.0805 0.0243 1.0000 7.250 1.1643 0.05953 0.05460 -0.0786 0.0240 1.0000 7.500 1.1649 0.06369 0.05894 -0.0765 0.0238 1.0000 7.750 1.1620 0.06793 0.06334 -0.0743 0.0236 1.0000 8.000 1.1548 0.07217 0.06774 -0.0720 0.0235 1.0000 8.250 1.1419 0.07636 0.07211 -0.0694 0.0235 1.0000 8.500 1.1182 0.08019 0.07607 -0.0659 0.0235 1.0000 8.750 1.0887 0.08482 0.08087 -0.0639 0.0236 1.0000 9.000 1.0555 0.09076 0.08697 -0.0641 0.0238 1.0000 9.250 1.0207 0.09822 0.09460 -0.0666 0.0241 1.0000