XFOIL Version 6.96 Calculated polar for: GOE 10K 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 = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -7.750 -0.5301 0.09049 0.08699 -0.0111 1.0000 0.0238 -7.500 -0.5321 0.08759 0.08415 -0.0117 1.0000 0.0240 -7.250 -0.5321 0.08422 0.08083 -0.0135 1.0000 0.0253 -7.000 -0.5282 0.08060 0.07723 -0.0161 1.0000 0.0256 -6.750 -0.5195 0.07668 0.07330 -0.0211 1.0000 0.0271 -6.500 -0.5075 0.07262 0.06919 -0.0256 1.0000 0.0275 -6.250 -0.4918 0.06860 0.06505 -0.0298 1.0000 0.0278 -6.000 -0.4906 0.06328 0.05975 -0.0293 1.0000 0.0289 -5.750 -0.4844 0.06023 0.05671 -0.0280 1.0000 0.0307 -5.500 -0.4713 0.05660 0.05300 -0.0290 1.0000 0.0330 -5.250 -0.4446 0.05249 0.04848 -0.0325 1.0000 0.0366 -5.000 -0.4340 0.04784 0.04377 -0.0323 1.0000 0.0376 -4.750 -0.4222 0.04489 0.04083 -0.0313 1.0000 0.0405 -4.500 -0.3984 0.04140 0.03681 -0.0318 1.0000 0.0455 -4.250 -0.3861 0.03794 0.03344 -0.0309 1.0000 0.0482 -4.000 -0.3623 0.03573 0.03055 -0.0301 1.0000 0.0560 -3.750 -0.3472 0.03194 0.02691 -0.0294 1.0000 0.0590 -3.500 -0.3263 0.02928 0.02388 -0.0284 1.0000 0.0669 -3.250 -0.3051 0.02692 0.02120 -0.0273 1.0000 0.0758 -3.000 -0.2847 0.02476 0.01891 -0.0263 1.0000 0.0883 -2.750 -0.2492 0.01927 0.01231 -0.0234 1.0000 0.0339 -2.500 -0.2215 0.01689 0.00936 -0.0216 1.0000 0.0275 -2.250 -0.1954 0.01489 0.00704 -0.0205 1.0000 0.0271 -2.000 -0.1707 0.01331 0.00537 -0.0195 1.0000 0.0308 -1.750 -0.1470 0.01241 0.00438 -0.0183 1.0000 0.0366 -1.500 -0.1238 0.01130 0.00346 -0.0171 1.0000 0.0850 -1.250 -0.0651 0.00811 0.00309 -0.0240 1.0000 1.0000 -1.000 -0.0430 0.00817 0.00293 -0.0229 1.0000 1.0000 -0.750 -0.0209 0.00824 0.00282 -0.0219 1.0000 1.0000 -0.500 0.0012 0.00833 0.00274 -0.0208 1.0000 1.0000 -0.250 0.0232 0.00843 0.00273 -0.0199 1.0000 1.0000 0.000 0.0452 0.00855 0.00276 -0.0189 1.0000 1.0000 0.250 0.0672 0.00868 0.00281 -0.0180 1.0000 1.0000 0.500 0.0892 0.00882 0.00290 -0.0171 1.0000 1.0000 0.750 0.1112 0.00897 0.00303 -0.0162 1.0000 1.0000 1.000 0.1331 0.00914 0.00320 -0.0153 1.0000 1.0000 1.250 0.1549 0.00932 0.00340 -0.0144 1.0000 1.0000 1.500 0.1766 0.00952 0.00364 -0.0136 1.0000 1.0000 1.750 0.2046 0.00976 0.00394 -0.0142 0.9982 1.0000 2.000 0.2735 0.00997 0.00437 -0.0231 0.9839 1.0000 2.250 0.3648 0.00939 0.00418 -0.0357 0.9557 1.0000 2.500 0.4677 0.00993 0.00278 -0.0479 0.3329 1.0000 2.750 0.4751 0.01246 0.00392 -0.0442 0.0406 1.0000 3.000 0.4953 0.01353 0.00498 -0.0423 0.0292 1.0000 3.250 0.5162 0.01489 0.00644 -0.0403 0.0267 1.0000 3.500 0.5396 0.01643 0.00810 -0.0388 0.0241 1.0000 3.750 0.5659 0.01866 0.01057 -0.0374 0.0244 1.0000 4.000 0.5926 0.02166 0.01392 -0.0360 0.0279 1.0000 4.250 0.6321 0.02807 0.02154 -0.0314 0.0903 1.0000 4.500 0.6511 0.03071 0.02449 -0.0295 0.0800 1.0000 5.250 0.7012 0.03918 0.03384 -0.0237 0.0532 1.0000 5.500 0.7166 0.04219 0.03700 -0.0223 0.0483 1.0000 5.750 0.6916 0.03203 0.02718 -0.0183 0.0464 1.0000 6.000 0.7028 0.03499 0.03087 -0.0155 0.0413 1.0000 6.250 0.7534 0.05578 0.05083 -0.0192 0.0379 1.0000 6.500 0.7633 0.05721 0.05315 -0.0163 0.0347 1.0000 6.750 0.7721 0.06095 0.05705 -0.0155 0.0323 1.0000 7.000 0.7789 0.06448 0.06062 -0.0147 0.0307 1.0000 7.250 0.7808 0.07124 0.06723 -0.0144 0.0296 1.0000 7.500 0.7824 0.07486 0.07112 -0.0138 0.0295 1.0000 7.750 0.7822 0.07885 0.07532 -0.0143 0.0292 1.0000 8.000 0.7803 0.08320 0.07980 -0.0160 0.0289 1.0000 8.250 0.7746 0.08742 0.08407 -0.0177 0.0287 1.0000 8.500 0.7672 0.09177 0.08843 -0.0201 0.0286 1.0000