XFOIL Version 6.96 Calculated polar for: GOE 620 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -18.500 -0.6320 0.10296 0.10033 -0.0934 0.9956 0.0273 -18.250 -0.6559 0.09478 0.09203 -0.0990 0.9936 0.0276 -18.000 -0.6728 0.08834 0.08549 -0.1032 0.9913 0.0277 -17.750 -0.6877 0.08209 0.07915 -0.1076 0.9892 0.0279 -17.500 -0.6986 0.07639 0.07335 -0.1120 0.9867 0.0281 -17.250 -0.7078 0.07090 0.06778 -0.1165 0.9841 0.0282 -16.750 -0.7321 0.06011 0.05680 -0.1251 0.9760 0.0283 -16.500 -0.7422 0.05482 0.05140 -0.1298 0.9711 0.0285 -16.250 -0.7506 0.04931 0.04578 -0.1355 0.9664 0.0286 -16.000 -0.7598 0.04354 0.03989 -0.1418 0.9582 0.0287 -15.750 -0.7581 0.03800 0.03423 -0.1495 0.9503 0.0289 -15.500 -0.7465 0.03290 0.02899 -0.1581 0.9403 0.0292 -15.250 -0.7188 0.02881 0.02476 -0.1675 0.9324 0.0294 -15.000 -0.6879 0.02541 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9.500 1.3504 0.02085 0.01481 -0.0804 0.1901 1.0000 9.750 1.3628 0.02177 0.01568 -0.0789 0.1824 1.0000 10.000 1.3721 0.02290 0.01675 -0.0772 0.1682 1.0000 10.250 1.3653 0.02509 0.01874 -0.0740 0.1350 1.0000 10.500 1.3735 0.02638 0.02000 -0.0723 0.1275 1.0000 10.750 1.3856 0.02744 0.02106 -0.0711 0.1240 1.0000 11.000 1.3954 0.02868 0.02230 -0.0698 0.1197 1.0000 11.250 1.4088 0.02969 0.02332 -0.0688 0.1178 1.0000 11.500 1.4218 0.03073 0.02440 -0.0678 0.1168 1.0000 11.750 1.4339 0.03186 0.02555 -0.0668 0.1153 1.0000 12.000 1.4447 0.03310 0.02681 -0.0658 0.1134 1.0000 12.250 1.4546 0.03445 0.02816 -0.0647 0.1115 1.0000 12.500 1.4650 0.03575 0.02949 -0.0637 0.1098 1.0000 12.750 1.4744 0.03716 0.03093 -0.0627 0.1081 1.0000 13.000 1.4843 0.03855 0.03234 -0.0617 0.1063 1.0000 13.250 1.4960 0.03982 0.03366 -0.0610 0.1059 1.0000 13.500 1.5077 0.04109 0.03498 -0.0602 0.1053 1.0000 13.750 1.5190 0.04240 0.03633 -0.0595 0.1045 1.0000 14.000 1.5287 0.04388 0.03786 -0.0588 0.1038 1.0000 14.250 1.5385 0.04534 0.03935 -0.0580 0.1027 1.0000 14.500 1.5473 0.04694 0.04099 -0.0573 0.1021 1.0000 14.750 1.5555 0.04859 0.04269 -0.0566 0.1010 1.0000 15.000 1.5624 0.05039 0.04452 -0.0559 0.0998 1.0000 15.250 1.5688 0.05227 0.04644 -0.0552 0.0985 1.0000 15.500 1.5742 0.05426 0.04847 -0.0545 0.0974 1.0000 15.750 1.5780 0.05642 0.05068 -0.0538 0.0956 1.0000 16.000 1.5844 0.05836 0.05267 -0.0533 0.0948 1.0000 16.250 1.5933 0.06007 0.05443 -0.0529 0.0941 1.0000 16.500 1.6009 0.06190 0.05631 -0.0525 0.0933 1.0000 16.750 1.6070 0.06393 0.05840 -0.0522 0.0920 1.0000 17.000 1.6128 0.06597 0.06048 -0.0518 0.0900 1.0000 17.250 1.6159 0.06833 0.06289 -0.0515 0.0885 1.0000 17.500 1.6183 0.07077 0.06536 -0.0511 0.0864 1.0000 17.750 1.6193 0.07343 0.06807 -0.0509 0.0848 1.0000 18.000 1.6239 0.07567 0.07037 -0.0507 0.0830 1.0000 18.250 1.6250 0.07836 0.07309 -0.0505 0.0791 1.0000 18.500 1.6213 0.08158 0.07631 -0.0504 0.0736 1.0000 18.750 1.6105 0.08569 0.08038 -0.0504 0.0649 1.0000 19.000 1.5998 0.08988 0.08458 -0.0505 0.0597 1.0000