XFOIL Version 6.94 Calculated polar for: GOE 286 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.3891 0.01138 0.00308 -0.0657 0.5987 0.0501 0.500 0.4456 0.01121 0.00299 -0.0658 0.5674 0.0751 1.000 0.5004 0.00874 0.00287 -0.0657 0.5308 1.0000 1.500 0.5557 0.00903 0.00280 -0.0654 0.4856 1.0000 2.000 0.6105 0.00944 0.00288 -0.0652 0.4463 1.0000 2.500 0.6651 0.00990 0.00306 -0.0650 0.4156 1.0000 3.000 0.7194 0.01038 0.00330 -0.0649 0.3864 1.0000 5.000 0.9208 0.01446 0.00564 -0.0631 0.1074 1.0000 5.500 0.9730 0.01507 0.00625 -0.0627 0.0978 1.0000 6.000 1.0212 0.01620 0.00713 -0.0619 0.0454 1.0000 6.500 1.0675 0.01749 0.00826 -0.0607 0.0067 1.0000 7.000 1.1157 0.01841 0.00929 -0.0597 0.0058 1.0000 7.500 1.1624 0.01943 0.01048 -0.0584 0.0058 1.0000 8.000 1.2070 0.02059 0.01182 -0.0569 0.0061 1.0000 8.500 1.2483 0.02196 0.01342 -0.0549 0.0065 1.0000 9.000 1.2841 0.02367 0.01537 -0.0523 0.0069 1.0000 9.500 1.3122 0.02573 0.01768 -0.0488 0.0073 1.0000 10.000 1.3337 0.02774 0.01991 -0.0445 0.0079 1.0000 10.500 1.3410 0.03100 0.02343 -0.0398 0.0085 1.0000 11.000 1.3368 0.03593 0.02858 -0.0365 0.0090 1.0000 11.500 1.3424 0.04079 0.03365 -0.0353 0.0095 1.0000 12.000 1.3426 0.04684 0.03997 -0.0350 0.0104 1.0000 12.500 1.3349 0.05393 0.04720 -0.0348 0.0111 1.0000 13.000 1.3413 0.05941 0.05291 -0.0343 0.0121 1.0000 13.500 1.3589 0.06235 0.05583 -0.0291 0.0139 1.0000 14.000 1.4051 0.06490 0.05894 -0.0209 0.0193 1.0000 14.500 1.1349 0.08595 0.08056 -0.0253 0.0120 1.0000 15.000 1.1621 0.08714 0.08187 -0.0204 0.0142 1.0000 15.500 1.1928 0.08936 0.08466 -0.0120 0.0194 1.0000 16.000 1.1354 0.10161 0.09739 -0.0179 0.0190 1.0000