XFOIL Version 6.94 Calculated polar for: GOE 431 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 0.8733 0.01253 0.00563 -0.2041 0.7742 0.2133 0.500 0.9251 0.01219 0.00533 -0.2028 0.7595 0.2184 1.000 0.9782 0.01179 0.00492 -0.2017 0.7407 0.2250 1.500 1.0239 0.01142 0.00460 -0.1991 0.7090 0.2344 2.000 1.0448 0.01174 0.00422 -0.1911 0.5608 0.2467 2.500 1.0481 0.01327 0.00516 -0.1808 0.4707 0.2646 3.000 1.0780 0.01398 0.00586 -0.1760 0.4404 0.3263 3.500 1.1166 0.01447 0.00665 -0.1730 0.4229 0.4628 4.500 1.1683 0.01613 0.00821 -0.1623 0.3470 0.6062 5.000 1.1996 0.01637 0.00886 -0.1582 0.3209 1.0000 5.500 1.2380 0.01713 0.00957 -0.1554 0.3053 1.0000 6.000 1.2692 0.01822 0.01045 -0.1516 0.2723 1.0000 6.500 1.2875 0.02031 0.01191 -0.1464 0.1714 1.0000 7.000 1.2866 0.02438 0.01510 -0.1392 0.0247 1.0000 7.500 1.3160 0.02610 0.01681 -0.1358 0.0056 1.0000 8.000 1.3483 0.02761 0.01844 -0.1329 0.0056 1.0000 8.500 1.3779 0.02936 0.02034 -0.1297 0.0059 1.0000 9.000 1.4049 0.03138 0.02254 -0.1265 0.0062 1.0000 9.500 1.4309 0.03354 0.02489 -0.1233 0.0069 1.0000 10.000 1.4512 0.03628 0.02786 -0.1197 0.0076 1.0000 10.500 1.4654 0.03969 0.03153 -0.1160 0.0081 1.0000 11.000 1.4778 0.04346 0.03555 -0.1125 0.0091 1.0000 11.500 1.4737 0.04912 0.04147 -0.1087 0.0098 1.0000 12.000 1.4726 0.05496 0.04758 -0.1058 0.0107 1.0000 12.500 1.4596 0.06278 0.05567 -0.1035 0.0117 1.0000 13.000 1.4449 0.07149 0.06452 -0.1021 0.0124 1.0000 13.500 1.4596 0.07671 0.06991 -0.0999 0.0146 1.0000