XFOIL Version 6.94 Calculated polar for: GOE 394 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 0.5673 0.01101 0.00379 -0.1165 0.7559 0.0750 0.500 0.6231 0.01166 0.00429 -0.1160 0.7439 0.1113 1.000 0.6781 0.01159 0.00422 -0.1155 0.7303 0.1233 1.500 0.7328 0.01142 0.00407 -0.1149 0.7156 0.1346 3.500 0.9422 0.01088 0.00316 -0.1105 0.5025 0.1774 4.000 0.9889 0.01058 0.00377 -0.1093 0.4101 1.0000 4.500 1.0394 0.01132 0.00422 -0.1084 0.3554 1.0000 5.000 1.0885 0.01221 0.00476 -0.1074 0.2872 1.0000 5.500 1.1160 0.01609 0.00690 -0.1044 0.0235 1.0000 6.500 1.2111 0.01797 0.00888 -0.1016 0.0061 1.0000 7.000 1.2569 0.01901 0.01012 -0.1000 0.0060 1.0000 7.500 1.3007 0.02019 0.01156 -0.0980 0.0062 1.0000 8.000 1.3410 0.02165 0.01330 -0.0955 0.0064 1.0000 8.500 1.3756 0.02353 0.01546 -0.0924 0.0068 1.0000 9.000 1.3992 0.02605 0.01829 -0.0879 0.0071 1.0000 9.500 1.4082 0.02928 0.02180 -0.0819 0.0075 1.0000 10.000 1.4127 0.03340 0.02629 -0.0766 0.0079 1.0000 10.500 1.4175 0.03821 0.03140 -0.0719 0.0083 1.0000 11.000 1.4239 0.04411 0.03766 -0.0670 0.0088 1.0000 11.500 1.4397 0.04809 0.04191 -0.0640 0.0095 1.0000 12.000 1.4496 0.05424 0.04865 -0.0600 0.0112 1.0000 12.500 1.4295 0.06438 0.05951 -0.0566 0.0129 1.0000 13.000 1.3973 0.07491 0.07055 -0.0559 0.0137 1.0000 13.500 1.3604 0.08650 0.08257 -0.0579 0.0142 1.0000 14.000 1.3217 0.09967 0.09613 -0.0628 0.0144 1.0000 14.500 1.2863 0.11442 0.11123 -0.0710 0.0144 1.0000 15.000 1.2521 0.13141 0.12854 -0.0820 0.0142 1.0000 15.500 1.2199 0.15097 0.14838 -0.0958 0.0137 1.0000 16.000 1.1836 0.17756 0.17515 -0.1132 0.0127 1.0000 16.500 1.1670 0.19907 0.19658 -0.1250 0.0139 1.0000