XFOIL Version 6.94 Calculated polar for: GOE 406c 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.2658 0.04543 0.04097 -0.0374 0.8022 0.9134 -2.750 0.2151 0.04621 0.04186 -0.0273 0.7896 0.8695 -2.500 0.1669 0.04626 0.04199 -0.0187 0.7776 0.8268 -2.250 0.1226 0.04608 0.04193 -0.0132 0.7607 0.7812 -2.000 0.4020 0.03951 0.03130 -0.1039 0.5307 0.2590 -1.750 0.4356 0.03860 0.02919 -0.1043 0.4726 0.2353 -1.500 0.4642 0.03763 0.02758 -0.1044 0.4464 0.2253 -1.250 0.5005 0.03763 0.02670 -0.1053 0.4303 0.2136 -1.000 0.5325 0.03685 0.02551 -0.1058 0.4195 0.2098 -0.750 0.5676 0.03630 0.02458 -0.1066 0.4104 0.2046 -0.500 0.6049 0.03608 0.02376 -0.1074 0.4030 0.2002 -0.250 0.6406 0.03590 0.02327 -0.1079 0.3962 0.1985 0.000 0.6760 0.03583 0.02289 -0.1084 0.3893 0.2010 0.250 0.7129 0.03602 0.02265 -0.1091 0.3829 0.2088 0.500 0.7436 0.03608 0.02272 -0.1089 0.3777 0.2171 0.750 0.7746 0.03623 0.02290 -0.1089 0.3731 0.2293 1.000 0.8067 0.03649 0.02323 -0.1093 0.3698 0.2519 1.250 0.8405 0.03628 0.02376 -0.1104 0.3670 0.3570 1.500 0.8621 0.03593 0.02395 -0.1070 0.3650 1.0001 1.750 0.8932 0.03744 0.02490 -0.1071 0.3630 1.0001 2.000 0.9196 0.03882 0.02616 -0.1070 0.3615 1.0001 2.250 0.9442 0.04031 0.02765 -0.1067 0.3598 1.0001 2.500 0.9673 0.04194 0.02930 -0.1063 0.3578 1.0001 2.750 0.9891 0.04370 0.03112 -0.1059 0.3556 1.0001 3.000 1.0097 0.04562 0.03309 -0.1053 0.3537 1.0001 3.250 1.0286 0.04774 0.03531 -0.1047 0.3528 1.0001 3.500 1.0445 0.05019 0.03790 -0.1039 0.3533 1.0001 3.750 1.0572 0.05297 0.04086 -0.1030 0.3543 1.0001 4.000 1.0665 0.05613 0.04421 -0.1019 0.3559 1.0001 4.250 1.0707 0.05980 0.04807 -0.1007 0.3581 1.0001 4.500 1.0752 0.06370 0.05210 -0.0997 0.3606 1.0001 4.750 1.0855 0.06749 0.05593 -0.0992 0.3626 1.0001 5.000 0.9607 0.08344 0.07282 -0.0958 0.3829 1.0001