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 = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2094 0.07594 0.07015 0.0448 0.9999 0.8261 -2.750 -0.2538 0.07533 0.06971 0.0520 0.9999 0.8037 -2.500 -0.2839 0.07401 0.06854 0.0547 0.9999 0.7707 -2.250 -0.3103 0.07267 0.06730 0.0536 0.9999 0.7316 -2.000 -0.2700 0.07000 0.06457 0.0314 0.9999 0.6377 -1.750 -0.1489 0.06697 0.06116 -0.0101 0.9999 0.5006 -1.500 0.2359 0.05868 0.05142 -0.0899 0.8328 0.3420 -1.250 0.4265 0.05015 0.04143 -0.1097 0.7133 0.3156 -1.000 0.4852 0.04833 0.03884 -0.1124 0.6627 0.3147 -0.750 0.5307 0.04745 0.03733 -0.1137 0.6298 0.3153 -0.500 0.5731 0.04678 0.03603 -0.1144 0.6058 0.3178 -0.250 0.6072 0.04677 0.03552 -0.1145 0.5855 0.3252 0.000 0.6331 0.04701 0.03566 -0.1139 0.5681 0.3373 0.250 0.6587 0.04751 0.03605 -0.1133 0.5536 0.3529 0.500 0.6927 0.04710 0.03558 -0.1130 0.5419 0.3836 0.750 0.7071 0.04883 0.03779 -0.1126 0.5330 0.4191 1.000 0.7170 0.04833 0.03900 -0.1093 0.5272 0.9214 1.250 0.7491 0.05044 0.04015 -0.1096 0.5213 1.0001 1.500 0.7503 0.05486 0.04438 -0.1091 0.5177 1.0001 1.750 0.7461 0.05972 0.04909 -0.1087 0.5151 1.0001 2.000 0.7332 0.06532 0.05467 -0.1084 0.5139 1.0001 2.250 0.7134 0.07149 0.06089 -0.1083 0.5141 1.0001 2.500 0.6916 0.07775 0.06722 -0.1084 0.5158 1.0001 2.750 0.6807 0.08338 0.07271 -0.1090 0.5174 1.0001 3.000 0.6779 0.08833 0.07739 -0.1096 0.5194 1.0001 3.250 0.6851 0.09272 0.08138 -0.1104 0.5220 1.0001 3.500 0.6453 0.10025 0.08929 -0.1122 0.5385 1.0001 4.000 0.6155 0.11186 0.10091 -0.1159 0.5804 1.0001