XFOIL Version 6.94 Calculated polar for: GOA 3 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1481 0.02714 0.01455 -0.0325 0.9999 0.1644 -2.750 -0.1215 0.02510 0.01291 -0.0315 0.9999 0.2156 -2.500 -0.1021 0.01996 0.01138 -0.0276 0.9999 1.0001 -2.250 -0.0755 0.02001 0.01059 -0.0263 0.9999 1.0001 -2.000 -0.0507 0.02007 0.01027 -0.0251 0.9999 1.0001 -1.750 -0.0260 0.02015 0.01005 -0.0239 0.9999 1.0001 -1.500 -0.0012 0.02024 0.00994 -0.0228 0.9999 1.0001 -1.250 0.0240 0.02033 0.00993 -0.0217 0.9999 1.0001 -1.000 0.0500 0.02041 0.01007 -0.0206 0.9999 1.0001 -0.500 0.1796 0.02382 0.01038 -0.0293 0.2950 1.0001 -0.250 0.2045 0.02488 0.01101 -0.0281 0.2793 1.0001 0.000 0.2312 0.02595 0.01172 -0.0272 0.2640 1.0001 0.250 0.2595 0.02708 0.01254 -0.0266 0.2481 1.0001 0.500 0.2898 0.02824 0.01348 -0.0264 0.2366 1.0001 0.750 0.3227 0.02941 0.01460 -0.0264 0.2329 1.0001 1.000 0.3566 0.03068 0.01593 -0.0266 0.2317 1.0001 1.250 0.3908 0.03202 0.01746 -0.0270 0.2324 1.0001 1.500 0.4242 0.03350 0.01917 -0.0273 0.2347 1.0001 1.750 0.4566 0.03518 0.02110 -0.0276 0.2382 1.0001 2.000 0.4876 0.03715 0.02323 -0.0278 0.2420 1.0001 2.250 0.5195 0.03887 0.02550 -0.0284 0.2501 1.0001 2.500 0.5491 0.04125 0.02812 -0.0288 0.2576 1.0001 2.750 0.5808 0.04351 0.03093 -0.0300 0.2706 1.0001 3.000 0.6117 0.04615 0.03406 -0.0315 0.2864 1.0001 3.250 0.6421 0.04920 0.03757 -0.0335 0.3062 1.0001 3.500 0.6747 0.05257 0.04158 -0.0377 0.3387 1.0001 3.750 0.7087 0.05676 0.04646 -0.0453 0.3898 1.0001 4.000 0.7327 0.06306 0.05353 -0.0633 0.4852 1.0001 4.250 0.7015 0.07102 0.06157 -0.0811 0.5854 1.0001 4.500 0.6496 0.07625 0.06657 -0.0867 0.6678 1.0001 4.750 0.5814 0.07970 0.06982 -0.0871 0.7714 1.0001