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 = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2901 0.07983 0.07324 0.0489 0.9999 0.7258 -2.750 -0.3097 0.07828 0.07181 0.0468 0.9999 0.6891 -2.250 -0.2484 0.07303 0.06640 0.0169 0.9999 0.5709 -2.000 -0.1904 0.07023 0.06347 0.0001 0.9999 0.5183 -1.750 -0.1272 0.06801 0.06106 -0.0164 0.9999 0.4736 -1.500 -0.0641 0.06667 0.05944 -0.0319 0.9999 0.4361 -1.250 -0.0245 0.06643 0.05919 -0.0409 0.9999 0.4138 -1.000 0.2818 0.06558 0.05702 -0.0992 0.7870 0.3678 -0.750 0.3876 0.06286 0.05348 -0.1099 0.7308 0.3637 -0.500 0.4326 0.06278 0.05292 -0.1126 0.6959 0.3634 -0.250 0.4820 0.06264 0.05223 -0.1155 0.6711 0.3692 0.000 0.4977 0.06451 0.05399 -0.1156 0.6546 0.3750 0.250 0.5231 0.06597 0.05523 -0.1165 0.6413 0.3849 0.500 0.5461 0.06757 0.05669 -0.1171 0.6292 0.3965 0.750 0.5525 0.07038 0.05947 -0.1169 0.6201 0.4068 1.000 0.5563 0.07335 0.06243 -0.1164 0.6120 0.4180 1.250 0.5853 0.07425 0.06355 -0.1169 0.6026 0.4554 1.500 0.5740 0.07859 0.06801 -0.1164 0.5994 0.4648 1.750 0.5724 0.08213 0.07186 -0.1164 0.5976 0.4931 2.250 0.5640 0.08808 0.07881 -0.1144 0.6012 1.0001 2.500 0.5769 0.09219 0.08234 -0.1156 0.6038 1.0001 2.750 0.5507 0.09715 0.08741 -0.1155 0.6150 1.0001 3.000 0.5591 0.10127 0.09103 -0.1164 0.6214 1.0001 3.500 0.5594 0.10966 0.09876 -0.1178 0.6441 1.0001 3.750 0.5474 0.11402 0.10306 -0.1187 0.6705 1.0001 4.500 0.3975 0.12120 0.11166 -0.1108 0.9659 1.0001 4.750 0.4082 0.12300 0.11296 -0.1104 0.9610 1.0001 5.000 0.4276 0.12617 0.11549 -0.1114 0.9553 1.0001