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 = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.2547 0.04732 0.04094 -0.0362 0.5129 0.9999 -2.750 0.2698 0.04596 0.03910 -0.0371 0.4684 0.9999 -2.500 0.2576 0.04520 0.03825 -0.0322 0.4548 0.9890 -2.250 0.1815 0.04573 0.03914 -0.0149 0.4677 0.9436 -2.000 0.1212 0.04507 0.03874 -0.0028 0.4750 0.8948 -1.750 0.0542 0.04393 0.03797 0.0088 0.4880 0.8503 -1.500 0.3517 0.04274 0.03301 -0.0968 0.4202 0.2789 -1.250 0.3990 0.04176 0.03120 -0.1002 0.4072 0.2346 -1.000 0.4397 0.04088 0.02970 -0.1021 0.3963 0.2112 -0.750 0.4794 0.04044 0.02863 -0.1035 0.3882 0.1952 -0.500 0.5157 0.03955 0.02748 -0.1046 0.3819 0.1876 -0.250 0.5528 0.03916 0.02666 -0.1056 0.3765 0.1836 0.000 0.5879 0.03897 0.02614 -0.1063 0.3722 0.1839 0.250 0.6213 0.03881 0.02581 -0.1065 0.3683 0.1847 0.500 0.6536 0.03881 0.02560 -0.1065 0.3641 0.1855 0.750 0.6847 0.03894 0.02553 -0.1063 0.3597 0.1884 1.000 0.7155 0.03936 0.02570 -0.1061 0.3554 0.1943 1.250 0.7457 0.03968 0.02609 -0.1062 0.3519 0.2068 1.500 0.7767 0.04008 0.02667 -0.1066 0.3491 0.2272 1.750 0.8094 0.04047 0.02745 -0.1075 0.3472 0.2766 2.000 0.8310 0.03939 0.02769 -0.1048 0.3459 0.9999 2.250 0.8611 0.04086 0.02875 -0.1049 0.3446 0.9999 2.500 0.8899 0.04247 0.03022 -0.1053 0.3437 0.9999 2.750 0.9176 0.04419 0.03191 -0.1057 0.3428 0.9999 3.000 0.9443 0.04602 0.03373 -0.1061 0.3417 0.9999 3.250 0.9701 0.04795 0.03566 -0.1064 0.3402 0.9999 3.500 0.9951 0.04998 0.03770 -0.1066 0.3387 0.9999 3.750 1.0190 0.05221 0.03997 -0.1068 0.3376 0.9999 4.000 1.0393 0.05497 0.04297 -0.1071 0.3387 0.9999 4.250 1.0547 0.05858 0.04694 -0.1074 0.3419 0.9999 4.500 1.0654 0.06288 0.05158 -0.1076 0.3463 0.9999 4.750 1.0779 0.06706 0.05591 -0.1079 0.3504 0.9999 5.000 1.0861 0.07180 0.06087 -0.1083 0.3563 0.9999