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 = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0024 0.06657 0.06188 0.0103 0.9999 0.9661 -2.750 -0.0396 0.06658 0.06206 0.0189 0.9935 0.9342 -2.500 -0.0795 0.06615 0.06175 0.0252 0.9763 0.8907 -2.250 -0.1199 0.06534 0.06107 0.0307 0.9579 0.8503 -2.000 -0.1563 0.06414 0.05997 0.0331 0.9371 0.8044 -1.750 0.4022 0.04262 0.03504 -0.1052 0.6992 0.2774 -1.500 0.4594 0.04071 0.03195 -0.1077 0.6101 0.2597 -1.250 0.4962 0.03930 0.02975 -0.1077 0.5585 0.2534 -1.000 0.5312 0.03866 0.02823 -0.1078 0.5254 0.2477 -0.750 0.5649 0.03834 0.02708 -0.1078 0.5023 0.2459 -0.500 0.5973 0.03821 0.02640 -0.1078 0.4849 0.2486 -0.250 0.6301 0.03823 0.02599 -0.1079 0.4727 0.2532 0.000 0.6640 0.03831 0.02556 -0.1079 0.4633 0.2582 0.250 0.6948 0.03841 0.02561 -0.1079 0.4555 0.2655 0.500 0.7270 0.03867 0.02573 -0.1081 0.4482 0.2813 0.750 0.7602 0.03870 0.02581 -0.1085 0.4417 0.3112 1.000 0.7892 0.03871 0.02661 -0.1090 0.4353 0.3815 1.250 0.8100 0.03818 0.02696 -0.1059 0.4299 1.0001 1.500 0.8442 0.03950 0.02727 -0.1056 0.4241 1.0001 1.750 0.8706 0.04118 0.02862 -0.1055 0.4203 1.0001 2.000 0.8937 0.04313 0.03050 -0.1055 0.4179 1.0001 2.250 0.9146 0.04532 0.03266 -0.1054 0.4161 1.0001 2.500 0.9326 0.04780 0.03518 -0.1051 0.4150 1.0001 2.750 0.9468 0.05068 0.03814 -0.1046 0.4145 1.0001 3.000 0.9553 0.05405 0.04163 -0.1040 0.4143 1.0001 3.250 0.9560 0.05812 0.04587 -0.1029 0.4147 1.0001 3.500 0.9427 0.06348 0.05145 -0.1016 0.4162 1.0001 3.750 0.9189 0.07003 0.05821 -0.1005 0.4190 1.0001 4.000 0.9034 0.07618 0.06444 -0.1001 0.4217 1.0001