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 = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.1518 0.05522 0.05062 -0.0189 1.0001 0.9999 -2.750 0.0915 0.05643 0.05205 -0.0044 1.0001 0.9777 -2.500 0.0389 0.05638 0.05219 0.0069 0.9873 0.9419 -2.250 -0.0085 0.05522 0.05116 0.0148 0.9620 0.8936 -2.000 -0.0572 0.05377 0.04983 0.0220 0.9369 0.8518 -1.750 0.3224 0.04464 0.03592 -0.0969 0.5551 0.2864 -1.500 0.3607 0.04379 0.03402 -0.0988 0.5128 0.2617 -1.250 0.4026 0.04332 0.03268 -0.1011 0.4879 0.2413 -1.000 0.4381 0.04236 0.03126 -0.1024 0.4706 0.2332 -0.750 0.4756 0.04181 0.03022 -0.1038 0.4557 0.2273 -0.500 0.5124 0.04154 0.02943 -0.1049 0.4435 0.2233 -0.250 0.5480 0.04155 0.02886 -0.1054 0.4330 0.2204 0.000 0.5823 0.04127 0.02827 -0.1058 0.4251 0.2194 0.250 0.6162 0.04120 0.02793 -0.1061 0.4189 0.2201 0.500 0.6488 0.04104 0.02761 -0.1063 0.4139 0.2254 0.750 0.6812 0.04124 0.02774 -0.1065 0.4098 0.2360 1.000 0.7130 0.04151 0.02813 -0.1069 0.4058 0.2489 1.250 0.7448 0.04191 0.02862 -0.1073 0.4014 0.2685 1.500 0.7775 0.04207 0.02918 -0.1082 0.3968 0.3237 1.750 0.7998 0.04107 0.02927 -0.1055 0.3932 0.9999 2.000 0.8302 0.04264 0.03029 -0.1053 0.3898 0.9999 2.250 0.8586 0.04439 0.03184 -0.1058 0.3873 0.9999 2.500 0.8861 0.04635 0.03373 -0.1064 0.3861 0.9999 2.750 0.9124 0.04852 0.03589 -0.1071 0.3854 0.9999 3.000 0.9372 0.05095 0.03836 -0.1077 0.3853 0.9999 3.250 0.9600 0.05366 0.04115 -0.1084 0.3858 0.9999 3.500 0.9799 0.05671 0.04432 -0.1090 0.3863 0.9999 3.750 0.9963 0.06014 0.04790 -0.1094 0.3868 0.9999 4.000 1.0086 0.06402 0.05195 -0.1098 0.3876 0.9999 4.250 1.0171 0.06837 0.05644 -0.1101 0.3890 0.9999 4.500 1.0258 0.07283 0.06099 -0.1104 0.3911 0.9999