XFOIL Version 6.94 Calculated polar for: COUDE2 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 -0.3515 0.07587 0.06654 0.0069 0.9999 0.4237 -2.750 -0.3076 0.07205 0.06244 -0.0014 0.9999 0.4000 -2.500 -0.2601 0.06878 0.05876 -0.0097 0.9999 0.3784 -2.250 -0.2222 0.06602 0.05585 -0.0134 0.9999 0.3661 -2.000 -0.1759 0.06389 0.05323 -0.0194 0.9999 0.3543 -1.750 -0.1378 0.06201 0.05116 -0.0221 0.9999 0.3512 -1.500 -0.0990 0.06048 0.04942 -0.0247 0.9999 0.3505 -1.250 -0.0601 0.05922 0.04800 -0.0270 0.9999 0.3507 -1.000 -0.0209 0.05814 0.04689 -0.0289 0.9999 0.3515 -0.750 0.0204 0.05723 0.04615 -0.0309 0.9999 0.3542 -0.500 0.0652 0.05635 0.04623 -0.0343 0.9999 0.3629 -0.250 0.3037 0.04535 0.03448 -0.0720 0.6187 0.4646 0.000 0.3671 0.04340 0.03336 -0.0759 0.5715 1.0001 0.250 0.4448 0.04571 0.03362 -0.0835 0.5405 1.0001 0.500 0.5136 0.04782 0.03461 -0.0909 0.5219 1.0001 0.750 0.5811 0.05018 0.03623 -0.0990 0.5069 1.0001 1.000 0.6497 0.05296 0.03839 -0.1076 0.4901 1.0001 1.250 0.6879 0.05491 0.04032 -0.1104 0.4803 1.0001 1.500 0.7356 0.05744 0.04273 -0.1152 0.4731 1.0001 1.750 0.7777 0.06010 0.04539 -0.1191 0.4711 1.0001 2.000 0.8044 0.06258 0.04799 -0.1201 0.4715 1.0001 2.250 0.8139 0.06491 0.05051 -0.1179 0.4737 1.0001 2.500 0.8104 0.06731 0.05313 -0.1134 0.4773 1.0001 2.750 0.8034 0.07009 0.05607 -0.1088 0.4815 1.0001 3.000 0.7987 0.07323 0.05931 -0.1050 0.4857 1.0001 3.250 0.7994 0.07673 0.06284 -0.1024 0.4893 1.0001 3.500 0.7133 0.08081 0.06737 -0.0885 0.5024 1.0001 3.750 0.7017 0.08529 0.07187 -0.0861 0.5102 1.0001 4.000 0.6316 0.09154 0.07839 -0.0799 0.5294 1.0001 4.500 0.5578 0.10380 0.09080 -0.0788 0.5760 1.0001 4.750 0.5162 0.10941 0.09650 -0.0789 0.6135 1.0001 5.000 0.4845 0.11450 0.10164 -0.0797 0.6614 1.0001