XFOIL Version 6.94 Calculated polar for: manu04 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4494 0.06485 0.05650 0.0429 0.9999 0.6147 -2.750 -0.4046 0.06344 0.05466 0.0287 0.9999 0.5386 -2.500 -0.3576 0.06179 0.05256 0.0172 0.9999 0.4796 -2.250 -0.3230 0.05933 0.04992 0.0127 0.9999 0.4525 -2.000 -0.2882 0.05740 0.04773 0.0083 0.9999 0.4338 -1.750 -0.2560 0.05544 0.04559 0.0055 0.9999 0.4208 -1.500 -0.2216 0.05365 0.04360 0.0028 0.9999 0.4094 -1.250 -0.1830 0.05202 0.04174 -0.0006 0.9999 0.3980 -1.000 -0.1422 0.05048 0.04006 -0.0037 0.9999 0.3919 -0.750 -0.0986 0.04885 0.03846 -0.0072 0.9999 0.3901 -0.250 0.2590 0.03739 0.02417 -0.0593 0.5680 0.4988 0.000 0.4306 0.03497 0.02172 -0.0887 0.5283 1.0001 0.250 0.4854 0.03611 0.02193 -0.0944 0.5172 1.0001 0.500 0.5338 0.03719 0.02232 -0.0985 0.5093 1.0001 0.750 0.5786 0.03828 0.02289 -0.1020 0.5027 1.0001 1.000 0.6168 0.03922 0.02359 -0.1042 0.4968 1.0001 1.250 0.6543 0.04030 0.02444 -0.1063 0.4909 1.0001 1.500 0.6919 0.04153 0.02543 -0.1084 0.4853 1.0001 1.750 0.7266 0.04278 0.02655 -0.1101 0.4815 1.0001 2.000 0.7561 0.04394 0.02775 -0.1108 0.4790 1.0001 2.250 0.7863 0.04522 0.02907 -0.1118 0.4785 1.0001 2.500 0.8153 0.04662 0.03058 -0.1127 0.4787 1.0001 2.750 0.8427 0.04815 0.03220 -0.1133 0.4794 1.0001 3.000 0.8685 0.04980 0.03396 -0.1138 0.4804 1.0001 3.250 0.8930 0.05159 0.03585 -0.1141 0.4820 1.0001 3.500 0.9107 0.05344 0.03792 -0.1135 0.4842 1.0001 3.750 0.9187 0.05553 0.04033 -0.1117 0.4878 1.0001 4.000 0.9244 0.05810 0.04314 -0.1099 0.4921 1.0001 4.250 0.9281 0.06101 0.04624 -0.1082 0.4969 1.0001 4.500 0.9354 0.06413 0.04947 -0.1071 0.5016 1.0001 4.750 0.9313 0.06772 0.05322 -0.1051 0.5074 1.0001 5.000 0.8754 0.07422 0.05993 -0.0992 0.5197 1.0001