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 = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.3567 0.07418 0.06596 0.0091 0.9999 0.4175 -2.750 -0.3008 0.06982 0.06116 -0.0040 0.9999 0.3745 -2.500 -0.2550 0.06652 0.05752 -0.0113 0.9999 0.3513 -2.250 -0.2079 0.06383 0.05438 -0.0180 0.9999 0.3320 -2.000 -0.1710 0.06144 0.05186 -0.0208 0.9999 0.3202 -1.750 -0.1279 0.05972 0.04970 -0.0247 0.9999 0.3080 -1.500 -0.0903 0.05814 0.04794 -0.0267 0.9999 0.3015 -1.250 -0.0519 0.05690 0.04658 -0.0285 0.9999 0.2987 -1.000 -0.0127 0.05584 0.04561 -0.0302 0.9999 0.3002 -0.750 0.0344 0.05479 0.04521 -0.0334 0.9999 0.3034 -0.500 0.2628 0.04613 0.03346 -0.0680 0.5345 0.3322 -0.250 0.3280 0.04590 0.03316 -0.0745 0.4989 0.3682 0.000 0.3920 0.04510 0.03333 -0.0810 0.4786 0.4689 0.250 0.4726 0.04488 0.03294 -0.0883 0.4608 1.0001 0.500 0.5482 0.04715 0.03397 -0.0969 0.4425 1.0001 0.750 0.6091 0.04912 0.03545 -0.1036 0.4283 1.0001 1.000 0.6886 0.05200 0.03788 -0.1148 0.4204 1.0001 1.250 0.7466 0.05440 0.04031 -0.1217 0.4178 1.0001 1.500 0.7982 0.05698 0.04297 -0.1275 0.4172 1.0001 1.750 0.8385 0.05951 0.04560 -0.1308 0.4174 1.0001 2.000 0.8654 0.06190 0.04813 -0.1315 0.4171 1.0001 2.250 0.8816 0.06425 0.05061 -0.1299 0.4168 1.0001 2.500 0.8898 0.06666 0.05314 -0.1268 0.4172 1.0001 2.750 0.8947 0.06924 0.05583 -0.1233 0.4183 1.0001 3.000 0.9041 0.07223 0.05888 -0.1210 0.4208 1.0001 3.250 0.8577 0.07390 0.06092 -0.1086 0.4278 1.0001 3.500 0.8173 0.07684 0.06409 -0.0987 0.4357 1.0001 3.750 0.8179 0.08070 0.06795 -0.0962 0.4431 1.0001 4.000 0.7256 0.08506 0.07271 -0.0823 0.4580 1.0001 4.250 0.6651 0.09108 0.07897 -0.0767 0.4779 1.0001 4.500 0.6175 0.09811 0.08615 -0.0758 0.5068 1.0001 4.750 0.5725 0.10606 0.09425 -0.0785 0.5554 1.0001 5.000 0.4741 0.11517 0.10378 -0.0820 0.6896 1.0001