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 = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.3922 0.07525 0.06813 0.0233 0.9999 0.4920 -2.750 -0.3021 0.06862 0.06073 -0.0038 0.9999 0.3761 -2.500 -0.2464 0.06486 0.05645 -0.0140 0.9999 0.3301 -2.250 -0.2019 0.06210 0.05332 -0.0195 0.9999 0.3075 -2.000 -0.1638 0.05972 0.05076 -0.0226 0.9999 0.2948 -1.750 -0.1236 0.05799 0.04872 -0.0256 0.9999 0.2838 -1.500 -0.0865 0.05632 0.04692 -0.0274 0.9999 0.2746 -1.250 -0.0449 0.05536 0.04562 -0.0296 0.9999 0.2656 -1.000 -0.0032 0.05401 0.04468 -0.0314 0.9999 0.2628 -0.750 0.2075 0.04566 0.03320 -0.0628 0.4931 0.2752 -0.500 0.2813 0.04548 0.03227 -0.0707 0.4579 0.2851 -0.250 0.3575 0.04535 0.03192 -0.0795 0.4355 0.3108 0.000 0.4294 0.04501 0.03211 -0.0880 0.4169 0.3782 0.250 0.5065 0.04414 0.03210 -0.0953 0.3989 1.0001 0.500 0.5741 0.04584 0.03308 -0.1024 0.3888 1.0001 0.750 0.6593 0.04835 0.03503 -0.1145 0.3817 1.0001 1.000 0.7632 0.05204 0.03832 -0.1313 0.3770 1.0001 1.250 0.8162 0.05436 0.04075 -0.1370 0.3754 1.0001 1.500 0.8488 0.05641 0.04295 -0.1383 0.3732 1.0001 1.750 0.8720 0.05849 0.04515 -0.1376 0.3714 1.0001 2.000 0.8927 0.06074 0.04751 -0.1366 0.3710 1.0001 2.250 0.9141 0.06334 0.05021 -0.1358 0.3733 1.0001 2.500 0.9399 0.06648 0.05335 -0.1361 0.3766 1.0001 2.750 0.9132 0.06759 0.05497 -0.1259 0.3841 1.0001 3.000 0.9018 0.07008 0.05768 -0.1194 0.3912 1.0001 3.250 0.9132 0.07354 0.06116 -0.1175 0.3978 1.0001 3.500 0.8539 0.07514 0.06318 -0.1031 0.4075 1.0001 3.750 0.8322 0.07871 0.06689 -0.0967 0.4180 1.0001 4.000 0.7592 0.08263 0.07117 -0.0844 0.4334 1.0001 4.250 0.6909 0.08871 0.07753 -0.0772 0.4575 1.0001 4.750 0.2549 0.11344 0.10466 -0.0621 0.9646 1.0001 5.000 0.2606 0.11602 0.10696 -0.0615 0.9697 1.0001