XFOIL Version 6.94 Calculated polar for: NACA 4412 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2533 0.03933 0.03025 -0.0518 1.0000 0.1944 -2.750 -0.2289 0.03805 0.02861 -0.0526 1.0000 0.2040 -2.500 -0.2049 0.03678 0.02711 -0.0533 1.0000 0.2129 -2.250 -0.1805 0.03609 0.02600 -0.0539 1.0000 0.2268 -2.000 -0.1443 0.03539 0.02518 -0.0566 0.9956 0.2455 -1.750 -0.0941 0.03481 0.02460 -0.0616 0.9851 0.2738 -1.500 -0.0464 0.03442 0.02412 -0.0661 0.9742 0.3109 -1.250 0.0032 0.03414 0.02381 -0.0705 0.9640 0.3636 -1.000 0.0452 0.03366 0.02365 -0.0736 0.9529 0.4309 -0.750 0.0833 0.03298 0.02369 -0.0757 0.9423 0.5509 -0.500 0.1196 0.03179 0.02346 -0.0757 0.9314 1.0000 -0.250 0.1528 0.03255 0.02372 -0.0779 0.9180 1.0000 0.000 0.1852 0.03338 0.02420 -0.0797 0.9053 1.0000 0.250 0.2294 0.03425 0.02474 -0.0834 0.8946 1.0000 0.500 0.2558 0.03504 0.02533 -0.0842 0.8819 1.0000 0.750 0.2835 0.03592 0.02601 -0.0851 0.8701 1.0000 1.000 0.3270 0.03668 0.02658 -0.0884 0.8601 1.0000 1.250 0.3432 0.03765 0.02743 -0.0876 0.8477 1.0000 1.500 0.3774 0.03850 0.02815 -0.0893 0.8377 1.0000 1.750 0.4024 0.03942 0.02897 -0.0898 0.8265 1.0000 2.000 0.4247 0.04047 0.02994 -0.0899 0.8160 1.0000 2.250 0.4595 0.04126 0.03065 -0.0915 0.8063 1.0000 2.500 0.4729 0.04254 0.03189 -0.0905 0.7952 1.0000 2.750 0.5137 0.04321 0.03249 -0.0927 0.7866 1.0000 3.000 0.5198 0.04474 0.03400 -0.0909 0.7750 1.0000 3.250 0.5660 0.04524 0.03448 -0.0936 0.7671 1.0000 3.500 0.5645 0.04711 0.03633 -0.0911 0.7552 1.0000 3.750 0.5989 0.04795 0.03716 -0.0923 0.7463 1.0000 4.000 0.6075 0.04963 0.03885 -0.0910 0.7354 1.0000 4.250 0.6315 0.05085 0.04009 -0.0911 0.7258 1.0000 4.500 0.6499 0.05225 0.04152 -0.0908 0.7155 1.0000 4.750 0.6639 0.05392 0.04321 -0.0901 0.7054 1.0000 5.000 0.6930 0.05490 0.04423 -0.0905 0.6954 1.0000