XFOIL Version 6.94 Calculated polar for: NACA4412c 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 -0.0893 0.04588 0.03379 -0.0443 0.9999 0.4020 -2.750 -0.0607 0.04440 0.03211 -0.0448 0.9999 0.4232 -2.500 -0.0332 0.04309 0.03072 -0.0449 0.9999 0.4509 -2.250 -0.0056 0.04190 0.02957 -0.0448 0.9999 0.4882 -2.000 0.0209 0.04071 0.02864 -0.0442 0.9999 0.5407 -1.750 0.0442 0.03925 0.02792 -0.0423 0.9999 0.6283 -1.500 0.0562 0.03674 0.02669 -0.0383 0.9999 1.0001 -1.250 0.1019 0.03768 0.02580 -0.0431 0.9999 1.0001 -1.000 0.1259 0.03849 0.02605 -0.0431 0.9999 1.0001 -0.750 0.1479 0.03939 0.02661 -0.0430 0.9999 1.0001 -0.500 0.1687 0.04042 0.02740 -0.0429 0.9999 1.0001 -0.250 0.1877 0.04167 0.02853 -0.0430 0.9999 1.0001 0.000 0.2030 0.04339 0.03026 -0.0433 0.9999 1.0001 0.250 0.1998 0.04743 0.03462 -0.0448 0.9999 1.0001 0.500 0.4787 0.04329 0.02974 -0.0894 0.8022 1.0001 0.750 0.5405 0.04284 0.02891 -0.0933 0.7541 1.0001 1.000 0.5781 0.04369 0.02947 -0.0944 0.7211 1.0001 1.250 0.6123 0.04466 0.03017 -0.0950 0.6956 1.0001 1.500 0.6426 0.04585 0.03113 -0.0952 0.6741 1.0001 1.750 0.6687 0.04744 0.03255 -0.0953 0.6554 1.0001 2.000 0.6917 0.04947 0.03446 -0.0956 0.6404 1.0001 2.250 0.7221 0.05099 0.03580 -0.0962 0.6296 1.0001 2.500 0.7308 0.05483 0.03975 -0.0971 0.6209 1.0001 2.750 0.7531 0.05742 0.04228 -0.0980 0.6132 1.0001 3.000 0.7658 0.06092 0.04578 -0.0986 0.6068 1.0001 3.250 0.7617 0.06585 0.05077 -0.0986 0.6017 1.0001 3.500 0.7631 0.07017 0.05510 -0.0986 0.5969 1.0001 3.750 0.7706 0.07407 0.05900 -0.0988 0.5931 1.0001 4.000 0.7322 0.08163 0.06659 -0.0976 0.5980 1.0001 4.250 0.7124 0.08765 0.07260 -0.0970 0.6039 1.0001 4.500 0.7154 0.09233 0.07727 -0.0979 0.6084 1.0001 4.750 0.6788 0.09898 0.08391 -0.0970 0.6234 1.0001 5.000 0.6803 0.10364 0.08857 -0.0982 0.6323 1.0001