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 = 0.060 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2187 0.03584 0.02666 -0.0582 0.9924 0.1742 -2.750 -0.1682 0.03466 0.02498 -0.0632 0.9831 0.1893 -2.500 -0.1226 0.03381 0.02391 -0.0672 0.9731 0.2075 -2.250 -0.0739 0.03298 0.02305 -0.0717 0.9644 0.2298 -2.000 -0.0327 0.03241 0.02248 -0.0748 0.9532 0.2559 -1.750 0.0078 0.03194 0.02197 -0.0777 0.9425 0.2893 -1.500 0.0551 0.03143 0.02159 -0.0815 0.9334 0.3377 -1.250 0.0884 0.03095 0.02130 -0.0829 0.9215 0.3925 -1.000 0.1389 0.03005 0.02108 -0.0869 0.9142 0.5126 -0.750 0.1535 0.02846 0.02106 -0.0825 0.9011 1.0000 -0.500 0.1927 0.02902 0.02099 -0.0853 0.8893 1.0000 -0.250 0.2379 0.02945 0.02102 -0.0888 0.8792 1.0000 0.250 0.3020 0.03049 0.02160 -0.0913 0.8558 1.0000 0.500 0.3343 0.03097 0.02190 -0.0925 0.8444 1.0000 0.750 0.3596 0.03162 0.02241 -0.0926 0.8324 1.0000 1.000 0.4032 0.03185 0.02249 -0.0953 0.8241 1.0000 1.500 0.4690 0.03271 0.02314 -0.0974 0.8045 1.0000 2.000 0.5169 0.03413 0.02441 -0.0970 0.7830 1.0000 2.250 0.5370 0.03497 0.02521 -0.0964 0.7715 1.0000 2.750 0.5912 0.03625 0.02642 -0.0967 0.7524 1.0000 3.250 0.6427 0.03763 0.02777 -0.0966 0.7334 1.0000 3.750 0.6923 0.03910 0.02926 -0.0961 0.7145 1.0000 4.250 0.7405 0.04062 0.03082 -0.0954 0.6955 1.0000 4.750 0.7879 0.04218 0.03246 -0.0944 0.6765 1.0000