XFOIL Version 6.94 Calculated polar for: NACA 4413 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.070 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1191 0.03299 0.02333 -0.0759 0.9397 0.1974 -2.750 -0.0702 0.03214 0.02242 -0.0801 0.9322 0.2177 -2.500 -0.0389 0.03161 0.02184 -0.0812 0.9201 0.2369 -2.250 0.0103 0.03090 0.02127 -0.0852 0.9134 0.2661 -2.000 0.0375 0.03055 0.02092 -0.0855 0.9006 0.2938 -1.750 0.0829 0.02993 0.02039 -0.0887 0.8930 0.3375 -1.500 0.1116 0.02942 0.02013 -0.0891 0.8810 0.3811 -1.250 0.1533 0.02869 0.01980 -0.0914 0.8734 0.4576 -1.000 0.1801 0.02792 0.01983 -0.0909 0.8624 0.5878 -0.750 0.2322 0.02665 0.01935 -0.0926 0.8571 1.0000 -0.250 0.2923 0.02728 0.01940 -0.0940 0.8341 1.0000 0.000 0.3237 0.02753 0.01943 -0.0948 0.8231 1.0000 0.500 0.3909 0.02793 0.01950 -0.0969 0.8039 1.0000 0.750 0.4403 0.02758 0.01898 -0.0999 0.7991 1.0000 1.000 0.4537 0.02838 0.01969 -0.0980 0.7851 1.0000 1.250 0.5027 0.02789 0.01907 -0.1007 0.7808 1.0000 1.500 0.5118 0.02896 0.02008 -0.0984 0.7665 1.0000 1.750 0.5590 0.02842 0.01944 -0.1006 0.7621 1.0000 2.000 0.5672 0.02964 0.02062 -0.0982 0.7481 1.0000 2.250 0.6126 0.02911 0.02002 -0.1001 0.7435 1.0000 2.750 0.6638 0.02993 0.02076 -0.0993 0.7251 1.0000 3.250 0.7122 0.03096 0.02175 -0.0982 0.7067 1.0000 3.750 0.7591 0.03211 0.02291 -0.0969 0.6883 1.0000 4.250 0.8047 0.03336 0.02418 -0.0954 0.6700 1.0000 4.500 0.8535 0.03237 0.02319 -0.0971 0.6658 1.0000 4.750 0.8495 0.03467 0.02556 -0.0939 0.6516 1.0000 5.000 0.8993 0.03356 0.02444 -0.0955 0.6472 1.0000