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 = 0.050 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2875 0.04056 0.03145 -0.0480 1.0000 0.2077 -2.750 -0.2615 0.03914 0.02947 -0.0493 1.0000 0.2173 -2.500 -0.2146 0.03827 0.02841 -0.0540 0.9925 0.2340 -2.250 -0.1633 0.03757 0.02738 -0.0593 0.9818 0.2564 -2.000 -0.1197 0.03697 0.02679 -0.0631 0.9706 0.2811 -1.750 -0.0762 0.03662 0.02638 -0.0668 0.9596 0.3123 -1.500 -0.0277 0.03631 0.02605 -0.0711 0.9497 0.3549 -1.250 0.0062 0.03596 0.02580 -0.0728 0.9376 0.4001 -1.000 0.0475 0.03564 0.02582 -0.0755 0.9273 0.4700 -0.750 0.0827 0.03500 0.02594 -0.0767 0.9170 0.5912 -0.500 0.1041 0.03389 0.02580 -0.0742 0.9052 1.0000 -0.250 0.1554 0.03479 0.02602 -0.0793 0.8950 1.0000 0.000 0.1762 0.03555 0.02649 -0.0793 0.8819 1.0000 0.250 0.2107 0.03641 0.02706 -0.0813 0.8713 1.0000 0.500 0.2427 0.03722 0.02763 -0.0829 0.8602 1.0000 0.750 0.2659 0.03814 0.02838 -0.0832 0.8491 1.0000 1.000 0.3049 0.03893 0.02897 -0.0856 0.8396 1.0000 1.250 0.3199 0.03999 0.02990 -0.0848 0.8282 1.0000 1.500 0.3630 0.04072 0.03047 -0.0876 0.8196 1.0000 1.750 0.3710 0.04196 0.03164 -0.0859 0.8082 1.0000 2.000 0.4160 0.04263 0.03218 -0.0888 0.8001 1.0000 2.250 0.4189 0.04409 0.03358 -0.0865 0.7887 1.0000 2.500 0.4640 0.04472 0.03413 -0.0893 0.7809 1.0000 2.750 0.4636 0.04639 0.03576 -0.0868 0.7697 1.0000 3.000 0.5082 0.04700 0.03631 -0.0893 0.7618 1.0000 3.250 0.5051 0.04890 0.03819 -0.0867 0.7509 1.0000 3.500 0.5490 0.04949 0.03874 -0.0889 0.7429 1.0000 3.750 0.5441 0.05162 0.04086 -0.0864 0.7323 1.0000 4.000 0.5860 0.05225 0.04148 -0.0883 0.7240 1.0000 4.250 0.5810 0.05453 0.04376 -0.0859 0.7135 1.0000 4.500 0.6213 0.05520 0.04444 -0.0875 0.7051 1.0000 4.750 0.6153 0.05768 0.04694 -0.0852 0.6948 1.0000 5.000 0.6561 0.05830 0.04759 -0.0866 0.6858 1.0000