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 = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2903 0.04220 0.03297 -0.0470 1.0000 0.2234 -2.750 -0.2656 0.04088 0.03120 -0.0482 1.0000 0.2344 -2.500 -0.2444 0.03979 0.03005 -0.0485 1.0000 0.2441 -2.250 -0.2204 0.03892 0.02885 -0.0493 1.0000 0.2584 -2.000 -0.1992 0.03837 0.02819 -0.0495 1.0000 0.2731 -1.750 -0.1514 0.03794 0.02759 -0.0543 0.9919 0.3006 -1.500 -0.0997 0.03775 0.02724 -0.0595 0.9808 0.3384 -1.250 -0.0569 0.03748 0.02703 -0.0631 0.9692 0.3815 -1.000 -0.0148 0.03731 0.02700 -0.0663 0.9578 0.4399 -0.750 0.0317 0.03703 0.02725 -0.0698 0.9477 0.5369 -0.500 0.0514 0.03593 0.02744 -0.0671 0.9376 0.7273 -0.250 0.0882 0.03605 0.02722 -0.0698 0.9241 1.0000 0.250 0.1576 0.03811 0.02846 -0.0748 0.9008 1.0000 0.500 0.1910 0.03919 0.02926 -0.0769 0.8899 1.0000 0.750 0.2237 0.04021 0.03004 -0.0788 0.8788 1.0000 1.000 0.2457 0.04127 0.03093 -0.0792 0.8681 1.0000 1.250 0.2851 0.04237 0.03182 -0.0820 0.8581 1.0000 1.500 0.2985 0.04348 0.03282 -0.0811 0.8473 1.0000 1.750 0.3419 0.04459 0.03377 -0.0844 0.8381 1.0000 2.000 0.3478 0.04583 0.03493 -0.0825 0.8276 1.0000 2.250 0.3923 0.04693 0.03590 -0.0857 0.8184 1.0000 2.500 0.3937 0.04833 0.03726 -0.0834 0.8085 1.0000 2.750 0.4388 0.04943 0.03826 -0.0865 0.7995 1.0000 3.000 0.4368 0.05101 0.03981 -0.0840 0.7897 1.0000 3.250 0.4804 0.05212 0.04087 -0.0868 0.7807 1.0000 3.500 0.4768 0.05389 0.04262 -0.0842 0.7714 1.0000 3.750 0.5178 0.05503 0.04373 -0.0866 0.7620 1.0000 4.000 0.5145 0.05699 0.04568 -0.0843 0.7533 1.0000 4.250 0.5506 0.05822 0.04690 -0.0860 0.7436 1.0000 4.500 0.5505 0.06029 0.04898 -0.0842 0.7351 1.0000 4.750 0.5790 0.06173 0.05042 -0.0852 0.7256 1.0000 5.000 0.5845 0.06380 0.05251 -0.0840 0.7167 1.0000