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 = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.3027 0.04688 0.03739 -0.0426 1.0000 0.2717 -2.750 -0.2766 0.04475 0.03488 -0.0445 1.0000 0.2801 -2.500 -0.2530 0.04341 0.03328 -0.0455 1.0000 0.2938 -2.250 -0.2307 0.04216 0.03189 -0.0459 1.0000 0.3070 -2.000 -0.2072 0.04117 0.03068 -0.0467 1.0000 0.3250 -1.750 -0.1839 0.04037 0.02969 -0.0473 1.0000 0.3451 -1.500 -0.1613 0.03979 0.02899 -0.0476 1.0000 0.3700 -1.250 -0.1370 0.03937 0.02840 -0.0483 1.0000 0.4008 -1.000 -0.1133 0.03909 0.02808 -0.0486 1.0000 0.4386 -0.750 -0.0907 0.03883 0.02804 -0.0486 1.0000 0.4858 -0.500 -0.0676 0.03856 0.02813 -0.0484 1.0000 0.5558 -0.250 -0.0484 0.03778 0.02831 -0.0462 1.0000 0.6906 0.000 -0.0406 0.03663 0.02736 -0.0445 1.0000 1.0000 0.250 -0.0169 0.03773 0.02776 -0.0461 1.0000 1.0000 0.500 0.0027 0.03884 0.02845 -0.0467 1.0000 1.0000 0.750 0.0498 0.04087 0.03002 -0.0524 0.9867 1.0000 1.000 0.0901 0.04272 0.03151 -0.0567 0.9742 1.0000 1.250 0.1283 0.04456 0.03306 -0.0605 0.9623 1.0000 1.500 0.1582 0.04613 0.03441 -0.0629 0.9525 1.0000 1.750 0.1845 0.04766 0.03576 -0.0647 0.9426 1.0000 2.000 0.2254 0.04975 0.03764 -0.0688 0.9312 1.0000 2.250 0.2362 0.05088 0.03868 -0.0680 0.9248 1.0000 2.500 0.2736 0.05292 0.04057 -0.0715 0.9139 1.0000 2.750 0.2848 0.05425 0.04183 -0.0709 0.9082 1.0000 3.000 0.3017 0.05582 0.04333 -0.0713 0.9025 1.0000 3.250 0.3332 0.05789 0.04531 -0.0738 0.8922 1.0000 3.500 0.3432 0.05949 0.04686 -0.0733 0.8895 1.0000 3.750 0.3552 0.06124 0.04859 -0.0732 0.8876 1.0000 4.000 0.3670 0.06315 0.05047 -0.0732 0.8879 1.0000 4.250 0.3820 0.06548 0.05277 -0.0740 0.8923 1.0000