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 = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2960 0.04440 0.03510 -0.0451 1.0000 0.2445 -2.750 -0.2692 0.04267 0.03284 -0.0469 1.0000 0.2548 -2.500 -0.2486 0.04148 0.03163 -0.0471 1.0000 0.2655 -2.250 -0.2239 0.04037 0.03017 -0.0482 1.0000 0.2799 -2.000 -0.2006 0.03968 0.02919 -0.0488 1.0000 0.2967 -1.750 -0.1792 0.03908 0.02854 -0.0490 1.0000 0.3154 -1.500 -0.1569 0.03863 0.02798 -0.0494 1.0000 0.3382 -1.250 -0.1357 0.03828 0.02766 -0.0495 1.0000 0.3637 -1.000 -0.1119 0.03815 0.02742 -0.0500 1.0000 0.3987 -0.750 -0.0625 0.03818 0.02769 -0.0546 0.9903 0.4645 -0.500 -0.0123 0.03803 0.02818 -0.0587 0.9791 0.5820 -0.250 0.0103 0.03646 0.02789 -0.0568 0.9677 1.0000 0.000 0.0538 0.03776 0.02830 -0.0618 0.9549 1.0000 0.250 0.0909 0.03909 0.02916 -0.0651 0.9429 1.0000 0.500 0.1361 0.04067 0.03033 -0.0696 0.9316 1.0000 0.750 0.1555 0.04162 0.03107 -0.0699 0.9202 1.0000 1.000 0.1898 0.04300 0.03219 -0.0724 0.9093 1.0000 1.250 0.2170 0.04423 0.03323 -0.0739 0.8988 1.0000 1.500 0.2425 0.04553 0.03436 -0.0750 0.8887 1.0000 1.750 0.2742 0.04690 0.03556 -0.0771 0.8784 1.0000 2.000 0.2923 0.04820 0.03675 -0.0771 0.8691 1.0000 2.250 0.3244 0.04964 0.03806 -0.0792 0.8591 1.0000 2.500 0.3392 0.05101 0.03936 -0.0788 0.8502 1.0000 2.750 0.3673 0.05249 0.04076 -0.0803 0.8408 1.0000 3.000 0.3837 0.05400 0.04219 -0.0801 0.8320 1.0000 3.250 0.4052 0.05554 0.04368 -0.0808 0.8234 1.0000 3.500 0.4267 0.05719 0.04529 -0.0813 0.8146 1.0000 3.750 0.4402 0.05880 0.04688 -0.0810 0.8065 1.0000 4.000 0.4692 0.06059 0.04863 -0.0825 0.7971 1.0000 4.250 0.4737 0.06235 0.05040 -0.0813 0.7906 1.0000 4.500 0.5112 0.06418 0.05220 -0.0836 0.7795 1.0000 4.750 0.5090 0.06615 0.05419 -0.0819 0.7747 1.0000 5.000 0.5214 0.06814 0.05620 -0.0817 0.7680 1.0000