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 = 0.060 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1898 0.03625 0.02670 -0.0651 0.9688 0.2000 -2.750 -0.1470 0.03550 0.02562 -0.0685 0.9585 0.2167 -2.500 -0.0979 0.03472 0.02480 -0.0730 0.9497 0.2385 -2.250 -0.0637 0.03417 0.02424 -0.0749 0.9378 0.2603 -2.000 -0.0203 0.03373 0.02380 -0.0781 0.9283 0.2905 -1.750 0.0168 0.03324 0.02338 -0.0803 0.9173 0.3245 -1.500 0.0521 0.03284 0.02311 -0.0820 0.9065 0.3670 -1.250 0.0950 0.03227 0.02280 -0.0847 0.8972 0.4315 -1.000 0.1241 0.03183 0.02284 -0.0851 0.8860 0.5179 -0.750 0.1570 0.03010 0.02276 -0.0831 0.8777 1.0000 -0.500 0.1871 0.03067 0.02275 -0.0846 0.8648 1.0000 -0.250 0.2385 0.03099 0.02265 -0.0889 0.8570 1.0000 0.250 0.2903 0.03217 0.02340 -0.0895 0.8335 1.0000 0.750 0.3458 0.03338 0.02427 -0.0905 0.8120 1.0000 1.000 0.3819 0.03374 0.02449 -0.0920 0.8029 1.0000 1.250 0.4009 0.03463 0.02527 -0.0913 0.7916 1.0000 1.500 0.4385 0.03493 0.02545 -0.0929 0.7835 1.0000 1.750 0.4537 0.03600 0.02644 -0.0918 0.7722 1.0000 2.000 0.4914 0.03624 0.02660 -0.0932 0.7644 1.0000 2.250 0.5040 0.03750 0.02781 -0.0919 0.7533 1.0000 2.500 0.5399 0.03778 0.02802 -0.0930 0.7455 1.0000 3.000 0.5850 0.03952 0.02969 -0.0923 0.7267 1.0000 3.500 0.6275 0.04145 0.03159 -0.0914 0.7078 1.0000 4.000 0.6669 0.04363 0.03376 -0.0902 0.6890 1.0000 4.250 0.7175 0.04289 0.03304 -0.0920 0.6840 1.0000 4.500 0.7042 0.04601 0.03617 -0.0888 0.6700 1.0000 4.750 0.7495 0.04551 0.03570 -0.0900 0.6642 1.0000 5.000 0.7402 0.04854 0.03874 -0.0874 0.6507 1.0000