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 = 0.080 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0693 0.03007 0.02063 -0.0828 0.9201 0.1946 -2.750 -0.0203 0.02927 0.01973 -0.0867 0.9138 0.2170 -2.500 0.0076 0.02880 0.01929 -0.0870 0.9016 0.2368 -2.250 0.0544 0.02801 0.01868 -0.0905 0.8955 0.2672 -2.000 0.0816 0.02769 0.01835 -0.0906 0.8831 0.2969 -1.750 0.1281 0.02680 0.01773 -0.0938 0.8770 0.3432 -1.500 0.1533 0.02640 0.01756 -0.0934 0.8647 0.3891 -1.250 0.1983 0.02536 0.01704 -0.0961 0.8589 0.4837 -0.750 0.2765 0.02364 0.01663 -0.0973 0.8420 1.0000 -0.250 0.3481 0.02363 0.01611 -0.0999 0.8232 1.0000 0.000 0.3971 0.02318 0.01544 -0.1030 0.8186 1.0000 0.500 0.4608 0.02315 0.01512 -0.1038 0.7997 1.0000 0.750 0.4772 0.02369 0.01557 -0.1021 0.7857 1.0000 1.000 0.5216 0.02321 0.01495 -0.1041 0.7808 1.0000 1.250 0.5373 0.02385 0.01551 -0.1023 0.7671 1.0000 1.500 0.5796 0.02340 0.01495 -0.1039 0.7620 1.0000 1.750 0.5949 0.02412 0.01563 -0.1021 0.7488 1.0000 2.000 0.6356 0.02370 0.01511 -0.1035 0.7433 1.0000 2.250 0.6508 0.02450 0.01589 -0.1017 0.7307 1.0000 2.500 0.6894 0.02416 0.01547 -0.1027 0.7247 1.0000 2.750 0.7055 0.02498 0.01629 -0.1011 0.7129 1.0000 3.000 0.7416 0.02476 0.01600 -0.1018 0.7063 1.0000 3.250 0.7594 0.02554 0.01678 -0.1005 0.6954 1.0000 3.500 0.7925 0.02546 0.01668 -0.1008 0.6880 1.0000 3.750 0.8137 0.02610 0.01733 -0.0998 0.6784 1.0000 4.000 0.8425 0.02625 0.01747 -0.0996 0.6698 1.0000 4.250 0.8702 0.02655 0.01775 -0.0993 0.6618 1.0000 4.500 0.8918 0.02712 0.01837 -0.0984 0.6517 1.0000 4.750 0.9316 0.02674 0.01793 -0.0994 0.6458 1.0000 5.000 0.9409 0.02800 0.01930 -0.0971 0.6335 1.0000