XFOIL Version 6.94 Calculated polar for: NACA 4412 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 -0.2632 0.04319 0.03392 -0.0489 1.0000 0.2319 -2.750 -0.2374 0.04138 0.03176 -0.0502 1.0000 0.2398 -2.500 -0.2126 0.03988 0.02999 -0.0512 1.0000 0.2517 -2.250 -0.1857 0.03868 0.02829 -0.0525 1.0000 0.2652 -2.000 -0.1639 0.03783 0.02740 -0.0527 1.0000 0.2821 -1.750 -0.1406 0.03709 0.02650 -0.0531 1.0000 0.3009 -1.500 -0.1171 0.03655 0.02580 -0.0536 1.0000 0.3252 -1.250 -0.0949 0.03613 0.02538 -0.0537 1.0000 0.3534 -1.000 -0.0705 0.03583 0.02505 -0.0542 1.0000 0.3905 -0.750 -0.0456 0.03561 0.02493 -0.0547 1.0000 0.4418 -0.500 -0.0204 0.03530 0.02505 -0.0550 1.0000 0.5194 -0.250 0.0024 0.03345 0.02515 -0.0521 0.9930 0.9126 0.000 0.0582 0.03486 0.02528 -0.0601 0.9782 1.0000 0.250 0.1053 0.03643 0.02632 -0.0652 0.9655 1.0000 0.500 0.1392 0.03763 0.02718 -0.0680 0.9525 1.0000 0.750 0.1705 0.03885 0.02814 -0.0702 0.9403 1.0000 1.000 0.2078 0.04025 0.02928 -0.0734 0.9286 1.0000 1.250 0.2413 0.04154 0.03036 -0.0759 0.9167 1.0000 1.500 0.2645 0.04274 0.03141 -0.0768 0.9057 1.0000 1.750 0.3040 0.04421 0.03271 -0.0801 0.8947 1.0000 2.000 0.3213 0.04536 0.03376 -0.0799 0.8839 1.0000 2.250 0.3503 0.04677 0.03506 -0.0816 0.8736 1.0000 2.500 0.3756 0.04810 0.03631 -0.0826 0.8631 1.0000 2.750 0.3961 0.04952 0.03766 -0.0830 0.8533 1.0000 3.000 0.4268 0.05098 0.03906 -0.0848 0.8431 1.0000 3.250 0.4400 0.05244 0.04049 -0.0842 0.8335 1.0000 3.500 0.4765 0.05400 0.04201 -0.0867 0.8233 1.0000 3.750 0.4811 0.05558 0.04357 -0.0851 0.8145 1.0000 4.000 0.5210 0.05719 0.04517 -0.0878 0.8036 1.0000 4.250 0.5193 0.05893 0.04692 -0.0857 0.7957 1.0000 4.500 0.5561 0.06059 0.04858 -0.0880 0.7845 1.0000 4.750 0.5550 0.06250 0.05051 -0.0861 0.7770 1.0000 5.000 0.5899 0.06425 0.05230 -0.0880 0.7657 1.0000