XFOIL Version 6.94 Calculated polar for: NACA4412c 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1265 0.03899 0.02957 -0.0499 1.0001 0.2166 -2.750 -0.0976 0.03742 0.02758 -0.0506 1.0001 0.2260 -2.500 -0.0730 0.03635 0.02649 -0.0504 1.0001 0.2383 -2.250 -0.0461 0.03534 0.02534 -0.0506 1.0001 0.2525 -2.000 -0.0191 0.03458 0.02449 -0.0508 1.0001 0.2710 -1.750 0.0072 0.03398 0.02387 -0.0508 1.0001 0.2926 -1.500 0.0338 0.03354 0.02345 -0.0508 1.0001 0.3206 -1.250 0.0610 0.03315 0.02318 -0.0509 1.0001 0.3574 -1.000 0.0890 0.03274 0.02308 -0.0512 1.0001 0.4116 -0.750 0.1261 0.03117 0.02292 -0.0525 0.9969 0.5820 -0.500 0.2328 0.02762 0.01936 -0.0643 0.9492 0.9999 -0.250 0.3484 0.02388 0.01527 -0.0760 0.8592 0.9999 0.000 0.4057 0.02229 0.01266 -0.0758 0.6650 0.9999 0.250 0.4318 0.02318 0.01241 -0.0738 0.5717 0.9999 0.500 0.4591 0.02408 0.01265 -0.0733 0.5321 0.9999 0.750 0.4880 0.02490 0.01302 -0.0732 0.5067 0.9999 1.000 0.5177 0.02568 0.01344 -0.0734 0.4883 0.9999 1.250 0.5482 0.02647 0.01389 -0.0738 0.4741 0.9999 1.500 0.5797 0.02723 0.01449 -0.0745 0.4625 0.9999 1.750 0.6113 0.02809 0.01513 -0.0752 0.4543 0.9999 2.000 0.6422 0.02894 0.01592 -0.0759 0.4466 0.9999 2.250 0.6729 0.02987 0.01668 -0.0764 0.4401 0.9999 2.500 0.7029 0.03088 0.01760 -0.0769 0.4341 0.9999 2.750 0.7321 0.03188 0.01863 -0.0774 0.4276 0.9999 3.000 0.7613 0.03293 0.01958 -0.0778 0.4216 0.9999 3.250 0.7904 0.03415 0.02067 -0.0783 0.4168 0.9999 3.500 0.8184 0.03543 0.02212 -0.0788 0.4133 0.9999 3.750 0.8460 0.03682 0.02363 -0.0794 0.4100 0.9999 4.000 0.8732 0.03830 0.02520 -0.0799 0.4070 0.9999 4.250 0.9001 0.03979 0.02677 -0.0803 0.4036 0.9999 4.500 0.9270 0.04132 0.02826 -0.0805 0.4001 0.9999 4.750 0.9518 0.04316 0.03017 -0.0808 0.3968 0.9999 5.000 0.9733 0.04526 0.03256 -0.0810 0.3941 0.9999