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 = 0.080 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0816 0.02972 0.02029 -0.0800 0.9438 0.1692 -2.750 -0.0387 0.02887 0.01938 -0.0831 0.9354 0.1875 -2.500 -0.0009 0.02822 0.01866 -0.0852 0.9258 0.2078 -2.250 0.0417 0.02743 0.01804 -0.0882 0.9173 0.2342 -2.000 0.0767 0.02699 0.01759 -0.0897 0.9069 0.2661 -1.750 0.1192 0.02627 0.01710 -0.0924 0.8989 0.3092 -1.500 0.1515 0.02574 0.01678 -0.0933 0.8881 0.3602 -1.250 0.1939 0.02475 0.01630 -0.0958 0.8806 0.4502 -1.000 0.2185 0.02352 0.01634 -0.0941 0.8712 0.6997 -0.500 0.3014 0.02305 0.01571 -0.0978 0.8518 1.0000 -0.250 0.3464 0.02289 0.01529 -0.1006 0.8442 1.0000 0.000 0.3747 0.02309 0.01533 -0.1006 0.8323 1.0000 0.250 0.4175 0.02283 0.01489 -0.1028 0.8252 1.0000 0.500 0.4430 0.02309 0.01503 -0.1024 0.8130 1.0000 0.750 0.4846 0.02278 0.01458 -0.1041 0.8063 1.0000 1.000 0.5080 0.02312 0.01483 -0.1034 0.7939 1.0000 1.250 0.5484 0.02277 0.01438 -0.1048 0.7874 1.0000 1.500 0.5703 0.02320 0.01475 -0.1038 0.7750 1.0000 1.750 0.6096 0.02285 0.01431 -0.1050 0.7686 1.0000 2.000 0.6298 0.02339 0.01481 -0.1039 0.7561 1.0000 2.250 0.6684 0.02304 0.01440 -0.1049 0.7498 1.0000 2.500 0.6872 0.02369 0.01503 -0.1036 0.7374 1.0000 2.750 0.7252 0.02336 0.01464 -0.1044 0.7311 1.0000 3.000 0.7430 0.02410 0.01539 -0.1031 0.7187 1.0000 3.250 0.7804 0.02380 0.01503 -0.1038 0.7124 1.0000 3.500 0.7974 0.02461 0.01588 -0.1024 0.7001 1.0000 3.750 0.8343 0.02434 0.01557 -0.1030 0.6937 1.0000 4.000 0.8505 0.02521 0.01650 -0.1015 0.6815 1.0000 4.250 0.8875 0.02494 0.01619 -0.1021 0.6750 1.0000 4.500 0.9029 0.02589 0.01721 -0.1005 0.6627 1.0000 4.750 0.9404 0.02558 0.01689 -0.1012 0.6561 1.0000 5.000 0.9547 0.02658 0.01799 -0.0994 0.6437 1.0000