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 = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0868 0.06391 0.04277 -0.0251 0.9999 1.0001 -2.750 -0.0840 0.06190 0.04023 -0.0278 0.9999 1.0001 -2.500 -0.0578 0.06108 0.03798 -0.0329 0.9999 1.0001 -2.250 -0.0292 0.06105 0.03651 -0.0357 0.9999 1.0001 -2.000 -0.0034 0.06127 0.03559 -0.0370 0.9999 1.0001 -1.750 0.0205 0.06161 0.03499 -0.0375 0.9999 1.0001 -1.500 0.0431 0.06204 0.03463 -0.0377 0.9999 1.0001 -1.250 0.0649 0.06255 0.03446 -0.0377 0.9999 1.0001 -1.000 0.0859 0.06314 0.03448 -0.0376 0.9999 1.0001 -0.750 0.1064 0.06382 0.03467 -0.0374 0.9999 1.0001 -0.500 0.1263 0.06459 0.03502 -0.0373 0.9999 1.0001 -0.250 0.1455 0.06546 0.03552 -0.0371 0.9999 1.0001 0.000 0.1640 0.06644 0.03620 -0.0370 0.9999 1.0001 0.250 0.1817 0.06754 0.03707 -0.0369 0.9999 1.0001 0.500 0.1984 0.06879 0.03813 -0.0369 0.9999 1.0001 0.750 0.2139 0.07021 0.03939 -0.0369 0.9999 1.0001 1.000 0.2281 0.07184 0.04092 -0.0371 0.9999 1.0001 1.250 0.2405 0.07372 0.04273 -0.0373 0.9999 1.0001 1.500 0.2510 0.07588 0.04484 -0.0378 0.9999 1.0001 1.750 0.2595 0.07834 0.04723 -0.0384 0.9999 1.0001 2.000 0.2664 0.08107 0.04990 -0.0393 0.9999 1.0001 2.250 0.2723 0.08399 0.05272 -0.0403 0.9999 1.0001 2.500 0.2781 0.08698 0.05560 -0.0414 0.9999 1.0001 2.750 0.2844 0.08998 0.05847 -0.0426 0.9999 1.0001 3.000 0.2914 0.09295 0.06128 -0.0438 0.9999 1.0001 3.250 0.2993 0.09589 0.06407 -0.0450 0.9999 1.0001 3.500 0.3079 0.09878 0.06679 -0.0462 0.9999 1.0001 3.750 0.3170 0.10164 0.06949 -0.0475 0.9999 1.0001 4.000 0.3267 0.10446 0.07216 -0.0487 0.9999 1.0001 4.250 0.3367 0.10725 0.07481 -0.0499 0.9999 1.0001 4.500 0.3470 0.11002 0.07744 -0.0511 0.9999 1.0001 4.750 0.3576 0.11278 0.08005 -0.0522 0.9999 1.0001 5.000 0.3685 0.11553 0.08267 -0.0534 0.9999 1.0001