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 = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0762 0.04390 0.03220 -0.0477 0.9999 0.3518 -2.750 -0.0487 0.04259 0.03071 -0.0479 0.9999 0.3708 -2.500 -0.0203 0.04149 0.02941 -0.0481 0.9999 0.3944 -2.250 0.0071 0.04047 0.02841 -0.0480 0.9999 0.4245 -2.000 0.0344 0.03954 0.02763 -0.0478 0.9999 0.4669 -1.750 0.0614 0.03848 0.02705 -0.0472 0.9999 0.5332 -1.500 0.0790 0.03648 0.02662 -0.0433 0.9999 0.7039 -1.250 0.1055 0.03603 0.02537 -0.0438 0.9999 1.0001 -1.000 0.1304 0.03688 0.02558 -0.0438 0.9999 1.0001 -0.750 0.1526 0.03780 0.02617 -0.0437 0.9999 1.0001 -0.500 0.1732 0.03887 0.02703 -0.0437 0.9999 1.0001 -0.250 0.1917 0.04021 0.02829 -0.0438 0.9999 1.0001 0.000 0.2032 0.04244 0.03062 -0.0444 0.9999 1.0001 0.250 0.4810 0.03579 0.02337 -0.0865 0.7927 1.0001 0.500 0.5313 0.03536 0.02237 -0.0876 0.7254 1.0001 0.750 0.5651 0.03596 0.02250 -0.0872 0.6846 1.0001 1.000 0.5953 0.03684 0.02299 -0.0867 0.6544 1.0001 1.250 0.6248 0.03779 0.02357 -0.0863 0.6309 1.0001 1.500 0.6531 0.03897 0.02451 -0.0863 0.6112 1.0001 1.750 0.6806 0.04051 0.02595 -0.0870 0.5957 1.0001 2.000 0.7086 0.04211 0.02745 -0.0877 0.5844 1.0001 2.250 0.7360 0.04382 0.02911 -0.0885 0.5743 1.0001 2.500 0.7623 0.04575 0.03101 -0.0894 0.5654 1.0001 2.750 0.7873 0.04784 0.03313 -0.0904 0.5562 1.0001 3.000 0.8146 0.04962 0.03483 -0.0909 0.5481 1.0001 3.250 0.8320 0.05272 0.03809 -0.0921 0.5411 1.0001 3.500 0.8511 0.05572 0.04118 -0.0931 0.5365 1.0001 3.750 0.8688 0.05895 0.04448 -0.0940 0.5335 1.0001 4.000 0.8849 0.06241 0.04801 -0.0949 0.5312 1.0001 4.250 0.8929 0.06670 0.05238 -0.0956 0.5294 1.0001 4.500 0.8851 0.07251 0.05832 -0.0959 0.5292 1.0001 4.750 0.8738 0.07855 0.06446 -0.0960 0.5298 1.0001