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 -0.1583 0.04684 0.03650 -0.0423 1.0001 0.3319 -2.750 -0.1293 0.04476 0.03411 -0.0438 1.0001 0.3437 -2.500 -0.1022 0.04300 0.03217 -0.0443 1.0001 0.3598 -2.250 -0.0734 0.04148 0.03041 -0.0451 1.0001 0.3800 -2.000 -0.0444 0.04019 0.02890 -0.0457 1.0001 0.4060 -1.750 -0.0175 0.03904 0.02776 -0.0456 1.0001 0.4391 -1.500 0.0101 0.03800 0.02683 -0.0455 1.0001 0.4838 -1.250 0.0367 0.03692 0.02611 -0.0449 1.0001 0.5491 -1.000 0.0569 0.03537 0.02556 -0.0420 1.0001 0.6673 -0.750 0.0819 0.03380 0.02394 -0.0423 1.0001 0.9999 -0.500 0.1080 0.03465 0.02407 -0.0431 1.0001 0.9999 -0.250 0.1295 0.03562 0.02471 -0.0434 1.0001 0.9999 0.000 0.1480 0.03687 0.02581 -0.0437 1.0001 0.9999 0.500 0.4500 0.03400 0.02214 -0.0866 0.7762 0.9999 0.750 0.4983 0.03356 0.02112 -0.0870 0.7165 0.9999 1.000 0.5320 0.03418 0.02129 -0.0867 0.6787 0.9999 1.250 0.5636 0.03502 0.02171 -0.0865 0.6525 0.9999 1.500 0.5924 0.03621 0.02263 -0.0866 0.6315 0.9999 1.750 0.6217 0.03733 0.02349 -0.0867 0.6145 0.9999 2.000 0.6475 0.03884 0.02489 -0.0870 0.5985 0.9999 2.250 0.6727 0.04043 0.02637 -0.0873 0.5848 0.9999 2.500 0.7011 0.04171 0.02744 -0.0873 0.5735 0.9999 2.750 0.7235 0.04401 0.02983 -0.0884 0.5642 0.9999 3.000 0.7511 0.04579 0.03153 -0.0892 0.5574 0.9999 3.250 0.7717 0.04848 0.03428 -0.0902 0.5520 0.9999 3.500 0.7850 0.05186 0.03779 -0.0912 0.5468 0.9999 3.750 0.8007 0.05484 0.04080 -0.0917 0.5413 0.9999 4.000 0.8286 0.05657 0.04243 -0.0918 0.5353 0.9999 4.250 0.8234 0.06171 0.04774 -0.0922 0.5328 0.9999 4.500 0.8167 0.06708 0.05321 -0.0927 0.5328 0.9999 4.750 0.8121 0.07239 0.05858 -0.0932 0.5347 0.9999