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 = 0.008 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0073 0.06006 0.04324 -0.0280 1.0001 0.9999 -2.750 -0.0008 0.05821 0.04161 -0.0271 1.0001 0.9999 -2.500 -0.0120 0.05632 0.03998 -0.0259 1.0001 0.9999 -2.250 -0.0272 0.05436 0.03831 -0.0241 1.0001 0.9999 -2.000 -0.0462 0.05230 0.03656 -0.0221 1.0001 0.9999 -1.750 -0.0598 0.05018 0.03447 -0.0222 1.0001 0.9999 -1.500 -0.0289 0.04898 0.03183 -0.0316 1.0001 0.9999 -1.250 0.0048 0.04920 0.03062 -0.0357 1.0001 0.9999 -1.000 0.0310 0.04965 0.03010 -0.0370 1.0001 0.9999 -0.750 0.0545 0.05020 0.02993 -0.0375 1.0001 0.9999 -0.500 0.0764 0.05083 0.02997 -0.0377 1.0001 0.9999 -0.250 0.0972 0.05156 0.03024 -0.0378 1.0001 0.9999 0.000 0.1170 0.05240 0.03071 -0.0378 1.0001 0.9999 0.250 0.1357 0.05337 0.03139 -0.0379 1.0001 0.9999 0.500 0.1530 0.05451 0.03229 -0.0380 1.0001 0.9999 0.750 0.1686 0.05586 0.03349 -0.0382 1.0001 0.9999 1.000 0.1818 0.05752 0.03504 -0.0386 1.0001 0.9999 1.250 0.1918 0.05959 0.03706 -0.0393 1.0001 0.9999 1.500 0.1978 0.06217 0.03961 -0.0402 1.0001 0.9999 1.750 0.2009 0.06518 0.04256 -0.0415 1.0001 0.9999 2.000 0.2037 0.06833 0.04562 -0.0430 1.0001 0.9999 2.250 0.2083 0.07142 0.04857 -0.0445 1.0001 0.9999 2.500 0.2146 0.07438 0.05138 -0.0460 1.0001 0.9999 2.750 0.2223 0.07724 0.05406 -0.0474 1.0001 0.9999 3.000 0.2311 0.08002 0.05666 -0.0487 1.0001 0.9999 3.250 0.2406 0.08274 0.05921 -0.0500 1.0001 0.9999 3.500 0.2508 0.08543 0.06171 -0.0512 1.0001 0.9999 3.750 0.2614 0.08810 0.06420 -0.0524 1.0001 0.9999 4.000 0.2723 0.09075 0.06668 -0.0535 1.0001 0.9999 4.250 0.2835 0.09340 0.06917 -0.0546 1.0001 0.9999 4.500 0.2949 0.09605 0.07167 -0.0557 1.0001 0.9999 4.750 0.3066 0.09871 0.07417 -0.0568 1.0001 0.9999 5.000 0.3183 0.10138 0.07670 -0.0579 1.0001 0.9999