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 = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2054 0.05346 0.04228 -0.0261 1.0001 0.4447 -2.750 -0.1759 0.05083 0.03944 -0.0285 1.0001 0.4588 -2.500 -0.1465 0.04854 0.03695 -0.0303 1.0001 0.4787 -2.250 -0.1182 0.04654 0.03482 -0.0313 1.0001 0.5046 -2.000 -0.0898 0.04473 0.03292 -0.0321 1.0001 0.5374 -1.750 -0.0651 0.04310 0.03139 -0.0314 1.0001 0.5804 -1.500 -0.0418 0.04152 0.03003 -0.0300 1.0001 0.6385 -1.250 -0.0253 0.03981 0.02889 -0.0263 1.0001 0.7217 -1.000 0.0381 0.03698 0.02577 -0.0375 1.0001 0.9999 -0.750 0.0743 0.03772 0.02518 -0.0408 1.0001 0.9999 -0.500 0.0982 0.03848 0.02537 -0.0413 1.0001 0.9999 -0.250 0.1196 0.03936 0.02588 -0.0414 1.0001 0.9999 0.000 0.1391 0.04041 0.02670 -0.0416 1.0001 0.9999 0.250 0.1559 0.04177 0.02794 -0.0419 1.0001 0.9999 0.500 0.1663 0.04387 0.03007 -0.0427 1.0001 0.9999 0.750 0.1546 0.04856 0.03505 -0.0452 1.0001 0.9999 1.000 0.3098 0.05330 0.03920 -0.0760 0.8849 0.9999 1.250 0.3798 0.05503 0.04058 -0.0849 0.8338 0.9999 1.500 0.4206 0.05677 0.04207 -0.0886 0.7989 0.9999 1.750 0.4710 0.05810 0.04313 -0.0928 0.7697 0.9999 2.000 0.4990 0.06021 0.04505 -0.0944 0.7495 0.9999 2.250 0.5146 0.06297 0.04767 -0.0949 0.7363 0.9999 2.500 0.5207 0.06619 0.05077 -0.0947 0.7277 0.9999 2.750 0.5243 0.06959 0.05405 -0.0944 0.7212 0.9999 3.000 0.5529 0.07220 0.05651 -0.0965 0.7126 0.9999 3.250 0.5458 0.07595 0.06019 -0.0953 0.7092 0.9999 3.500 0.5455 0.07941 0.06357 -0.0947 0.7053 0.9999 3.750 0.5506 0.08275 0.06682 -0.0947 0.7019 0.9999 4.000 0.5529 0.08629 0.07028 -0.0948 0.7020 0.9999 4.250 0.5539 0.09005 0.07398 -0.0951 0.7070 0.9999 4.500 0.5657 0.09387 0.07772 -0.0966 0.7124 0.9999 4.750 0.5469 0.09772 0.08156 -0.0954 0.7281 0.9999 5.000 0.5364 0.10142 0.08523 -0.0951 0.7461 0.9999