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 = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1773 0.04969 0.03904 -0.0362 1.0001 0.3784 -2.750 -0.1456 0.04730 0.03629 -0.0388 1.0001 0.3923 -2.500 -0.1146 0.04531 0.03396 -0.0407 1.0001 0.4101 -2.250 -0.0872 0.04360 0.03213 -0.0411 1.0001 0.4326 -2.000 -0.0609 0.04213 0.03062 -0.0410 1.0001 0.4617 -1.750 -0.0334 0.04078 0.02925 -0.0411 1.0001 0.4990 -1.500 -0.0071 0.03952 0.02815 -0.0406 1.0001 0.5504 -1.250 0.0154 0.03815 0.02729 -0.0387 1.0001 0.6255 -1.000 0.0270 0.03623 0.02649 -0.0332 1.0001 0.7652 -0.750 0.0788 0.03547 0.02442 -0.0417 1.0001 0.9999 -0.500 0.1035 0.03627 0.02460 -0.0423 1.0001 0.9999 -0.250 0.1251 0.03720 0.02518 -0.0425 1.0001 0.9999 0.000 0.1442 0.03834 0.02613 -0.0427 1.0001 0.9999 0.250 0.1589 0.03998 0.02773 -0.0432 1.0001 0.9999 0.750 0.4389 0.04186 0.02893 -0.0891 0.7959 0.9999 1.000 0.4980 0.04144 0.02812 -0.0921 0.7486 0.9999 1.250 0.5397 0.04184 0.02818 -0.0931 0.7158 0.9999 1.500 0.5748 0.04279 0.02883 -0.0938 0.6929 0.9999 1.750 0.6027 0.04432 0.03017 -0.0942 0.6743 0.9999 2.000 0.6219 0.04663 0.03238 -0.0945 0.6587 0.9999 2.250 0.6408 0.04899 0.03464 -0.0948 0.6454 0.9999 2.500 0.6714 0.05032 0.03576 -0.0951 0.6331 0.9999 2.750 0.6785 0.05378 0.03923 -0.0952 0.6221 0.9999 3.000 0.7053 0.05566 0.04097 -0.0957 0.6128 0.9999 3.250 0.7004 0.06043 0.04581 -0.0959 0.6084 0.9999 3.500 0.6956 0.06520 0.05059 -0.0961 0.6063 0.9999 3.750 0.6859 0.07028 0.05568 -0.0960 0.6064 0.9999 4.000 0.6741 0.07534 0.06072 -0.0958 0.6078 0.9999 4.250 0.6647 0.08011 0.06544 -0.0954 0.6095 0.9999 4.500 0.6612 0.08450 0.06979 -0.0955 0.6108 0.9999 4.750 0.6619 0.08867 0.07392 -0.0958 0.6123 0.9999