XFOIL Version 6.94 Calculated polar for: NACA 4412 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.070 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1315 0.03236 0.02295 -0.0727 0.9651 0.1727 -2.750 -0.0921 0.03157 0.02190 -0.0753 0.9545 0.1881 -2.500 -0.0481 0.03078 0.02101 -0.0787 0.9459 0.2081 -2.250 -0.0072 0.03006 0.02039 -0.0816 0.9358 0.2321 -2.000 0.0311 0.02959 0.01991 -0.0838 0.9255 0.2620 -1.750 0.0753 0.02899 0.01939 -0.0870 0.9166 0.3029 -1.500 0.1077 0.02848 0.01910 -0.0882 0.9052 0.3476 -1.250 0.1541 0.02762 0.01863 -0.0914 0.8977 0.4279 -1.000 0.1816 0.02677 0.01863 -0.0913 0.8862 0.5698 -0.750 0.2250 0.02558 0.01822 -0.0919 0.8789 1.0000 -0.500 0.2527 0.02596 0.01827 -0.0923 0.8654 1.0000 -0.250 0.2945 0.02613 0.01815 -0.0948 0.8563 1.0000 0.000 0.3278 0.02636 0.01818 -0.0960 0.8448 1.0000 0.250 0.3578 0.02671 0.01836 -0.0966 0.8334 1.0000 0.500 0.4009 0.02667 0.01815 -0.0990 0.8252 1.0000 0.750 0.4248 0.02715 0.01851 -0.0986 0.8129 1.0000 1.000 0.4698 0.02694 0.01817 -0.1010 0.8063 1.0000 1.250 0.4900 0.02756 0.01871 -0.1001 0.7935 1.0000 1.500 0.5340 0.02723 0.01828 -0.1021 0.7874 1.0000 1.750 0.5512 0.02801 0.01901 -0.1008 0.7745 1.0000 2.000 0.5946 0.02762 0.01855 -0.1025 0.7688 1.0000 2.250 0.6097 0.02854 0.01944 -0.1010 0.7559 1.0000 2.500 0.6522 0.02812 0.01897 -0.1025 0.7503 1.0000 2.750 0.6658 0.02917 0.02001 -0.1008 0.7374 1.0000 3.000 0.7076 0.02871 0.01952 -0.1021 0.7318 1.0000 3.250 0.7198 0.02991 0.02073 -0.1003 0.7191 1.0000 3.500 0.7614 0.02940 0.02020 -0.1014 0.7133 1.0000 3.750 0.7724 0.03073 0.02156 -0.0995 0.7008 1.0000 4.000 0.8138 0.03018 0.02102 -0.1005 0.6949 1.0000 4.250 0.8239 0.03159 0.02248 -0.0986 0.6823 1.0000 4.500 0.8655 0.03099 0.02190 -0.0995 0.6763 1.0000 4.750 0.8747 0.03249 0.02345 -0.0975 0.6636 1.0000 5.000 0.9174 0.03178 0.02278 -0.0984 0.6576 1.0000