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 = 0.012 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2056 0.05510 0.04183 -0.0044 0.9999 0.6496 -2.750 -0.1816 0.05269 0.03952 -0.0042 0.9999 0.6829 -2.500 -0.1582 0.05042 0.03741 -0.0034 0.9999 0.7296 -2.250 -0.1331 0.04814 0.03546 -0.0021 0.9999 0.8000 -2.000 -0.0375 0.04424 0.03164 -0.0223 0.9999 1.0001 -1.750 0.0322 0.04407 0.02890 -0.0391 0.9999 1.0001 -1.500 0.0622 0.04470 0.02840 -0.0404 0.9999 1.0001 -1.250 0.0867 0.04537 0.02835 -0.0406 0.9999 1.0001 -1.000 0.1093 0.04610 0.02856 -0.0405 0.9999 1.0001 -0.750 0.1308 0.04691 0.02898 -0.0404 0.9999 1.0001 -0.500 0.1514 0.04782 0.02959 -0.0403 0.9999 1.0001 -0.250 0.1709 0.04887 0.03041 -0.0402 0.9999 1.0001 0.000 0.1890 0.05010 0.03150 -0.0402 0.9999 1.0001 0.250 0.2050 0.05161 0.03295 -0.0403 0.9999 1.0001 0.500 0.2174 0.05361 0.03495 -0.0407 0.9999 1.0001 0.750 0.2227 0.05653 0.03798 -0.0416 0.9999 1.0001 1.000 0.2178 0.06079 0.04237 -0.0434 0.9999 1.0001 1.250 0.2137 0.06522 0.04681 -0.0456 0.9999 1.0001 1.500 0.2156 0.06902 0.05054 -0.0475 0.9999 1.0001 1.750 0.2212 0.07242 0.05381 -0.0492 0.9999 1.0001 2.000 0.2289 0.07557 0.05682 -0.0507 0.9999 1.0001 2.250 0.2380 0.07858 0.05968 -0.0521 0.9999 1.0001 2.500 0.2480 0.08149 0.06245 -0.0535 0.9999 1.0001 2.750 0.2586 0.08435 0.06515 -0.0547 0.9999 1.0001 3.000 0.2698 0.08717 0.06782 -0.0559 0.9999 1.0001 3.250 0.2826 0.08998 0.07048 -0.0574 0.9987 1.0001 3.500 0.2930 0.09277 0.07313 -0.0582 0.9999 1.0001 3.750 0.3049 0.09555 0.07578 -0.0593 0.9999 1.0001 4.000 0.3169 0.09834 0.07843 -0.0605 0.9999 1.0001 4.250 0.3291 0.10113 0.08110 -0.0616 0.9999 1.0001 4.500 0.3413 0.10392 0.08378 -0.0626 0.9999 1.0001 4.750 0.3536 0.10673 0.08648 -0.0637 0.9999 1.0001 5.000 0.3660 0.10954 0.08920 -0.0648 0.9999 1.0001