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 = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1476 0.04470 0.03468 -0.0452 1.0001 0.2963 -2.750 -0.1162 0.04269 0.03218 -0.0472 1.0001 0.3073 -2.500 -0.0914 0.04113 0.03055 -0.0471 1.0001 0.3223 -2.250 -0.0645 0.03976 0.02900 -0.0474 1.0001 0.3399 -2.000 -0.0360 0.03865 0.02768 -0.0478 1.0001 0.3643 -1.750 -0.0079 0.03765 0.02659 -0.0480 1.0001 0.3934 -1.500 0.0204 0.03678 0.02578 -0.0481 1.0001 0.4330 -1.250 0.0477 0.03592 0.02520 -0.0479 1.0001 0.4892 -1.000 0.0729 0.03474 0.02482 -0.0467 1.0001 0.5868 -0.500 0.1117 0.03339 0.02369 -0.0439 1.0001 0.9999 -0.250 0.1331 0.03441 0.02438 -0.0442 1.0001 0.9999 0.000 0.1503 0.03583 0.02570 -0.0446 1.0001 0.9999 0.250 0.4136 0.03030 0.01953 -0.0821 0.8112 0.9999 0.500 0.4694 0.02927 0.01770 -0.0826 0.7090 0.9999 0.750 0.5018 0.02982 0.01760 -0.0816 0.6577 0.9999 1.000 0.5314 0.03071 0.01802 -0.0812 0.6246 0.9999 1.250 0.5608 0.03168 0.01862 -0.0811 0.6009 0.9999 1.500 0.5899 0.03273 0.01938 -0.0813 0.5818 0.9999 1.750 0.6185 0.03384 0.02027 -0.0814 0.5656 0.9999 2.000 0.6475 0.03492 0.02112 -0.0816 0.5518 0.9999 2.250 0.6756 0.03616 0.02226 -0.0820 0.5397 0.9999 2.500 0.7045 0.03760 0.02362 -0.0828 0.5312 0.9999 2.750 0.7324 0.03924 0.02530 -0.0838 0.5235 0.9999 3.000 0.7616 0.04070 0.02666 -0.0844 0.5172 0.9999 3.250 0.7860 0.04277 0.02883 -0.0853 0.5111 0.9999 3.500 0.8092 0.04487 0.03103 -0.0860 0.5045 0.9999 3.750 0.8359 0.04651 0.03259 -0.0863 0.4982 0.9999 4.000 0.8568 0.04899 0.03515 -0.0869 0.4935 0.9999 4.250 0.8711 0.05236 0.03874 -0.0877 0.4911 0.9999 4.500 0.8804 0.05630 0.04288 -0.0885 0.4900 0.9999 4.750 0.8822 0.06103 0.04778 -0.0891 0.4901 0.9999 5.000 0.8752 0.06657 0.05350 -0.0895 0.4910 0.9999