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 = 0.090 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1049 0.03029 0.02167 -0.0519 0.9995 0.1342 -2.750 -0.0437 0.02854 0.01966 -0.0581 0.9907 0.1483 -2.500 0.0185 0.02688 0.01792 -0.0644 0.9802 0.1667 -2.250 0.0743 0.02530 0.01643 -0.0695 0.9681 0.1885 -2.000 0.1335 0.02364 0.01490 -0.0750 0.9538 0.2185 -1.750 0.1973 0.02159 0.01309 -0.0808 0.9330 0.2618 -1.500 0.2444 0.01985 0.01170 -0.0828 0.8956 0.3170 -1.250 0.2804 0.01804 0.01057 -0.0822 0.8222 0.4531 -1.000 0.2962 0.01600 0.00905 -0.0750 0.6188 0.9999 -0.500 0.3454 0.01809 0.00899 -0.0732 0.4025 0.9999 -0.250 0.3738 0.01863 0.00912 -0.0732 0.3854 0.9999 0.000 0.4025 0.01917 0.00929 -0.0732 0.3745 0.9999 0.250 0.4317 0.01966 0.00951 -0.0733 0.3659 0.9999 0.500 0.4608 0.02032 0.00981 -0.0735 0.3592 0.9999 0.750 0.4906 0.02078 0.01011 -0.0737 0.3537 0.9999 1.000 0.5204 0.02130 0.01047 -0.0740 0.3484 0.9999 1.250 0.5500 0.02191 0.01086 -0.0743 0.3434 0.9999 1.500 0.5798 0.02276 0.01143 -0.0747 0.3387 0.9999 1.750 0.6095 0.02322 0.01187 -0.0750 0.3346 0.9999 2.000 0.6391 0.02379 0.01237 -0.0754 0.3306 0.9999 2.250 0.6688 0.02443 0.01292 -0.0757 0.3273 0.9999 2.500 0.6984 0.02514 0.01353 -0.0761 0.3243 0.9999 2.750 0.7281 0.02598 0.01424 -0.0765 0.3215 0.9999 3.000 0.7576 0.02699 0.01514 -0.0770 0.3191 0.9999 3.250 0.7867 0.02762 0.01584 -0.0773 0.3168 0.9999 3.500 0.8156 0.02832 0.01661 -0.0776 0.3138 0.9999 3.750 0.8442 0.02909 0.01740 -0.0779 0.3107 0.9999 4.000 0.8729 0.02993 0.01823 -0.0782 0.3078 0.9999 4.250 0.9013 0.03086 0.01917 -0.0785 0.3056 0.9999 4.500 0.9297 0.03194 0.02023 -0.0788 0.3036 0.9999 4.750 0.9577 0.03336 0.02163 -0.0793 0.3017 0.9999 5.000 0.9848 0.03435 0.02281 -0.0794 0.3002 0.9999