XFOIL Version 6.94 Calculated polar for: NASA SC(2)-0414 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.2542 0.01321 0.00854 -0.0725 0.8734 0.7123 0.500 0.3128 0.01299 0.00826 -0.0724 0.8484 0.7188 1.000 0.3686 0.01286 0.00810 -0.0716 0.8203 0.7226 1.500 0.4268 0.01270 0.00787 -0.0715 0.7811 0.7280 2.000 0.4835 0.01262 0.00762 -0.0711 0.7198 0.7320 2.500 0.5312 0.01325 0.00754 -0.0688 0.5429 0.7357 3.000 0.5762 0.01504 0.00799 -0.0679 0.2876 0.7408 3.500 0.6230 0.01626 0.00862 -0.0665 0.1762 0.7438 4.000 0.6787 0.01705 0.00910 -0.0670 0.1348 0.7485 4.500 0.7320 0.01777 0.00973 -0.0668 0.1141 0.7520 5.000 0.7824 0.01867 0.01052 -0.0658 0.0997 0.7561 5.500 0.8379 0.01939 0.01119 -0.0661 0.0875 0.7610 6.000 0.8863 0.02037 0.01216 -0.0648 0.0787 0.7644 6.500 0.9409 0.02125 0.01307 -0.0650 0.0715 0.7695 7.000 0.9888 0.02234 0.01426 -0.0635 0.0659 0.7730 7.500 1.0376 0.02357 0.01545 -0.0625 0.0609 0.7772 8.000 1.0872 0.02491 0.01694 -0.0616 0.0572 0.7816 8.500 1.1335 0.02603 0.01817 -0.0601 0.0533 0.7858 9.000 1.1795 0.02825 0.02046 -0.0591 0.0492 0.7907 9.500 1.2209 0.02995 0.02250 -0.0567 0.0460 0.7946 10.000 1.2629 0.03219 0.02495 -0.0551 0.0438 0.8001 10.500 1.2973 0.03447 0.02743 -0.0522 0.0422 0.8039 11.000 1.3232 0.03928 0.03260 -0.0492 0.0403 0.8090 11.500 1.3331 0.04406 0.03791 -0.0441 0.0398 0.8133 12.000 1.3289 0.04820 0.04249 -0.0375 0.0396 0.8186 12.500 1.3131 0.05373 0.04851 -0.0317 0.0394 0.8228 13.000 1.2824 0.06116 0.05648 -0.0279 0.0394 0.8273 13.500 1.2341 0.07188 0.06772 -0.0280 0.0394 0.8311 14.000 1.1510 0.09062 0.08702 -0.0367 0.0398 0.8333