XFOIL Version 6.94 Calculated polar for: EPPLER 541 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.500 0.3074 0.01600 0.01119 -0.0460 0.7588 0.8103 1.000 0.3677 0.01590 0.01104 -0.0472 0.7538 0.8117 1.500 0.4114 0.01567 0.01088 -0.0454 0.7468 0.8136 2.000 0.4607 0.01537 0.01060 -0.0447 0.7396 0.8158 2.500 0.5173 0.01502 0.01021 -0.0456 0.7332 0.8180 3.000 0.5565 0.01461 0.00989 -0.0434 0.7244 0.8207 3.500 0.6089 0.01412 0.00941 -0.0438 0.7156 0.8237 4.000 0.6598 0.01367 0.00902 -0.0439 0.7056 0.8256 4.500 0.7128 0.01319 0.00857 -0.0443 0.6936 0.8273 5.000 0.7602 0.01278 0.00826 -0.0434 0.6790 0.8291 5.500 0.8029 0.01229 0.00787 -0.0411 0.6603 0.8308 6.000 0.8392 0.01194 0.00759 -0.0376 0.6317 0.8322 6.500 0.8615 0.01178 0.00738 -0.0312 0.5844 0.8340 7.000 0.8573 0.01220 0.00751 -0.0200 0.5129 0.8364 7.500 0.8460 0.01344 0.00840 -0.0087 0.4360 0.8392 8.000 0.8371 0.01506 0.00971 0.0011 0.3625 0.8425 8.500 0.8387 0.01679 0.01115 0.0082 0.2949 0.8448 9.000 0.8464 0.01859 0.01270 0.0138 0.2335 0.8466 9.500 0.8588 0.02044 0.01430 0.0184 0.1784 0.8480 10.000 0.8732 0.02221 0.01587 0.0224 0.1308 0.8499 10.500 0.8886 0.02408 0.01757 0.0262 0.0902 0.8518 11.000 0.9027 0.02621 0.01955 0.0298 0.0595 0.8538 11.500 0.9185 0.02841 0.02170 0.0329 0.0394 0.8558 12.000 0.9361 0.03063 0.02393 0.0355 0.0290 0.8583 14.000 0.9905 0.04249 0.03622 0.0436 0.0081 0.8669 15.000 0.9976 0.05140 0.04555 0.0460 0.0062 0.8713 15.500 0.9908 0.05740 0.05177 0.0465 0.0056 0.8737 16.000 0.9910 0.06314 0.05778 0.0462 0.0053 0.8762 16.500 0.9873 0.06982 0.06475 0.0451 0.0050 0.8787 17.000 0.9836 0.07695 0.07214 0.0432 0.0050 0.8811 17.500 0.9751 0.08535 0.08085 0.0402 0.0048 0.8835 18.000 0.9622 0.09501 0.09082 0.0359 0.0047 0.8856 18.500 0.9461 0.10581 0.10195 0.0306 0.0046 0.8875 19.000 0.9247 0.11809 0.11457 0.0239 0.0046 0.8901 19.500 0.8985 0.13181 0.12865 0.0161 0.0047 0.8932 20.000 0.8637 0.14817 0.14538 0.0067 0.0049 0.8962 20.500 0.8233 0.16694 0.16446 -0.0038 0.0052 0.8993