XFOIL Version 6.94 Calculated polar for: EPPLER 858 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 0.5133 0.01536 0.00716 -0.1004 0.4456 0.4876 0.500 0.5594 0.01550 0.00723 -0.0984 0.4381 0.4955 1.000 0.6079 0.01557 0.00726 -0.0971 0.4306 0.5035 1.500 0.6620 0.01595 0.00757 -0.0969 0.4231 0.5119 2.000 0.7064 0.01615 0.00776 -0.0947 0.4172 0.5202 2.500 0.7525 0.01629 0.00795 -0.0930 0.4113 0.5297 3.000 0.8055 0.01671 0.00827 -0.0927 0.4054 0.5396 3.500 0.8572 0.01716 0.00875 -0.0923 0.4000 0.5498 4.000 0.8957 0.01745 0.00914 -0.0891 0.3950 0.5614 4.500 0.9389 0.01777 0.00951 -0.0870 0.3900 0.5733 5.000 0.9849 0.01818 0.00994 -0.0856 0.3855 0.5872 5.500 1.0490 0.01899 0.01074 -0.0878 0.3798 0.6046 6.000 1.0798 0.01937 0.01131 -0.0835 0.3768 0.6228 6.500 1.1128 0.01978 0.01191 -0.0798 0.3725 0.6454 7.000 1.1490 0.02022 0.01253 -0.0768 0.3683 0.6771 7.500 1.1877 0.02062 0.01314 -0.0742 0.3645 0.7276 8.500 1.2988 0.02210 0.01512 -0.0763 0.3563 1.0000 9.000 1.3197 0.02292 0.01606 -0.0713 0.3528 1.0000 9.500 1.3462 0.02387 0.01708 -0.0675 0.3490 1.0000 10.000 1.3773 0.02482 0.01808 -0.0646 0.3454 1.0000 10.500 1.4156 0.02571 0.01896 -0.0629 0.3420 1.0000 11.000 1.4663 0.02695 0.02015 -0.0633 0.3376 1.0000 11.500 1.4726 0.02863 0.02203 -0.0575 0.3345 1.0000 12.000 1.4699 0.03073 0.02434 -0.0514 0.3311 1.0000 12.500 1.4725 0.03304 0.02683 -0.0466 0.3273 1.0000 13.000 1.4882 0.03500 0.02888 -0.0436 0.3236 1.0000 13.500 1.5279 0.03595 0.02982 -0.0430 0.3200 1.0000 14.000 1.5532 0.03847 0.03241 -0.0415 0.3158 1.0000 14.500 1.4440 0.04838 0.04281 -0.0325 0.3104 1.0000 15.000 1.3407 0.06182 0.05657 -0.0293 0.3015 1.0000 15.500 1.4179 0.05910 0.05379 -0.0295 0.2997 1.0000