XFOIL Version 6.94 Calculated polar for: kut1 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4807 0.03896 0.02559 0.0438 1.0000 0.5865 -2.750 -0.4564 0.03826 0.02442 0.0442 1.0000 0.5966 -2.500 -0.4247 0.03767 0.02316 0.0425 1.0000 0.5956 -2.250 -0.3909 0.03706 0.02183 0.0404 1.0000 0.5970 -2.000 -0.3563 0.03647 0.02074 0.0380 1.0000 0.6012 -1.750 -0.3201 0.03586 0.01978 0.0354 1.0000 0.6083 -1.500 -0.2837 0.03525 0.01894 0.0329 1.0000 0.6203 -1.250 -0.2433 0.03447 0.01825 0.0297 1.0000 0.6404 -1.000 -0.0987 0.03378 0.01867 0.0070 1.0000 0.9366 -0.750 -0.0137 0.03426 0.01829 -0.0067 1.0000 1.0000 -0.500 0.0128 0.03460 0.01819 -0.0072 1.0000 1.0000 -0.250 0.0362 0.03493 0.01824 -0.0069 1.0000 1.0000 0.000 0.0574 0.03524 0.01842 -0.0061 1.0000 1.0000 0.250 0.0774 0.03554 0.01871 -0.0051 1.0000 1.0000 0.500 0.0967 0.03584 0.01913 -0.0039 1.0000 1.0000 0.750 0.1156 0.03616 0.01965 -0.0027 1.0000 1.0000 1.000 0.1344 0.03647 0.02026 -0.0015 1.0000 1.0000 1.250 0.1529 0.03683 0.02099 -0.0003 1.0000 1.0000 1.500 0.1712 0.03723 0.02185 0.0009 1.0000 1.0000 1.750 0.1888 0.03772 0.02294 0.0021 1.0000 1.0000 2.000 0.2050 0.03833 0.02414 0.0033 1.0000 1.0000 2.250 0.2178 0.03924 0.02565 0.0045 1.0000 1.0000 2.500 0.2183 0.04115 0.02802 0.0055 1.0000 1.0000 2.750 0.4633 0.03593 0.02477 -0.0198 0.7311 1.0000 3.000 0.4547 0.03663 0.02571 -0.0136 0.6649 1.0000 3.250 0.5292 0.03641 0.02386 -0.0134 0.3670 1.0000 3.500 0.5416 0.03813 0.02515 -0.0099 0.2945 1.0000 3.750 0.5552 0.03992 0.02659 -0.0069 0.2319 1.0000 4.000 0.5752 0.04185 0.02835 -0.0044 0.2006 1.0000 4.250 0.6015 0.04399 0.03046 -0.0027 0.1861 1.0000 4.500 0.6357 0.04656 0.03340 -0.0020 0.1791 1.0000 4.750 0.6727 0.04969 0.03701 -0.0019 0.1762 1.0000 5.000 0.7066 0.05338 0.04133 -0.0020 0.1762 1.0000