XFOIL Version 6.94 Calculated polar for: GOA 2 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.012 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1674 0.02996 0.01561 -0.0260 1.0000 1.0000 -2.750 -0.1391 0.02985 0.01403 -0.0253 1.0000 1.0000 -2.500 -0.1139 0.02976 0.01300 -0.0244 1.0000 1.0000 -2.250 -0.0893 0.02970 0.01218 -0.0234 1.0000 1.0000 -2.000 -0.0651 0.02966 0.01143 -0.0226 1.0000 1.0000 -1.750 -0.0411 0.02964 0.01087 -0.0217 1.0000 1.0000 -1.500 -0.0172 0.02965 0.01041 -0.0208 1.0000 1.0000 -1.250 0.0066 0.02968 0.01007 -0.0199 1.0000 1.0000 -1.000 0.0304 0.02972 0.00983 -0.0191 1.0000 1.0000 -0.750 0.0541 0.02979 0.00971 -0.0182 1.0000 1.0000 -0.500 0.0779 0.02988 0.00963 -0.0173 1.0000 1.0000 -0.250 0.1017 0.02999 0.00972 -0.0164 1.0000 1.0000 0.000 0.1258 0.03012 0.00995 -0.0155 1.0000 1.0000 0.250 0.1503 0.03026 0.01032 -0.0145 1.0000 1.0000 0.500 0.1757 0.03040 0.01087 -0.0136 1.0000 1.0000 0.750 0.2034 0.03051 0.01167 -0.0128 1.0000 1.0000 1.000 0.2369 0.03053 0.01275 -0.0126 1.0000 1.0000 1.250 0.2775 0.03064 0.01437 -0.0151 1.0000 1.0000 1.500 0.4121 0.03443 0.01670 -0.0307 0.5787 1.0000 1.750 0.4438 0.03684 0.01863 -0.0309 0.5439 1.0000 2.000 0.4766 0.03914 0.02086 -0.0319 0.5224 1.0000 2.250 0.5096 0.04153 0.02335 -0.0333 0.5087 1.0000 2.500 0.5426 0.04399 0.02609 -0.0354 0.5000 1.0000 2.750 0.5751 0.04672 0.02905 -0.0375 0.4952 1.0000 3.000 0.6058 0.04958 0.03239 -0.0406 0.4927 1.0000 3.250 0.6332 0.05270 0.03589 -0.0434 0.4896 1.0000 3.500 0.6574 0.05603 0.03953 -0.0461 0.4863 1.0000 3.750 0.6788 0.05957 0.04338 -0.0486 0.4840 1.0000 4.000 0.6967 0.06334 0.04741 -0.0517 0.4849 1.0000 4.250 0.7114 0.06737 0.05164 -0.0550 0.4896 1.0000 4.500 0.7248 0.07160 0.05603 -0.0587 0.4976 1.0000 4.750 0.7312 0.07605 0.06057 -0.0629 0.5102 1.0000 5.000 0.7299 0.08047 0.06502 -0.0670 0.5248 1.0000