XFOIL Version 6.94 Calculated polar for: GOA 1 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 -0.2269 0.03692 0.02102 -0.0241 1.0000 0.4509 -2.750 -0.2069 0.03352 0.01877 -0.0204 1.0000 0.5512 -2.500 -0.1317 0.02922 0.01340 -0.0248 1.0000 1.0000 -2.250 -0.1038 0.02914 0.01186 -0.0235 1.0000 1.0000 -2.000 -0.0782 0.02908 0.01094 -0.0224 1.0000 1.0000 -1.750 -0.0532 0.02904 0.01026 -0.0214 1.0000 1.0000 -1.500 -0.0285 0.02903 0.00972 -0.0204 1.0000 1.0000 -1.250 -0.0040 0.02904 0.00930 -0.0195 1.0000 1.0000 -1.000 0.0202 0.02906 0.00899 -0.0187 1.0000 1.0000 -0.750 0.0443 0.02911 0.00880 -0.0178 1.0000 1.0000 -0.500 0.0684 0.02917 0.00862 -0.0170 1.0000 1.0000 -0.250 0.0925 0.02925 0.00856 -0.0161 1.0000 1.0000 0.000 0.1169 0.02934 0.00840 -0.0151 1.0000 1.0000 0.250 0.1410 0.02947 0.00861 -0.0143 1.0000 1.0000 0.500 0.1654 0.02961 0.00896 -0.0134 1.0000 1.0000 0.750 0.1902 0.02976 0.00949 -0.0124 1.0000 1.0000 1.000 0.2166 0.02990 0.01025 -0.0115 1.0000 1.0000 1.250 0.2471 0.02998 0.01136 -0.0110 1.0000 1.0000 1.500 0.2870 0.03002 0.01303 -0.0127 1.0000 1.0000 1.750 0.4199 0.03354 0.01609 -0.0296 0.6028 1.0000 2.000 0.4511 0.03610 0.01803 -0.0294 0.5600 1.0000 2.250 0.4833 0.03854 0.02047 -0.0303 0.5346 1.0000 2.500 0.5163 0.04102 0.02307 -0.0319 0.5185 1.0000 2.750 0.5492 0.04365 0.02593 -0.0340 0.5086 1.0000 3.000 0.5817 0.04646 0.02913 -0.0369 0.5026 1.0000 3.250 0.6115 0.04953 0.03244 -0.0389 0.4961 1.0000 3.500 0.6382 0.05277 0.03606 -0.0420 0.4904 1.0000 3.750 0.6621 0.05622 0.03981 -0.0446 0.4852 1.0000 4.000 0.6839 0.05991 0.04377 -0.0475 0.4834 1.0000 4.250 0.7018 0.06391 0.04805 -0.0514 0.4876 1.0000 4.500 0.7208 0.06816 0.05250 -0.0551 0.4957 1.0000 4.750 0.7272 0.07259 0.05709 -0.0603 0.5083 1.0000 5.000 0.7307 0.07702 0.06159 -0.0648 0.5220 1.0000