XFOIL Version 6.94 Calculated polar for: GOA 3 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1490 0.02466 0.01324 -0.0320 0.9999 0.1104 -2.750 -0.1221 0.02334 0.01195 -0.0309 0.9999 0.1247 -2.500 -0.0944 0.02177 0.01068 -0.0301 0.9999 0.1675 -2.250 -0.0750 0.01838 0.01001 -0.0281 0.9999 0.6206 -2.000 -0.0496 0.01768 0.00954 -0.0257 0.9999 1.0001 -1.750 -0.0243 0.01777 0.00945 -0.0245 0.9999 1.0001 -1.500 0.0013 0.01787 0.00950 -0.0233 0.9999 1.0001 -1.000 0.1313 0.02073 0.00898 -0.0324 0.2473 1.0001 -0.750 0.1549 0.02163 0.00936 -0.0312 0.2281 1.0001 -0.500 0.1794 0.02255 0.00982 -0.0301 0.2103 1.0001 -0.250 0.2062 0.02332 0.01034 -0.0293 0.1988 1.0001 0.000 0.2350 0.02432 0.01104 -0.0287 0.1942 1.0001 0.250 0.2664 0.02533 0.01189 -0.0285 0.1918 1.0001 0.500 0.2994 0.02635 0.01289 -0.0285 0.1909 1.0001 0.750 0.3329 0.02751 0.01405 -0.0287 0.1911 1.0001 1.000 0.3649 0.02886 0.01541 -0.0287 0.1922 1.0001 1.250 0.3956 0.03014 0.01685 -0.0285 0.1947 1.0001 1.500 0.4265 0.03148 0.01854 -0.0283 0.1999 1.0001 1.750 0.4565 0.03324 0.02051 -0.0283 0.2054 1.0001 2.000 0.4859 0.03518 0.02259 -0.0283 0.2110 1.0001 2.250 0.5172 0.03710 0.02500 -0.0285 0.2228 1.0001 2.500 0.5486 0.03927 0.02766 -0.0290 0.2375 1.0001 2.750 0.5798 0.04190 0.03071 -0.0299 0.2565 1.0001 3.000 0.6147 0.04489 0.03432 -0.0321 0.2904 1.0001 3.250 0.6648 0.04863 0.03920 -0.0404 0.3778 1.0001 3.750 0.5985 0.06562 0.05740 -0.0880 0.7220 1.0001 4.000 0.4985 0.06787 0.05946 -0.0829 0.8702 1.0001 4.500 0.3974 0.06661 0.05793 -0.0620 0.9999 1.0001 4.750 0.4091 0.06957 0.06083 -0.0633 0.9999 1.0001 5.000 0.4207 0.07258 0.06379 -0.0646 0.9999 1.0001