XFOIL Version 6.94 Calculated polar for: GOA 1b 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 -0.2183 0.03029 0.01892 -0.0115 1.0000 0.9614 -2.750 -0.1523 0.02932 0.01373 -0.0243 1.0000 1.0000 -2.500 -0.1257 0.02921 0.01242 -0.0234 1.0000 1.0000 -2.250 -0.1004 0.02912 0.01144 -0.0225 1.0000 1.0000 -2.000 -0.0756 0.02906 0.01057 -0.0216 1.0000 1.0000 -1.750 -0.0512 0.02902 0.00991 -0.0208 1.0000 1.0000 -1.500 -0.0269 0.02900 0.00936 -0.0200 1.0000 1.0000 -1.250 -0.0027 0.02900 0.00892 -0.0192 1.0000 1.0000 -1.000 0.0214 0.02903 0.00860 -0.0184 1.0000 1.0000 -0.750 0.0453 0.02907 0.00838 -0.0176 1.0000 1.0000 -0.500 0.0693 0.02914 0.00820 -0.0167 1.0000 1.0000 -0.250 0.0931 0.02923 0.00819 -0.0159 1.0000 1.0000 0.000 0.1170 0.02934 0.00831 -0.0151 1.0000 1.0000 0.250 0.1410 0.02946 0.00854 -0.0142 1.0000 1.0000 0.500 0.1653 0.02960 0.00892 -0.0133 1.0000 1.0000 0.750 0.1901 0.02975 0.00943 -0.0124 1.0000 1.0000 1.000 0.2163 0.02990 0.01020 -0.0115 1.0000 1.0000 1.250 0.2468 0.02998 0.01132 -0.0110 1.0000 1.0000 1.500 0.2866 0.03002 0.01298 -0.0127 1.0000 1.0000 1.750 0.4197 0.03352 0.01604 -0.0296 0.6032 1.0000 2.000 0.4509 0.03608 0.01797 -0.0294 0.5602 1.0000 2.250 0.4833 0.03851 0.02034 -0.0303 0.5347 1.0000 2.500 0.5163 0.04099 0.02296 -0.0319 0.5185 1.0000 2.750 0.5493 0.04363 0.02589 -0.0339 0.5086 1.0000 3.000 0.5818 0.04643 0.02912 -0.0368 0.5027 1.0000 3.250 0.6117 0.04951 0.03247 -0.0389 0.4961 1.0000 3.500 0.6384 0.05274 0.03616 -0.0418 0.4904 1.0000 3.750 0.6624 0.05619 0.03993 -0.0445 0.4852 1.0000 4.000 0.6841 0.05988 0.04391 -0.0473 0.4833 1.0000 4.250 0.7021 0.06387 0.04820 -0.0512 0.4875 1.0000 4.500 0.7213 0.06812 0.05264 -0.0548 0.4955 1.0000 4.750 0.7274 0.07254 0.05725 -0.0601 0.5081 1.0000 5.000 0.7310 0.07697 0.06176 -0.0646 0.5217 1.0000