XFOIL Version 6.94 Calculated polar for: BOEING BACXXX AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -1.700 -0.0248 0.02373 0.01398 -0.0181 1.0000 1.0000 -1.600 -0.0253 0.02353 0.01377 -0.0168 1.0000 1.0000 -1.500 -0.0261 0.02334 0.01356 -0.0155 1.0000 1.0000 -1.400 -0.0274 0.02314 0.01334 -0.0142 1.0000 1.0000 -1.300 -0.0290 0.02294 0.01313 -0.0127 1.0000 1.0000 -1.200 -0.0310 0.02273 0.01291 -0.0112 1.0000 1.0000 -1.100 -0.0335 0.02252 0.01269 -0.0097 1.0000 1.0000 -1.000 -0.0365 0.02230 0.01246 -0.0080 1.0000 1.0000 0.000 -0.0448 0.02082 0.01065 0.0059 1.0000 1.0000 1.000 0.0243 0.02172 0.01107 0.0085 1.0000 1.0000 2.000 0.0957 0.02398 0.01321 0.0091 1.0000 1.0000 3.000 0.3647 0.02720 0.01696 -0.0247 0.8825 1.0000 4.000 0.6160 0.02197 0.01239 -0.0391 0.7021 1.0000 5.000 0.7399 0.02379 0.01343 -0.0368 0.5223 1.0000 6.000 0.8331 0.02818 0.01772 -0.0328 0.4179 1.0000 7.000 0.9137 0.03340 0.02325 -0.0274 0.3364 1.0000 8.000 0.9855 0.03953 0.02990 -0.0210 0.2665 1.0000 9.000 1.0433 0.04762 0.03871 -0.0135 0.2107 1.0000 9.300 1.0402 0.05151 0.04320 -0.0096 0.2007 1.0000 9.400 1.0509 0.05204 0.04366 -0.0094 0.1954 1.0000 9.500 1.0443 0.05376 0.04564 -0.0078 0.1935 1.0000 9.600 1.0393 0.05539 0.04746 -0.0064 0.1913 1.0000 9.700 1.0376 0.05679 0.04899 -0.0053 0.1887 1.0000 9.900 1.0567 0.05866 0.05082 -0.0048 0.1820 1.0000 10.000 1.0478 0.06068 0.05304 -0.0034 0.1815 1.0000