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 = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -1.900 -0.0289 0.02476 0.01452 -0.0202 1.0000 1.0000 -1.800 -0.0288 0.02457 0.01430 -0.0190 1.0000 1.0000 -1.700 -0.0290 0.02437 0.01408 -0.0178 1.0000 1.0000 -1.600 -0.0296 0.02418 0.01387 -0.0165 1.0000 1.0000 -1.500 -0.0304 0.02398 0.01366 -0.0152 1.0000 1.0000 -1.400 -0.0316 0.02378 0.01344 -0.0138 1.0000 1.0000 -1.300 -0.0332 0.02358 0.01322 -0.0124 1.0000 1.0000 -1.200 -0.0351 0.02338 0.01301 -0.0109 1.0000 1.0000 -1.100 -0.0374 0.02317 0.01279 -0.0093 1.0000 1.0000 -1.000 -0.0401 0.02296 0.01256 -0.0077 1.0000 1.0000 0.000 -0.0425 0.02161 0.01085 0.0056 1.0000 1.0000 1.000 0.0252 0.02247 0.01122 0.0085 1.0000 1.0000 2.000 0.0962 0.02466 0.01327 0.0093 1.0000 1.0000 3.000 0.3199 0.02935 0.01841 -0.0194 0.9055 1.0000 4.000 0.5942 0.02471 0.01472 -0.0387 0.7270 1.0000 5.000 0.7383 0.02521 0.01468 -0.0376 0.5490 1.0000 6.000 0.8305 0.02992 0.01936 -0.0332 0.4389 1.0000 7.000 0.9138 0.03565 0.02531 -0.0280 0.3559 1.0000 8.000 0.9784 0.04266 0.03304 -0.0207 0.2846 1.0000 9.000 1.0101 0.05267 0.04412 -0.0113 0.2306 1.0000 9.200 1.0207 0.05500 0.04655 -0.0100 0.2235 1.0000 9.300 1.0132 0.05693 0.04869 -0.0086 0.2221 1.0000 9.400 1.0060 0.05887 0.05079 -0.0074 0.2206 1.0000 9.500 0.9992 0.06077 0.05284 -0.0063 0.2192 1.0000 9.600 0.9933 0.06260 0.05479 -0.0053 0.2176 1.0000 5.000 0.7399 0.02379 0.01343 -0.0368 0.5222 1.0000