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 = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -1.800 -0.0336 0.02535 0.01444 -0.0187 1.0000 1.0000 -1.700 -0.0338 0.02515 0.01422 -0.0174 1.0000 1.0000 -1.600 -0.0343 0.02496 0.01401 -0.0162 1.0000 1.0000 -1.500 -0.0351 0.02476 0.01379 -0.0148 1.0000 1.0000 -1.400 -0.0362 0.02457 0.01357 -0.0134 1.0000 1.0000 -1.300 -0.0376 0.02437 0.01336 -0.0120 1.0000 1.0000 -1.200 -0.0393 0.02417 0.01314 -0.0105 1.0000 1.0000 -1.100 -0.0413 0.02396 0.01292 -0.0090 1.0000 1.0000 -1.000 -0.0437 0.02376 0.01270 -0.0074 1.0000 1.0000 -1.000 -0.0437 0.02376 0.01270 -0.0074 1.0000 1.0000 0.000 -0.0400 0.02256 0.01108 0.0052 1.0000 1.0000 1.000 0.0263 0.02338 0.01141 0.0085 1.0000 1.0000 2.000 0.0968 0.02548 0.01336 0.0095 1.0000 1.0000 3.000 0.2575 0.03109 0.01932 -0.0105 0.9414 1.0000 4.000 0.5580 0.02884 0.01819 -0.0378 0.7534 1.0000 5.000 0.7287 0.02739 0.01669 -0.0384 0.5824 1.0000 6.000 0.8253 0.03213 0.02141 -0.0338 0.4655 1.0000 7.000 0.9021 0.03882 0.02858 -0.0279 0.3818 1.0000 8.000 0.9610 0.04698 0.03741 -0.0202 0.3099 1.0000 9.000 0.9695 0.05952 0.05097 -0.0107 0.2612 1.0000 9.200 0.9767 0.06163 0.05316 -0.0091 0.2511 1.0000 9.300 0.9667 0.06376 0.05542 -0.0081 0.2500 1.0000 9.400 0.9564 0.06597 0.05772 -0.0072 0.2492 1.0000 9.500 0.9451 0.06828 0.06010 -0.0065 0.2487 1.0000 9.600 0.9327 0.07073 0.06261 -0.0059 0.2488 1.0000 9.700 0.9187 0.07335 0.06526 -0.0056 0.2493 1.0000 9.800 0.9039 0.07596 0.06787 -0.0052 0.2502 1.0000 9.900 0.8899 0.07870 0.07061 -0.0052 0.2510 1.0000 10.000 0.8768 0.08160 0.07349 -0.0056 0.2518 1.0000