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 = 0.060 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -1.900 -0.1095 0.02302 0.01463 -0.0042 1.0000 0.9778 -1.800 -0.0880 0.02285 0.01431 -0.0069 1.0000 0.9832 -1.700 -0.0663 0.02268 0.01401 -0.0096 1.0000 0.9885 -1.600 -0.0442 0.02251 0.01372 -0.0125 1.0000 0.9941 -1.500 -0.0211 0.02233 0.01344 -0.0156 1.0000 0.9995 -1.400 -0.0198 0.02214 0.01322 -0.0147 1.0000 1.0000 -1.300 -0.0214 0.02194 0.01301 -0.0133 1.0000 1.0000 -1.200 -0.0235 0.02173 0.01279 -0.0118 1.0000 1.0000 -1.100 -0.0262 0.02152 0.01257 -0.0103 1.0000 1.0000 -1.000 -0.0294 0.02130 0.01235 -0.0086 1.0000 1.0000 -0.900 -0.0332 0.02107 0.01211 -0.0069 1.0000 1.0000 -0.800 -0.0377 0.02083 0.01187 -0.0051 1.0000 1.0000 -0.700 -0.0428 0.02057 0.01161 -0.0031 1.0000 1.0000 -0.600 -0.0481 0.02032 0.01135 -0.0012 1.0000 1.0000 -0.500 -0.0530 0.02008 0.01110 0.0008 1.0000 1.0000 -0.400 -0.0564 0.01987 0.01087 0.0026 1.0000 1.0000 -0.300 -0.0580 0.01971 0.01066 0.0041 1.0000 1.0000 -0.200 -0.0569 0.01961 0.01051 0.0052 1.0000 1.0000 -0.100 -0.0536 0.01957 0.01041 0.0059 1.0000 1.0000 0.000 -0.0488 0.01956 0.01034 0.0065 1.0000 1.0000 1.000 0.0228 0.02053 0.01085 0.0085 1.0000 1.0000 2.000 0.1577 0.02349 0.01372 -0.0026 0.9756 1.0000 3.000 0.4217 0.02327 0.01407 -0.0303 0.8553 1.0000 4.000 0.6367 0.01894 0.00985 -0.0393 0.6528 1.0000 5.000 0.7412 0.02177 0.01170 -0.0360 0.4827 1.0000 6.000 0.8340 0.02564 0.01540 -0.0323 0.3875 1.0000 7.000 0.9162 0.02985 0.01983 -0.0273 0.3082 1.0000 8.100 0.9984 0.03535 0.02586 -0.0207 0.2340 1.0000 8.200 1.0075 0.03568 0.02612 -0.0203 0.2271 1.0000 8.300 1.0123 0.03640 0.02701 -0.0194 0.2210 1.0000 8.400 1.0193 0.03692 0.02758 -0.0187 0.2148 1.0000 8.500 1.0284 0.03761 0.02821 -0.0184 0.2097 1.0000 8.600 1.0308 0.03857 0.02945 -0.0172 0.2047 1.0000 8.700 1.0398 0.03909 0.02993 -0.0168 0.1995 1.0000 8.800 1.0461 0.03992 0.03083 -0.0162 0.1949 1.0000 8.900 1.0481 0.04087 0.03200 -0.0149 0.1902 1.0000 9.000 1.0585 0.04118 0.03222 -0.0147 0.1848 1.0000 9.100 1.0628 0.04210 0.03324 -0.0139 0.1806 1.0000 9.200 1.0635 0.04318 0.03458 -0.0125 0.1767 1.0000 9.300 1.0701 0.04393 0.03538 -0.0119 0.1727 1.0000 9.400 1.0835 0.04459 0.03586 -0.0122 0.1688 1.0000 9.500 1.0815 0.04605 0.03762 -0.0106 0.1668 1.0000 9.600 1.0791 0.04755 0.03941 -0.0091 0.1647 1.0000 9.700 1.0780 0.04901 0.04109 -0.0078 0.1627 1.0000 9.800 1.0789 0.05032 0.04255 -0.0067 0.1604 1.0000 10.000 1.0969 0.05187 0.04402 -0.0062 0.1534 1.0000