XFOIL Version 6.94 Calculated polar for: AIRFOIL PROFILE12A 9.00% 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.500 0.2726 0.00742 0.00293 -0.0432 0.8518 0.9475 1.000 0.3401 0.00744 0.00280 -0.0449 0.8108 0.9640 1.500 0.4131 0.00754 0.00272 -0.0483 0.7638 0.9726 2.000 0.4816 0.00772 0.00270 -0.0509 0.7122 0.9818 2.500 0.5520 0.00797 0.00277 -0.0542 0.6577 0.9903 3.000 0.6179 0.00826 0.00288 -0.0567 0.5995 1.0000 3.500 0.6476 0.00859 0.00307 -0.0520 0.5482 1.0000 4.000 0.6860 0.00903 0.00334 -0.0488 0.4918 1.0000 4.500 0.7317 0.00963 0.00373 -0.0471 0.4261 1.0000 5.000 0.7768 0.01045 0.00424 -0.0454 0.3397 1.0000 5.500 0.8223 0.01141 0.00488 -0.0439 0.2573 1.0000 6.000 0.8671 0.01256 0.00570 -0.0424 0.1755 1.0000 6.500 0.9100 0.01401 0.00682 -0.0406 0.0948 1.0000 7.000 0.9466 0.01629 0.00879 -0.0377 0.0356 1.0000 8.500 1.0569 0.02303 0.01614 -0.0288 0.0233 1.0000 9.000 1.0885 0.02893 0.02247 -0.0257 0.0214 1.0000 9.500 1.1233 0.03098 0.02494 -0.0228 0.0194 1.0000 10.000 1.1470 0.03535 0.02983 -0.0190 0.0176 1.0000 10.500 1.1555 0.04042 0.03541 -0.0139 0.0165 1.0000 11.000 1.1418 0.04576 0.04128 -0.0071 0.0159 1.0000 11.500 1.1215 0.05122 0.04709 -0.0027 0.0152 1.0000 12.000 1.0780 0.06111 0.05756 -0.0018 0.0159 1.0000 12.500 1.0064 0.07835 0.07537 -0.0105 0.0174 1.0000 13.000 0.9546 0.09769 0.09501 -0.0243 0.0181 1.0000 13.500 0.8891 0.12730 0.12472 -0.0425 0.0201 1.0000 15.000 0.6732 0.16252 0.16026 -0.0626 0.0349 1.0000 15.500 0.6917 0.16812 0.16593 -0.0636 0.0323 1.0000 16.000 0.6917 0.17974 0.17753 -0.0685 0.0313 1.0000 16.500 0.6899 0.18769 0.18545 -0.0753 0.0293 1.0000 17.000 0.6999 0.19448 0.19228 -0.0790 0.0238 1.0000 17.500 0.7149 0.20101 0.19885 -0.0813 0.0224 1.0000