XFOIL Version 6.94 Calculated polar for: pp14 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.008 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1897 0.06297 0.04687 0.0157 0.9996 0.8940 -2.750 -0.1925 0.06208 0.04614 0.0165 0.9996 0.8781 -2.500 -0.1780 0.06121 0.04536 0.0131 0.9996 0.8605 -2.250 -0.1435 0.06042 0.04457 0.0052 0.9996 0.8422 -2.000 -0.0944 0.05956 0.04364 -0.0054 0.9996 0.8257 -1.750 -0.0351 0.05893 0.04292 -0.0179 0.9996 0.8107 -1.500 0.0335 0.05866 0.04252 -0.0323 0.9996 0.7960 -1.250 0.0993 0.05834 0.04213 -0.0452 0.9996 0.7863 -1.000 0.1624 0.05811 0.04187 -0.0568 0.9996 0.7840 -0.750 0.2208 0.05792 0.04169 -0.0669 0.9996 0.7865 -0.500 0.2777 0.05791 0.04173 -0.0764 0.9996 0.7922 -0.250 0.3298 0.05790 0.04184 -0.0844 0.9996 0.8052 0.000 0.3754 0.05780 0.04197 -0.0907 0.9996 0.8286 0.250 0.4142 0.05764 0.04218 -0.0954 0.9996 0.8659 0.500 0.4367 0.05763 0.04258 -0.0975 0.9996 1.0004 0.750 0.5070 0.05990 0.04469 -0.1099 0.9996 1.0004 1.000 0.5699 0.06221 0.04680 -0.1203 0.9996 1.0004 1.250 0.6247 0.06449 0.04883 -0.1285 0.9996 1.0004 1.500 0.6717 0.06667 0.05068 -0.1345 0.9996 1.0004 1.750 0.7116 0.06872 0.05233 -0.1384 0.9996 1.0004 2.000 0.7452 0.07062 0.05378 -0.1406 0.9996 1.0004 2.250 0.7730 0.07239 0.05511 -0.1414 0.9996 1.0004 2.500 0.7966 0.07408 0.05658 -0.1413 0.9996 1.0004 2.750 0.8181 0.07578 0.05828 -0.1409 0.9996 1.0004 3.000 0.8386 0.07753 0.06025 -0.1404 0.9996 1.0004 3.250 0.8586 0.07938 0.06252 -0.1401 0.9996 1.0004 3.500 0.8777 0.08149 0.06527 -0.1399 0.9996 1.0004 3.750 0.8933 0.08444 0.06907 -0.1405 0.9996 1.0004 4.000 0.8957 0.08995 0.07549 -0.1433 0.9996 1.0004 4.250 0.8794 0.09882 0.08473 -0.1488 0.9996 1.0004 4.500 0.8653 0.10742 0.09337 -0.1540 0.9996 1.0004 4.750 0.9848 0.12069 0.10666 -0.1807 0.7835 1.0004 5.000 1.0081 0.12578 0.11160 -0.1831 0.7408 1.0004