XFOIL Version 6.94 Calculated polar for: eif385mod5 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2155 0.04889 0.04100 0.0466 0.9993 0.8552 -2.750 -0.1805 0.04740 0.03959 0.0368 0.9993 0.8206 -2.500 -0.1103 0.04646 0.03873 0.0194 0.9993 0.7792 -2.250 -0.0372 0.04584 0.03826 0.0038 0.9993 0.7473 -2.000 0.2379 0.03977 0.03191 -0.0475 0.6756 0.6839 -1.750 0.3235 0.04031 0.03007 -0.0584 0.5318 0.6664 -1.500 0.3664 0.04093 0.02986 -0.0622 0.5022 0.6578 -1.250 0.4127 0.04185 0.03005 -0.0669 0.4830 0.6468 -1.000 0.4569 0.04264 0.03044 -0.0710 0.4681 0.6419 -0.750 0.4981 0.04338 0.03082 -0.0742 0.4563 0.6363 -0.500 0.5394 0.04426 0.03149 -0.0774 0.4457 0.6279 -0.250 0.5800 0.04540 0.03225 -0.0803 0.4372 0.6225 0.000 0.6182 0.04635 0.03323 -0.0829 0.4320 0.6222 0.250 0.6557 0.04747 0.03444 -0.0854 0.4290 0.6214 0.500 0.6921 0.04884 0.03589 -0.0878 0.4269 0.6185 0.750 0.7264 0.05035 0.03753 -0.0899 0.4257 0.6158 1.000 0.7585 0.05202 0.03942 -0.0917 0.4253 0.6140 1.250 0.7889 0.05395 0.04160 -0.0936 0.4257 0.6133 1.500 0.8177 0.05616 0.04408 -0.0954 0.4269 0.6142 1.750 0.8437 0.05860 0.04685 -0.0971 0.4288 0.6189 2.000 0.8671 0.06126 0.04988 -0.0985 0.4312 0.6274 2.250 0.8878 0.06417 0.05318 -0.0997 0.4336 0.6409 2.500 0.9046 0.06705 0.05647 -0.1002 0.4358 0.7512 2.750 0.8914 0.07318 0.06310 -0.1015 0.4436 1.0007 3.000 0.8757 0.07992 0.07006 -0.1024 0.4523 1.0007 3.250 0.8787 0.08518 0.07530 -0.1035 0.4579 1.0007 3.750 0.7733 0.10518 0.09562 -0.1070 0.5016 0.7152 4.250 0.7078 0.12251 0.11293 -0.1151 0.5896 0.6414