XFOIL Version 6.94 Calculated polar for: Eif385mod2 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0882 0.04272 0.03139 -0.0173 0.9993 0.4294 -2.750 -0.0569 0.04211 0.03089 -0.0177 0.9993 0.4352 -2.500 -0.0237 0.04161 0.03053 -0.0185 0.9993 0.4407 -2.250 0.2007 0.03401 0.02246 -0.0472 0.5744 0.4793 -2.000 0.2318 0.03473 0.02188 -0.0463 0.4879 0.4967 -1.750 0.2588 0.03505 0.02190 -0.0455 0.4594 0.5215 -1.500 0.2860 0.03508 0.02204 -0.0448 0.4407 0.5620 -1.250 0.3294 0.03407 0.02210 -0.0462 0.4239 1.0007 -1.000 0.3669 0.03519 0.02251 -0.0477 0.4144 1.0007 -0.750 0.4030 0.03616 0.02308 -0.0487 0.4078 1.0007 -0.500 0.4381 0.03721 0.02376 -0.0496 0.4025 1.0007 -0.250 0.4726 0.03834 0.02458 -0.0504 0.3983 1.0007 0.000 0.5068 0.03959 0.02556 -0.0513 0.3952 1.0007 0.250 0.5402 0.04088 0.02670 -0.0522 0.3931 1.0007 0.500 0.5723 0.04217 0.02799 -0.0530 0.3919 1.0007 0.750 0.6037 0.04359 0.02944 -0.0538 0.3910 1.0007 1.000 0.6341 0.04515 0.03106 -0.0547 0.3898 1.0007 1.250 0.6628 0.04684 0.03283 -0.0554 0.3886 1.0007 1.500 0.6900 0.04870 0.03479 -0.0560 0.3874 1.0007 1.750 0.7158 0.05073 0.03692 -0.0566 0.3865 1.0007 2.000 0.7404 0.05298 0.03928 -0.0572 0.3866 1.0007 2.250 0.7639 0.05548 0.04190 -0.0579 0.3882 1.0007 2.500 0.7864 0.05822 0.04473 -0.0586 0.3908 1.0007 2.750 0.8087 0.06115 0.04769 -0.0593 0.3933 1.0007 3.000 0.8086 0.06562 0.05277 -0.0601 0.4014 1.0007 3.250 0.8056 0.07084 0.05826 -0.0606 0.4100 1.0007 3.500 0.8178 0.07501 0.06244 -0.0612 0.4160 1.0007 3.750 0.7566 0.08507 0.07292 -0.0615 0.4366 1.0007 4.250 0.5911 0.10926 0.09743 -0.0677 0.5377 1.0007