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 = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.1138 0.03539 0.02202 -0.0473 0.3293 0.3239 -2.750 0.1424 0.03520 0.02157 -0.0470 0.3212 0.3248 -2.500 0.1718 0.03512 0.02125 -0.0468 0.3151 0.3291 -2.250 0.2043 0.03494 0.02095 -0.0471 0.3106 0.3349 -2.000 0.2382 0.03490 0.02075 -0.0475 0.3066 0.3398 -1.750 0.2738 0.03497 0.02066 -0.0483 0.3032 0.3445 -1.500 0.3089 0.03478 0.02056 -0.0491 0.3004 0.3541 -1.250 0.3438 0.03492 0.02069 -0.0498 0.2979 0.3647 -1.000 0.3771 0.03514 0.02094 -0.0504 0.2953 0.3763 -0.750 0.4094 0.03560 0.02143 -0.0509 0.2924 0.3928 -0.500 0.4408 0.03624 0.02224 -0.0514 0.2898 0.4130 -0.250 0.4688 0.03610 0.02284 -0.0513 0.2882 0.4798 0.250 0.5388 0.03675 0.02453 -0.0527 0.2870 1.0007 0.500 0.5688 0.03799 0.02568 -0.0528 0.2873 1.0007 0.750 0.5983 0.03934 0.02697 -0.0530 0.2878 1.0007 1.000 0.6274 0.04083 0.02841 -0.0533 0.2885 1.0007 1.250 0.6559 0.04250 0.03004 -0.0535 0.2895 1.0007 1.500 0.6844 0.04392 0.03150 -0.0537 0.2908 1.0007 1.750 0.7127 0.04485 0.03263 -0.0538 0.2937 1.0007 2.000 0.7390 0.04639 0.03453 -0.0540 0.2982 1.0007 2.250 0.7644 0.04845 0.03683 -0.0543 0.3030 1.0007 2.500 0.7889 0.05077 0.03929 -0.0546 0.3075 1.0007 2.750 0.8129 0.05331 0.04189 -0.0549 0.3112 1.0007 3.000 0.8366 0.05617 0.04475 -0.0552 0.3140 1.0007 3.250 0.8528 0.05831 0.04753 -0.0554 0.3231 1.0007 3.500 0.8642 0.06241 0.05199 -0.0558 0.3334 1.0007 3.750 0.8851 0.06620 0.05574 -0.0564 0.3398 1.0007 4.000 0.8461 0.07732 0.06795 -0.0605 0.3966 1.0007 4.500 0.4966 0.11709 0.10861 -0.0740 0.6749 1.0007 4.750 0.5137 0.11987 0.11129 -0.0740 0.6577 1.0007 5.000 0.5307 0.12265 0.11396 -0.0739 0.6401 1.0007