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 = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.1181 0.03452 0.02202 -0.0477 0.4566 0.3717 -2.750 0.1417 0.03503 0.02166 -0.0466 0.4044 0.3766 -2.500 0.1697 0.03524 0.02142 -0.0463 0.3829 0.3815 -2.250 0.1991 0.03551 0.02129 -0.0461 0.3691 0.3864 -2.000 0.2313 0.03557 0.02115 -0.0463 0.3607 0.3950 -1.750 0.2630 0.03545 0.02107 -0.0465 0.3539 0.4064 -1.500 0.2961 0.03560 0.02114 -0.0469 0.3484 0.4190 -1.250 0.3292 0.03571 0.02134 -0.0474 0.3441 0.4378 -1.000 0.3627 0.03579 0.02168 -0.0480 0.3410 0.4659 -0.750 0.3916 0.03523 0.02209 -0.0475 0.3389 0.5572 -0.500 0.4417 0.03504 0.02258 -0.0505 0.3367 1.0007 -0.250 0.4766 0.03600 0.02334 -0.0514 0.3347 1.0007 0.000 0.5097 0.03703 0.02422 -0.0520 0.3325 1.0007 0.250 0.5414 0.03814 0.02522 -0.0524 0.3301 1.0007 0.500 0.5720 0.03934 0.02633 -0.0527 0.3278 1.0007 0.750 0.6019 0.04063 0.02757 -0.0531 0.3262 1.0007 1.000 0.6312 0.04203 0.02897 -0.0534 0.3261 1.0007 1.250 0.6600 0.04355 0.03057 -0.0538 0.3272 1.0007 1.500 0.6880 0.04522 0.03232 -0.0542 0.3289 1.0007 1.750 0.7152 0.04706 0.03426 -0.0546 0.3310 1.0007 2.000 0.7415 0.04907 0.03638 -0.0550 0.3335 1.0007 2.250 0.7668 0.05129 0.03868 -0.0554 0.3362 1.0007 2.500 0.7916 0.05373 0.04116 -0.0558 0.3390 1.0007 2.750 0.8163 0.05650 0.04391 -0.0563 0.3416 1.0007 3.000 0.8290 0.05907 0.04724 -0.0568 0.3513 1.0007 3.250 0.8412 0.06296 0.05142 -0.0574 0.3598 1.0007 3.500 0.8589 0.06657 0.05509 -0.0578 0.3655 1.0007 3.750 0.8324 0.07388 0.06306 -0.0586 0.3848 1.0007 4.000 0.8374 0.07896 0.06823 -0.0596 0.3965 1.0007 5.000 0.3055 0.11704 0.10842 -0.0463 0.6972 1.0007