XFOIL Version 6.94 Calculated polar for: Eif385mod4 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 -0.2042 0.04351 0.03505 0.0301 0.9993 0.7718 -2.750 -0.1554 0.04235 0.03387 0.0213 0.9993 0.7550 -2.500 -0.1009 0.04160 0.03304 0.0116 0.9993 0.7306 -2.250 -0.0515 0.04109 0.03256 0.0046 0.9993 0.7144 -2.000 -0.0063 0.04075 0.03242 -0.0008 0.9993 0.7045 -1.500 0.2855 0.03478 0.02433 -0.0397 0.5266 0.6905 -1.250 0.3217 0.03538 0.02452 -0.0410 0.5002 0.7096 -1.000 0.3708 0.03567 0.02484 -0.0446 0.4808 0.8586 -0.750 0.4106 0.03665 0.02542 -0.0474 0.4687 1.0007 -0.500 0.4483 0.03799 0.02621 -0.0493 0.4589 1.0007 -0.250 0.4845 0.03924 0.02726 -0.0512 0.4505 1.0007 0.000 0.5192 0.04059 0.02839 -0.0526 0.4424 1.0007 0.250 0.5534 0.04210 0.02957 -0.0538 0.4353 1.0007 0.500 0.5858 0.04358 0.03106 -0.0551 0.4305 1.0007 0.750 0.6172 0.04520 0.03273 -0.0563 0.4274 1.0007 1.000 0.6476 0.04698 0.03460 -0.0575 0.4262 1.0007 1.250 0.6763 0.04896 0.03669 -0.0586 0.4258 1.0007 1.500 0.7040 0.05117 0.03905 -0.0598 0.4260 1.0007 1.750 0.7300 0.05364 0.04166 -0.0610 0.4271 1.0007 2.000 0.7538 0.05638 0.04454 -0.0622 0.4287 1.0007 2.250 0.7755 0.05938 0.04767 -0.0633 0.4310 1.0007 2.500 0.7960 0.06260 0.05097 -0.0642 0.4335 1.0007 2.750 0.7927 0.06766 0.05649 -0.0652 0.4404 1.0007 3.000 0.7775 0.07393 0.06304 -0.0658 0.4498 1.0007 3.250 0.7824 0.07868 0.06783 -0.0665 0.4560 1.0007 3.500 0.7157 0.08871 0.07811 -0.0666 0.4756 1.0007 4.000 0.6387 0.10442 0.09386 -0.0690 0.5213 1.0007