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 = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0974 0.04414 0.03218 -0.0150 0.9993 0.4787 -2.750 -0.0663 0.04351 0.03163 -0.0157 0.9993 0.4841 -2.500 -0.0339 0.04299 0.03125 -0.0166 0.9993 0.4916 -2.250 -0.0016 0.04247 0.03105 -0.0173 0.9993 0.5032 -2.000 0.0303 0.04209 0.03115 -0.0180 0.9993 0.5176 -1.750 0.2509 0.03355 0.02336 -0.0447 0.5905 0.6448 -1.500 0.3067 0.03363 0.02268 -0.0479 0.5232 1.0007 -1.250 0.3429 0.03494 0.02294 -0.0491 0.4968 1.0007 -1.000 0.3771 0.03619 0.02339 -0.0498 0.4782 1.0007 -0.750 0.4103 0.03744 0.02398 -0.0503 0.4638 1.0007 -0.500 0.4434 0.03859 0.02479 -0.0510 0.4529 1.0007 -0.250 0.4767 0.03985 0.02568 -0.0517 0.4453 1.0007 0.000 0.5101 0.04121 0.02669 -0.0525 0.4403 1.0007 0.250 0.5423 0.04253 0.02799 -0.0535 0.4367 1.0007 0.500 0.5736 0.04400 0.02944 -0.0544 0.4340 1.0007 0.750 0.6039 0.04561 0.03106 -0.0553 0.4320 1.0007 1.000 0.6329 0.04738 0.03289 -0.0562 0.4309 1.0007 1.250 0.6603 0.04935 0.03494 -0.0571 0.4306 1.0007 1.500 0.6868 0.05156 0.03727 -0.0582 0.4305 1.0007 1.750 0.7111 0.05401 0.03984 -0.0592 0.4305 1.0007 2.000 0.7326 0.05673 0.04269 -0.0601 0.4305 1.0007 2.250 0.7510 0.05978 0.04587 -0.0609 0.4310 1.0007 2.500 0.7646 0.06328 0.04953 -0.0616 0.4324 1.0007 2.750 0.7745 0.06719 0.05356 -0.0621 0.4350 1.0007 3.000 0.7843 0.07124 0.05768 -0.0627 0.4384 1.0007 3.250 0.7465 0.07906 0.06588 -0.0629 0.4503 1.0007 3.500 0.7358 0.08504 0.07189 -0.0635 0.4601 1.0007 3.750 0.6751 0.09450 0.08147 -0.0637 0.4808 1.0007 4.000 0.6434 0.10219 0.08915 -0.0654 0.5031 1.0007 4.250 0.5976 0.11018 0.09718 -0.0676 0.5410 1.0007 4.500 0.5536 0.11864 0.10572 -0.0719 0.6179 1.0007 5.000 0.3814 0.12542 0.11277 -0.0633 0.9211 1.0007