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 = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1442 0.04076 0.03326 0.0083 0.9993 0.6591 -2.750 -0.1082 0.03984 0.03250 0.0057 0.9993 0.6674 -2.500 0.1258 0.03248 0.02355 -0.0276 0.5033 0.6895 -2.250 0.1581 0.03317 0.02310 -0.0292 0.4432 0.6888 -2.000 0.1941 0.03349 0.02286 -0.0314 0.4211 0.6894 -1.750 0.2338 0.03392 0.02278 -0.0344 0.4068 0.6813 -1.500 0.2762 0.03435 0.02285 -0.0378 0.3947 0.6652 -1.250 0.3155 0.03496 0.02309 -0.0402 0.3845 0.6574 -1.000 0.3541 0.03538 0.02345 -0.0422 0.3764 0.6546 -0.750 0.3912 0.03586 0.02395 -0.0438 0.3693 0.6572 -0.500 0.4276 0.03629 0.02456 -0.0450 0.3644 0.6960 -0.250 0.4766 0.03677 0.02522 -0.0488 0.3609 1.0007 0.000 0.5139 0.03797 0.02623 -0.0505 0.3588 1.0007 0.250 0.5505 0.03927 0.02741 -0.0522 0.3573 1.0007 0.500 0.5851 0.04067 0.02875 -0.0535 0.3565 1.0007 0.750 0.6181 0.04212 0.03019 -0.0545 0.3562 1.0007 1.000 0.6496 0.04360 0.03172 -0.0554 0.3565 1.0007 1.250 0.6799 0.04512 0.03335 -0.0562 0.3574 1.0007 1.500 0.7090 0.04675 0.03515 -0.0570 0.3590 1.0007 1.750 0.7365 0.04860 0.03722 -0.0577 0.3610 1.0007 2.000 0.7620 0.05072 0.03956 -0.0584 0.3629 1.0007 2.250 0.7856 0.05308 0.04214 -0.0589 0.3646 1.0007 2.500 0.8068 0.05571 0.04499 -0.0594 0.3665 1.0007 2.750 0.8253 0.05866 0.04815 -0.0599 0.3689 1.0007 3.000 0.8405 0.06200 0.05170 -0.0602 0.3725 1.0007 3.250 0.8546 0.06566 0.05550 -0.0606 0.3773 1.0007 3.500 0.8739 0.06925 0.05909 -0.0611 0.3818 1.0007 3.750 0.8282 0.07792 0.06848 -0.0616 0.4018 1.0007 4.000 0.8438 0.08233 0.07290 -0.0628 0.4109 1.0007 5.000 0.4620 0.12584 0.11721 -0.0740 0.8113 1.0007