XFOIL Version 6.94 Calculated polar for: Eif385mod3 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 -0.0973 0.04069 0.03163 -0.0102 0.9993 0.5183 -2.750 0.0816 0.03547 0.02600 -0.0368 0.8797 0.4905 -2.500 0.1743 0.03431 0.02225 -0.0440 0.4660 0.4847 -2.250 0.2040 0.03507 0.02212 -0.0441 0.4329 0.4805 -2.000 0.2356 0.03546 0.02208 -0.0444 0.4149 0.4791 -1.750 0.2683 0.03580 0.02212 -0.0449 0.4018 0.4809 -1.500 0.3009 0.03613 0.02226 -0.0453 0.3903 0.4833 -1.250 0.3337 0.03649 0.02252 -0.0459 0.3812 0.4889 -1.000 0.3668 0.03674 0.02288 -0.0464 0.3736 0.4961 -0.750 0.4002 0.03712 0.02337 -0.0471 0.3678 0.5045 -0.250 0.4758 0.03696 0.02461 -0.0496 0.3612 1.0007 0.000 0.5113 0.03816 0.02562 -0.0507 0.3591 1.0007 0.250 0.5463 0.03947 0.02681 -0.0519 0.3577 1.0007 0.500 0.5796 0.04087 0.02815 -0.0529 0.3570 1.0007 0.750 0.6114 0.04231 0.02959 -0.0536 0.3568 1.0007 1.000 0.6420 0.04378 0.03110 -0.0543 0.3571 1.0007 1.250 0.6715 0.04528 0.03269 -0.0549 0.3579 1.0007 1.500 0.7000 0.04689 0.03445 -0.0555 0.3592 1.0007 1.750 0.7273 0.04871 0.03641 -0.0560 0.3602 1.0007 2.000 0.7527 0.05073 0.03860 -0.0565 0.3611 1.0007 2.250 0.7763 0.05296 0.04103 -0.0570 0.3621 1.0007 2.500 0.7974 0.05548 0.04378 -0.0574 0.3636 1.0007 2.750 0.8150 0.05846 0.04703 -0.0579 0.3668 1.0007 3.000 0.8290 0.06197 0.05080 -0.0583 0.3720 1.0007 3.250 0.8419 0.06576 0.05474 -0.0588 0.3779 1.0007 3.500 0.8603 0.06939 0.05838 -0.0593 0.3826 1.0007 3.750 0.8143 0.07811 0.06779 -0.0597 0.4026 1.0007 4.000 0.8288 0.08259 0.07227 -0.0609 0.4120 1.0007