XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY LS(1)-0413 (GA(W)-2) AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -10.250 -0.3919 0.09506 0.09176 -0.0395 1.0000 0.0748 -10.000 -0.4004 0.09297 0.08971 -0.0377 1.0000 0.0758 -9.750 -0.4116 0.09040 0.08718 -0.0363 1.0000 0.0771 -9.500 -0.4159 0.08474 0.08154 -0.0398 0.9981 0.0795 -9.250 -0.4671 0.07030 0.06707 -0.0543 0.9940 0.0825 -9.000 -0.5188 0.06190 0.05839 -0.0634 0.9844 0.0832 -8.750 -0.5813 0.06917 0.06569 -0.0543 0.9934 0.0775 -8.250 -0.5714 0.05601 0.05179 -0.0686 0.9808 0.0863 -8.000 -0.5506 0.05346 0.04926 -0.0696 0.9767 0.0883 -7.750 -0.5270 0.05055 0.04619 -0.0723 0.9731 0.0926 -7.500 -0.5025 0.04618 0.04141 -0.0771 0.9702 0.1010 -7.250 -0.4746 0.03311 0.02616 -0.0786 0.9654 0.0505 -7.000 -0.4455 0.03028 0.02318 -0.0798 0.9622 0.0493 -6.750 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