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 = 0.050 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -10.750 -0.4484 0.10255 0.09480 -0.0499 1.0000 0.0657 -10.500 -0.4580 0.09715 0.08948 -0.0522 1.0000 0.0643 -10.000 -0.5493 0.08005 0.07244 -0.0616 1.0000 0.0598 -9.750 -0.5707 0.07735 0.06981 -0.0597 1.0000 0.0596 -9.500 -0.5901 0.07437 0.06684 -0.0582 1.0000 0.0593 -9.250 -0.6075 0.07138 0.06380 -0.0564 1.0000 0.0590 -9.000 -0.6221 0.06827 0.06059 -0.0548 1.0000 0.0588 -8.750 -0.6327 0.06507 0.05723 -0.0533 1.0000 0.0586 -8.500 -0.6389 0.06177 0.05371 -0.0521 1.0000 0.0584 -8.250 -0.6405 0.05845 0.05012 -0.0512 1.0000 0.0582 -8.000 -0.6376 0.05519 0.04653 -0.0506 1.0000 0.0583 -7.750 -0.6304 0.05204 0.04298 -0.0502 1.0000 0.0588 -7.500 -0.6191 0.04909 0.03955 -0.0500 1.0000 0.0596 -7.250 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