XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY LS(1)-0421 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -16.750 -0.6408 0.10208 0.09533 -0.0804 1.0000 0.0490 -16.500 -0.7002 0.08850 0.08134 -0.0882 1.0000 0.0489 -16.250 -0.7307 0.08089 0.07345 -0.0920 1.0000 0.0492 -16.000 -0.7535 0.07510 0.06740 -0.0944 1.0000 0.0495 -15.750 -0.7731 0.07041 0.06247 -0.0956 1.0000 0.0500 -15.500 -0.7899 0.06695 0.05881 -0.0955 1.0000 0.0504 -15.250 -0.7975 0.06554 0.05742 -0.0936 1.0000 0.0508 -15.000 -0.7937 0.06350 0.05535 -0.0941 0.9959 0.0515 -14.750 -0.7830 0.06087 0.05261 -0.0964 0.9909 0.0524 -14.500 -0.7736 0.05822 0.04981 -0.0982 0.9846 0.0533 -14.250 -0.7628 0.05557 0.04695 -0.1002 0.9796 0.0544 -14.000 -0.7521 0.05306 0.04419 -0.1016 0.9737 0.0555 -13.750 -0.7341 0.05136 0.04246 -0.1029 0.9682 0.0566 -13.500 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1.0000 17.250 1.3540 0.09309 0.08546 -0.0276 0.0996 1.0000 17.500 1.3509 0.09718 0.08976 -0.0289 0.0978 1.0000 17.750 1.3498 0.10100 0.09374 -0.0304 0.0961 1.0000 18.000 1.3518 0.10437 0.09720 -0.0318 0.0945 1.0000 18.250 1.3591 0.10690 0.09974 -0.0329 0.0931 1.0000 18.500 1.3633 0.10991 0.10280 -0.0343 0.0917 1.0000 18.750 1.3434 0.11684 0.11007 -0.0375 0.0903 1.0000 19.000 1.3232 0.12410 0.11761 -0.0413 0.0888 1.0000