XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY LS(1)-0417 (GA(W)-1) 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 ------ -------- --------- --------- -------- -------- -------- -12.500 -0.3985 0.12006 0.11630 -0.0538 1.0000 0.0769 -12.250 -0.4214 0.11990 0.11624 -0.0488 1.0000 0.0776 -12.000 -0.7747 0.05532 0.04985 -0.0871 0.9823 0.0421 -11.750 -0.7636 0.05163 0.04600 -0.0889 0.9770 0.0418 -11.500 -0.7563 0.04791 0.04197 -0.0905 0.9713 0.0417 -11.250 -0.7438 0.04424 0.03788 -0.0928 0.9680 0.0418 -11.000 -0.7360 0.04167 0.03494 -0.0924 0.9612 0.0420 -10.750 -0.7177 0.03904 0.03192 -0.0935 0.9571 0.0422 -10.500 -0.6912 0.03652 0.02905 -0.0953 0.9546 0.0423 -10.250 -0.6598 0.03412 0.02652 -0.0972 0.9530 0.0427 -10.000 -0.6348 0.03257 0.02490 -0.0976 0.9490 0.0432 -9.750 -0.6118 0.03137 0.02363 -0.0974 0.9440 0.0438 -9.500 -0.5832 0.03008 0.02224 -0.0982 0.9409 0.0445 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