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 = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.750 -1.0906 0.08018 0.07589 -0.0722 1.0000 0.0238 -19.500 -1.0931 0.07641 0.07204 -0.0739 1.0000 0.0239 -19.250 -1.0939 0.07291 0.06846 -0.0754 1.0000 0.0241 -19.000 -1.0935 0.06958 0.06505 -0.0768 1.0000 0.0242 -18.750 -1.0920 0.06640 0.06180 -0.0782 1.0000 0.0244 -18.500 -1.0894 0.06341 0.05874 -0.0794 1.0000 0.0245 -18.250 -1.0925 0.05989 0.05515 -0.0807 1.0000 0.0247 -18.000 -1.0949 0.05649 0.05170 -0.0818 1.0000 0.0250 -17.750 -1.0935 0.05357 0.04873 -0.0828 1.0000 0.0252 -17.500 -1.0880 0.05096 0.04607 -0.0842 0.9793 0.0255 -17.250 -1.0832 0.04873 0.04375 -0.0845 0.9545 0.0258 -17.000 -1.0779 0.04686 0.04179 -0.0842 0.9374 0.0261 -16.750 -1.0717 0.04499 0.03982 -0.0841 0.9248 0.0264 -16.500 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-0.0754 0.0932 0.7577 12.250 1.5018 0.03476 0.02825 -0.0745 0.0912 0.7601 12.500 1.5168 0.03578 0.02931 -0.0737 0.0894 0.7623 12.750 1.5305 0.03692 0.03046 -0.0729 0.0874 0.7644 13.000 1.5414 0.03829 0.03183 -0.0718 0.0853 0.7664 13.250 1.5534 0.03959 0.03317 -0.0709 0.0834 0.7699 13.500 1.5682 0.04066 0.03430 -0.0703 0.0819 0.7728 13.750 1.5802 0.04200 0.03568 -0.0694 0.0801 0.7754 14.000 1.5910 0.04345 0.03714 -0.0685 0.0783 0.7780 14.250 1.5994 0.04510 0.03881 -0.0675 0.0763 0.7805 14.500 1.6124 0.04641 0.04018 -0.0669 0.0752 0.7831 14.750 1.6240 0.04782 0.04163 -0.0662 0.0738 0.7853 15.000 1.6349 0.04931 0.04315 -0.0655 0.0723 0.7878 15.250 1.6430 0.05110 0.04497 -0.0647 0.0707 0.7911 15.500 1.6493 0.05305 0.04697 -0.0638 0.0692 0.7940 15.750 1.6614 0.05449 0.04847 -0.0633 0.0681 0.7970 16.000 1.6703 0.05625 0.05029 -0.0627 0.0669 0.8000 16.250 1.6785 0.05808 0.05215 -0.0621 0.0657 0.8030 16.500 1.6851 0.06008 0.05418 -0.0615 0.0644 0.8055 16.750 1.6872 0.06260 0.05674 -0.0608 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