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 = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.750 -0.9798 0.09144 0.08680 -0.0747 1.0000 0.0298 -19.500 -0.9849 0.08708 0.08233 -0.0770 1.0000 0.0300 -19.250 -0.9887 0.08299 0.07813 -0.0791 1.0000 0.0302 -19.000 -0.9913 0.07915 0.07419 -0.0809 1.0000 0.0305 -18.750 -0.9928 0.07550 0.07044 -0.0827 1.0000 0.0307 -18.500 -0.9935 0.07200 0.06682 -0.0843 1.0000 0.0310 -18.250 -0.9934 0.06866 0.06338 -0.0858 1.0000 0.0313 -18.000 -0.9925 0.06551 0.06012 -0.0871 1.0000 0.0315 -17.750 -0.9909 0.06248 0.05696 -0.0884 1.0000 0.0318 -17.500 -0.9878 0.05970 0.05408 -0.0895 1.0000 0.0320 -17.250 -0.9784 0.05796 0.05235 -0.0896 1.0000 0.0324 -17.000 -0.9693 0.05604 0.05042 -0.0901 1.0000 0.0328 -16.750 -0.9605 0.05403 0.04840 -0.0909 1.0000 0.0331 -16.500 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-0.0642 0.1133 0.7930 12.250 1.4012 0.04195 0.03498 -0.0630 0.1104 0.7956 12.500 1.4157 0.04311 0.03622 -0.0624 0.1082 0.7985 12.750 1.4285 0.04442 0.03756 -0.0617 0.1057 0.8013 13.000 1.4393 0.04593 0.03907 -0.0609 0.1033 0.8040 13.250 1.4461 0.04780 0.04093 -0.0599 0.1009 0.8062 13.500 1.4600 0.04903 0.04226 -0.0594 0.0989 0.8089 13.750 1.4723 0.05041 0.04371 -0.0587 0.0967 0.8115 14.000 1.4813 0.05211 0.04544 -0.0580 0.0947 0.8142 14.250 1.4872 0.05412 0.04745 -0.0571 0.0927 0.8169 14.500 1.4974 0.05575 0.04915 -0.0565 0.0910 0.8199 14.750 1.5081 0.05741 0.05089 -0.0561 0.0893 0.8229 15.000 1.5179 0.05918 0.05269 -0.0557 0.0875 0.8256 15.250 1.5249 0.06114 0.05470 -0.0551 0.0857 0.8287 15.500 1.5273 0.06356 0.05713 -0.0542 0.0840 0.8316 15.750 1.5371 0.06533 0.05901 -0.0539 0.0827 0.8350 16.000 1.5453 0.06731 0.06107 -0.0535 0.0812 0.8386 16.250 1.5526 0.06942 0.06324 -0.0533 0.0797 0.8419 16.500 1.5584 0.07167 0.06553 -0.0530 0.0784 0.8450 16.750 1.5604 0.07428 0.06817 -0.0526 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