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 = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.750 -1.1634 0.07109 0.06646 -0.0748 1.0000 0.0195 -19.500 -1.1655 0.06754 0.06283 -0.0764 0.9870 0.0196 -19.250 -1.1655 0.06429 0.05948 -0.0778 0.9558 0.0198 -19.000 -1.1666 0.06142 0.05647 -0.0782 0.9354 0.0199 -18.750 -1.1659 0.05866 0.05360 -0.0787 0.9215 0.0201 -18.250 -1.1598 0.05360 0.04833 -0.0798 0.9002 0.0205 -18.000 -1.1547 0.05134 0.04597 -0.0802 0.8912 0.0206 -17.750 -1.1489 0.04915 0.04369 -0.0806 0.8844 0.0208 -17.500 -1.1410 0.04719 0.04165 -0.0810 0.8775 0.0210 -17.250 -1.1337 0.04523 0.03960 -0.0813 0.8705 0.0212 -17.000 -1.1266 0.04328 0.03758 -0.0815 0.8650 0.0215 -16.750 -1.1187 0.04143 0.03568 -0.0817 0.8595 0.0218 -16.500 -1.1087 0.03982 0.03401 -0.0818 0.8535 0.0221 -16.250 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-0.0746 0.0759 0.7269 12.500 1.5128 0.03614 0.02933 -0.0738 0.0744 0.7296 12.750 1.5271 0.03724 0.03046 -0.0730 0.0728 0.7322 13.000 1.5395 0.03849 0.03174 -0.0721 0.0713 0.7347 13.250 1.5519 0.03976 0.03302 -0.0712 0.0697 0.7371 13.500 1.5657 0.04092 0.03422 -0.0705 0.0684 0.7393 13.750 1.5780 0.04224 0.03557 -0.0697 0.0670 0.7414 14.000 1.5900 0.04359 0.03695 -0.0689 0.0655 0.7433 14.250 1.6016 0.04497 0.03835 -0.0681 0.0640 0.7452 14.500 1.6118 0.04651 0.03993 -0.0673 0.0627 0.7478 14.750 1.6235 0.04792 0.04139 -0.0666 0.0617 0.7505 15.000 1.6351 0.04935 0.04287 -0.0659 0.0605 0.7531 15.250 1.6448 0.05098 0.04454 -0.0652 0.0592 0.7557 15.500 1.6534 0.05272 0.04632 -0.0645 0.0579 0.7584 15.750 1.6625 0.05443 0.04805 -0.0638 0.0568 0.7611 16.000 1.6716 0.05618 0.04986 -0.0632 0.0559 0.7636 16.250 1.6803 0.05795 0.05167 -0.0626 0.0550 0.7658 16.500 1.6888 0.05975 0.05353 -0.0621 0.0539 0.7687 16.750 1.6947 0.06188 0.05571 -0.0615 0.0529 0.7716 17.000 1.7008 0.06397 0.05785 -0.0610 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