XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY LS(1)-0413 (GA(W)-2) 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 = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.500 -0.4516 0.09762 0.09264 -0.0421 1.0000 0.1572 -9.250 -0.3810 0.09551 0.09101 -0.0312 1.0000 0.1637 -9.000 -0.4534 0.09266 0.08780 -0.0373 1.0000 0.1645 -8.750 -0.4675 0.09091 0.08614 -0.0347 1.0000 0.1693 -8.500 -0.5127 0.08902 0.08440 -0.0323 1.0000 0.1723 -8.250 -0.5731 0.08555 0.08107 -0.0329 1.0000 0.1735 -8.000 -0.5858 0.08176 0.07733 -0.0317 1.0000 0.1770 -7.750 -0.5597 0.08184 0.07745 -0.0253 1.0000 0.1836 -7.500 -0.6229 0.07492 0.07040 -0.0329 1.0000 0.1920 -7.250 -0.5944 0.07442 0.07009 -0.0264 1.0000 0.1970 -6.500 -0.5794 0.04394 0.03687 -0.0475 1.0000 0.0957 -6.250 -0.5494 0.03988 0.03184 -0.0488 1.0000 0.0837 -6.000 -0.5246 0.03655 0.02836 -0.0496 1.0000 0.0817 -5.750 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-1.000 -0.0441 0.02680 0.02013 -0.0300 0.9553 0.9031 -0.750 -0.0300 0.02662 0.01990 -0.0278 0.9487 0.9159 -0.500 -0.0043 0.02661 0.01981 -0.0278 0.9434 0.9290 -0.250 0.0246 0.02670 0.01983 -0.0286 0.9380 0.9393 0.000 0.0511 0.02669 0.01977 -0.0292 0.9302 0.9495 0.250 0.0956 0.02695 0.01995 -0.0328 0.9247 0.9597 0.500 0.1197 0.02693 0.01991 -0.0332 0.9149 0.9695 0.750 0.1714 0.02723 0.02016 -0.0383 0.9089 0.9766 1.000 0.2013 0.02740 0.02033 -0.0399 0.8985 0.9848 1.500 0.2775 0.02786 0.02079 -0.0457 0.8809 1.0000 1.750 0.3165 0.02804 0.02098 -0.0484 0.8710 1.0000 2.000 0.3653 0.02799 0.02094 -0.0524 0.8615 1.0000 2.250 0.3980 0.02807 0.02105 -0.0539 0.8490 1.0000 2.500 0.4404 0.02793 0.02096 -0.0566 0.8377 1.0000 2.750 0.5068 0.02680 0.01987 -0.0617 0.8281 1.0000 3.000 0.5452 0.02625 0.01939 -0.0631 0.8139 1.0000 3.250 0.5848 0.02568 0.01890 -0.0646 0.8010 1.0000 3.500 0.6562 0.02355 0.01687 -0.0694 0.7950 1.0000 3.750 0.6949 0.02251 0.01592 -0.0701 0.7806 1.0000 4.000 0.7341 0.02141 0.01493 -0.0707 0.7660 1.0000 4.250 0.7740 0.02024 0.01386 -0.0714 0.7501 1.0000 4.500 0.8152 0.01897 0.01268 -0.0721 0.7315 1.0000 4.750 0.8432 0.01830 0.01212 -0.0712 0.7037 1.0000 5.000 0.8753 0.01747 0.01132 -0.0708 0.6654 1.0000 5.250 0.9016 0.01700 0.01075 -0.0694 0.5924 1.0000 5.500 0.9224 0.01753 0.01026 -0.0671 0.4442 1.0000 5.750 0.9317 0.01927 0.01113 -0.0645 0.3392 1.0000 6.000 0.9473 0.02085 0.01215 -0.0632 0.2828 1.0000 6.250 0.9688 0.02224 0.01319 -0.0629 0.2483 1.0000 6.500 0.9937 0.02340 0.01420 -0.0631 0.2234 1.0000 6.750 1.0206 0.02460 0.01523 -0.0636 0.2055 1.0000 7.000 1.0489 0.02580 0.01634 -0.0643 0.1907 1.0000 7.250 1.0786 0.02706 0.01755 -0.0652 0.1785 1.0000 7.500 1.1101 0.02848 0.01882 -0.0664 0.1680 1.0000 7.750 1.1382 0.02966 0.02007 -0.0670 0.1583 1.0000 8.000 1.1691 0.03122 0.02166 -0.0680 0.1501 1.0000 8.250 1.1968 0.03252 0.02298 -0.0686 0.1422 1.0000 8.500 1.2258 0.03430 0.02487 -0.0693 0.1355 1.0000 8.750 1.2502 0.03567 0.02639 -0.0692 0.1290 1.0000 9.000 1.2788 0.03771 0.02839 -0.0701 0.1232 1.0000 9.250 1.2982 0.03936 0.03041 -0.0691 0.1181 1.0000 9.500 1.3211 0.04090 0.03201 -0.0689 0.1131 1.0000 9.750 1.3427 0.04347 0.03473 -0.0688 0.1089 1.0000 10.000 1.3542 0.04567 0.03738 -0.0667 0.1053 1.0000 10.250 1.3705 0.04762 0.03952 -0.0656 0.1015 1.0000 10.500 1.3914 0.04973 0.04161 -0.0655 0.0980 1.0000 10.750 1.3984 0.05330 0.04553 -0.0636 0.0959 1.0000 11.000 1.3934 0.05634 0.04908 -0.0599 0.0944 1.0000 11.250 1.3849 0.05965 0.05284 -0.0561 0.0930 1.0000 11.500 1.3714 0.06296 0.05650 -0.0520 0.0918 1.0000 11.750 1.3528 0.06637 0.06022 -0.0478 0.0910 1.0000 12.000 1.3304 0.07023 0.06439 -0.0442 0.0905 1.0000 12.250 1.3021 0.07489 0.06934 -0.0412 0.0904 1.0000 12.500 1.2663 0.08065 0.07541 -0.0394 0.0909 1.0000 12.750 1.2236 0.08789 0.08292 -0.0393 0.0918 1.0000 13.000 1.1773 0.09681 0.09206 -0.0415 0.0932 1.0000 13.250 1.1337 0.10717 0.10258 -0.0459 0.0947 1.0000 13.500 1.0995 0.11802 0.11353 -0.0514 0.0959 1.0000 13.750 1.0805 0.12761 0.12315 -0.0558 0.0967 1.0000