XFOIL Version 6.94 Calculated polar for: NASA/LANGLEY MS(1)-0413 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.3693 0.01122 0.00617 -0.0782 0.8219 0.7754 0.500 0.4247 0.01107 0.00600 -0.0774 0.8049 0.7829 1.000 0.4838 0.01089 0.00576 -0.0777 0.7857 0.7910 1.500 0.5393 0.01069 0.00555 -0.0771 0.7623 0.7977 2.000 0.5951 0.01057 0.00542 -0.0765 0.7315 0.8047 2.500 0.6523 0.01055 0.00526 -0.0764 0.6804 0.8129 3.000 0.6978 0.01110 0.00519 -0.0739 0.5318 0.8193 3.500 0.7344 0.01296 0.00597 -0.0710 0.3280 0.8271 4.000 0.7802 0.01432 0.00672 -0.0700 0.2240 0.8352 4.500 0.8258 0.01521 0.00741 -0.0683 0.1831 0.8422 5.000 0.8746 0.01606 0.00817 -0.0674 0.1613 0.8512 5.500 0.9220 0.01690 0.00897 -0.0662 0.1480 0.8588 6.000 0.9668 0.01790 0.00994 -0.0645 0.1380 0.8680 7.000 1.0584 0.01989 0.01199 -0.0615 0.1246 0.8870 7.500 1.1045 0.02068 0.01294 -0.0600 0.1194 0.8982 8.000 1.1484 0.02175 0.01404 -0.0583 0.1148 0.9112 8.500 1.1907 0.02300 0.01543 -0.0562 0.1105 0.9258 9.000 1.2291 0.02381 0.01640 -0.0533 0.1060 0.9483 9.500 1.2763 0.02539 0.01796 -0.0528 0.1011 1.0000 10.000 1.3247 0.02660 0.01944 -0.0528 0.0969 1.0000 10.500 1.3706 0.02799 0.02084 -0.0524 0.0926 1.0000 11.000 1.4106 0.02983 0.02289 -0.0510 0.0883 1.0000 11.500 1.4410 0.03120 0.02439 -0.0480 0.0841 1.0000 12.000 1.4715 0.03318 0.02650 -0.0455 0.0801 1.0000 12.500 1.4916 0.03493 0.02849 -0.0418 0.0759 1.0000 13.000 1.5141 0.03718 0.03083 -0.0389 0.0721 1.0000 13.500 1.5263 0.03964 0.03359 -0.0355 0.0680 1.0000 14.000 1.5406 0.04252 0.03658 -0.0329 0.0645 1.0000 14.500 1.5475 0.04600 0.04036 -0.0305 0.0607 1.0000 15.000 1.5539 0.05000 0.04451 -0.0287 0.0576 1.0000 15.500 1.5536 0.05496 0.04979 -0.0276 0.0543 1.0000 16.000 1.5530 0.06032 0.05526 -0.0270 0.0517 1.0000 16.500 1.5398 0.06796 0.06332 -0.0279 0.0490 1.0000 17.000 1.5334 0.07524 0.07072 -0.0298 0.0468 1.0000 17.500 1.5071 0.08633 0.08220 -0.0339 0.0449 1.0000 18.000 1.4780 0.09892 0.09515 -0.0399 0.0432 1.0000 18.500 1.4564 0.11056 0.10695 -0.0461 0.0417 1.0000 19.000 1.4171 0.12611 0.12279 -0.0551 0.0405 1.0000 19.500 1.3389 0.15150 0.14868 -0.0721 0.0401 1.0000