XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY LS(1)-0421MOD 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 -1.2058 0.07315 0.06843 -0.0537 1.0000 0.0194 -19.500 -1.2097 0.06926 0.06447 -0.0552 1.0000 0.0195 -19.250 -1.2119 0.06567 0.06080 -0.0566 1.0000 0.0196 -19.000 -1.2125 0.06234 0.05740 -0.0577 1.0000 0.0197 -18.750 -1.2114 0.05928 0.05428 -0.0587 1.0000 0.0199 -18.500 -1.2085 0.05649 0.05143 -0.0596 1.0000 0.0200 -18.250 -1.2049 0.05385 0.04873 -0.0603 1.0000 0.0202 -18.000 -1.2006 0.05133 0.04616 -0.0609 1.0000 0.0204 -17.750 -1.1941 0.04910 0.04388 -0.0614 1.0000 0.0205 -17.500 -1.1875 0.04691 0.04164 -0.0618 1.0000 0.0207 -17.250 -1.1795 0.04493 0.03961 -0.0621 1.0000 0.0209 -17.000 -1.1711 0.04303 0.03766 -0.0624 1.0000 0.0211 -16.750 -1.1614 0.04128 0.03587 -0.0625 1.0000 0.0213 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1.4137 0.01705 0.01008 -0.0930 0.1323 0.7028 8.500 1.4300 0.01751 0.01056 -0.0912 0.1293 0.7050 8.750 1.4482 0.01799 0.01105 -0.0899 0.1266 0.7070 9.000 1.4655 0.01855 0.01161 -0.0885 0.1236 0.7086 9.250 1.4817 0.01919 0.01225 -0.0872 0.1206 0.7107 9.500 1.4991 0.01979 0.01288 -0.0860 0.1183 0.7131 9.750 1.5162 0.02044 0.01356 -0.0849 0.1159 0.7155 10.000 1.5318 0.02120 0.01434 -0.0837 0.1133 0.7178 10.250 1.5463 0.02206 0.01521 -0.0824 0.1106 0.7201 10.500 1.5618 0.02290 0.01608 -0.0814 0.1085 0.7224 10.750 1.5773 0.02376 0.01698 -0.0804 0.1066 0.7246 11.000 1.5917 0.02473 0.01797 -0.0794 0.1046 0.7267 11.250 1.6049 0.02581 0.01907 -0.0784 0.1026 0.7286 11.500 1.6169 0.02702 0.02029 -0.0774 0.1006 0.7303 11.750 1.6306 0.02812 0.02144 -0.0765 0.0992 0.7325 12.000 1.6437 0.02929 0.02266 -0.0757 0.0977 0.7352 12.250 1.6554 0.03060 0.02402 -0.0749 0.0961 0.7381 12.500 1.6660 0.03201 0.02547 -0.0740 0.0945 0.7409 12.750 1.6750 0.03360 0.02709 -0.0732 0.0928 0.7436 13.000 1.6834 0.03525 0.02879 -0.0723 0.0913 0.7461 13.250 1.6936 0.03676 0.03035 -0.0716 0.0901 0.7484 13.500 1.7019 0.03848 0.03212 -0.0710 0.0887 0.7505 13.750 1.7093 0.04032 0.03400 -0.0703 0.0873 0.7523 14.000 1.7153 0.04232 0.03605 -0.0696 0.0858 0.7550 14.250 1.7198 0.04452 0.03830 -0.0690 0.0844 0.7577 14.500 1.7250 0.04668 0.04052 -0.0685 0.0833 0.7605 14.750 1.7313 0.04879 0.04270 -0.0681 0.0824 0.7635 15.000 1.7363 0.05106 0.04504 -0.0678 0.0813 0.7666 15.250 1.7408 0.05339 0.04743 -0.0675 0.0802 0.7694 15.500 1.7429 0.05604 0.05013 -0.0672 0.0792 0.7719 16.000 1.7453 0.06162 0.05582 -0.0669 0.0770 0.7769 16.250 1.7445 0.06472 0.05899 -0.0669 0.0760 0.7798 16.500 1.7471 0.06738 0.06173 -0.0668 0.0753 0.7830 16.750 1.7475 0.07038 0.06481 -0.0669 0.0745 0.7861 17.000 1.7475 0.07342 0.06792 -0.0670 0.0735 0.7889 17.250 1.7466 0.07662 0.07118 -0.0672 0.0726 0.7913 17.500 1.7435 0.08010 0.07472 -0.0675 0.0718 0.7933 17.750 1.7413 0.08346 0.07815 -0.0678 0.0710 0.7955 18.000 1.7360 0.08728 0.08205 -0.0683 0.0702 0.7979