XFOIL Version 6.94 Calculated polar for: NASA/LANGLEY MS(1)-0317 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 0.3580 0.01377 0.00840 -0.0728 0.7480 0.7358 0.500 0.4158 0.01365 0.00821 -0.0726 0.7330 0.7411 1.000 0.4760 0.01345 0.00795 -0.0733 0.7154 0.7478 1.500 0.5392 0.01321 0.00767 -0.0750 0.6961 0.7540 2.000 0.5944 0.01306 0.00753 -0.0744 0.6743 0.7583 2.500 0.6514 0.01298 0.00741 -0.0743 0.6479 0.7628 3.000 0.7084 0.01296 0.00732 -0.0745 0.6083 0.7684 3.500 0.7637 0.01326 0.00727 -0.0746 0.5288 0.7744 4.000 0.7990 0.01460 0.00779 -0.0713 0.3728 0.7788 4.500 0.8334 0.01612 0.00871 -0.0681 0.2658 0.7836 5.000 0.8746 0.01729 0.00954 -0.0662 0.2101 0.7885 5.500 0.9190 0.01829 0.01036 -0.0649 0.1816 0.7940 6.000 0.9618 0.01924 0.01122 -0.0633 0.1650 0.7990 6.500 0.9967 0.02036 0.01227 -0.0602 0.1535 0.8034 7.000 1.0338 0.02116 0.01316 -0.0573 0.1464 0.8084 7.500 1.0681 0.02253 0.01445 -0.0546 0.1392 0.8140 8.000 1.1110 0.02360 0.01562 -0.0533 0.1344 0.8198 8.500 1.1489 0.02474 0.01683 -0.0511 0.1301 0.8248 9.000 1.1881 0.02629 0.01834 -0.0494 0.1257 0.8303 9.500 1.2301 0.02764 0.01982 -0.0482 0.1225 0.8370 10.000 1.2698 0.02896 0.02127 -0.0466 0.1193 0.8430 10.500 1.3069 0.03037 0.02275 -0.0448 0.1161 0.8493 11.000 1.3562 0.03223 0.02455 -0.0446 0.1124 0.8560 11.500 1.3871 0.03377 0.02635 -0.0423 0.1100 0.8631 12.000 1.4158 0.03550 0.02829 -0.0399 0.1070 0.8721 12.500 1.4469 0.03730 0.03017 -0.0381 0.1040 0.8812 13.000 1.4854 0.03942 0.03233 -0.0367 0.1008 0.8906 13.500 1.5028 0.04181 0.03505 -0.0340 0.0986 0.9020 14.000 1.5204 0.04440 0.03790 -0.0314 0.0961 0.9160 15.000 1.5621 0.04903 0.04270 -0.0264 0.0907 1.0000 15.500 1.5648 0.05352 0.04759 -0.0255 0.0887 1.0000 16.000 1.5683 0.05830 0.05265 -0.0250 0.0863 1.0000 16.500 1.5776 0.06258 0.05706 -0.0248 0.0839 1.0000 17.000 1.5979 0.06628 0.06073 -0.0244 0.0813 1.0000 17.500 1.5730 0.07442 0.06934 -0.0253 0.0795 1.0000 18.000 1.5521 0.08311 0.07839 -0.0273 0.0773 1.0000 18.500 1.5437 0.09065 0.08612 -0.0297 0.0753 1.0000 19.000 1.5587 0.09490 0.09032 -0.0306 0.0732 1.0000 19.500 1.5007 0.11094 0.10691 -0.0384 0.0718 1.0000