XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY MS(1)-0313 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -13.000 -0.6729 0.08134 0.07874 -0.0447 1.0000 0.0200 -12.750 -0.7111 0.06851 0.06576 -0.0551 1.0000 0.0196 -12.500 -0.7468 0.05917 0.05621 -0.0620 1.0000 0.0193 -12.250 -0.7803 0.05184 0.04862 -0.0659 1.0000 0.0191 -12.000 -0.8093 0.04609 0.04259 -0.0673 1.0000 0.0190 -11.750 -0.8322 0.04195 0.03821 -0.0666 1.0000 0.0190 -11.500 -0.8506 0.03883 0.03484 -0.0643 1.0000 0.0191 -11.250 -0.8673 0.03666 0.03246 -0.0599 1.0000 0.0192 -11.000 -0.8690 0.03386 0.02934 -0.0582 0.9980 0.0195 -10.750 -0.8480 0.03052 0.02558 -0.0604 0.9948 0.0198 -10.500 -0.8227 0.02818 0.02289 -0.0621 0.9918 0.0201 -10.250 -0.7966 0.02585 0.02026 -0.0637 0.9885 0.0205 -10.000 -0.7677 0.02389 0.01821 -0.0654 0.9862 0.0211 -9.750 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0.1411 0.8054 5.000 0.9138 0.01392 0.00649 -0.0778 0.1346 0.8090 5.250 0.9376 0.01434 0.00688 -0.0772 0.1281 0.8120 5.500 0.9636 0.01457 0.00715 -0.0769 0.1248 0.8155 5.750 0.9891 0.01486 0.00747 -0.0765 0.1210 0.8194 6.000 1.0139 0.01526 0.00784 -0.0762 0.1173 0.8233 6.250 1.0369 0.01577 0.00836 -0.0755 0.1136 0.8268 6.500 1.0619 0.01602 0.00868 -0.0750 0.1116 0.8304 6.750 1.0866 0.01632 0.00902 -0.0745 0.1091 0.8346 7.000 1.1112 0.01667 0.00940 -0.0741 0.1068 0.8389 7.250 1.1344 0.01709 0.00984 -0.0735 0.1045 0.8427 7.500 1.1552 0.01769 0.01045 -0.0724 0.1020 0.8467 7.750 1.1766 0.01829 0.01110 -0.0714 0.1001 0.8515 8.000 1.2010 0.01861 0.01148 -0.0710 0.0987 0.8565 8.250 1.2235 0.01891 0.01187 -0.0702 0.0969 0.8609 8.500 1.2461 0.01926 0.01227 -0.0694 0.0948 0.8659 8.750 1.2685 0.01964 0.01269 -0.0687 0.0928 0.8714 9.000 1.2882 0.02014 0.01322 -0.0675 0.0909 0.8768 9.250 1.3053 0.02108 0.01418 -0.0660 0.0883 0.8831 9.500 1.3260 0.02148 0.01467 -0.0650 0.0873 0.8896 9.750 1.3455 0.02181 0.01511 -0.0637 0.0860 0.8968 10.000 1.3634 0.02217 0.01557 -0.0621 0.0844 0.9050 10.250 1.3790 0.02255 0.01604 -0.0602 0.0828 0.9152 10.500 1.3940 0.02294 0.01651 -0.0582 0.0813 0.9291 10.750 1.4043 0.02328 0.01693 -0.0553 0.0798 0.9666 11.000 1.4207 0.02415 0.01780 -0.0541 0.0777 1.0000 11.250 1.4388 0.02506 0.01877 -0.0532 0.0760 1.0000 11.500 1.4580 0.02554 0.01936 -0.0525 0.0747 1.0000 11.750 1.4763 0.02612 0.02001 -0.0516 0.0729 1.0000 12.000 1.4936 0.02675 0.02070 -0.0507 0.0711 1.0000 12.250 1.5100 0.02743 0.02142 -0.0497 0.0694 1.0000 12.500 1.5232 0.02842 0.02241 -0.0484 0.0674 1.0000 12.750 1.5362 0.02950 0.02356 -0.0471 0.0655 1.0000 13.000 1.5526 0.03019 0.02437 -0.0462 0.0639 1.0000 13.250 1.5680 0.03098 0.02524 -0.0453 0.0620 1.0000 13.500 1.5827 0.03183 0.02613 -0.0444 0.0600 1.0000 13.750 1.5927 0.03309 0.02739 -0.0431 0.0579 1.0000 14.000 1.6043 0.03429 0.02868 -0.0420 0.0559 1.0000 14.250 1.6190 0.03521 0.02971 -0.0413 0.0537 1.0000 14.500 1.6318 0.03629 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