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 = 1.000 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -18.250 -0.9770 0.11413 0.11160 -0.0215 1.0000 0.0086 -18.000 -1.0483 0.09596 0.09317 -0.0314 1.0000 0.0084 -17.750 -1.1105 0.07973 0.07670 -0.0410 1.0000 0.0084 -17.500 -1.1454 0.06904 0.06579 -0.0478 1.0000 0.0084 -17.250 -1.1673 0.06138 0.05797 -0.0526 1.0000 0.0084 -17.000 -1.1816 0.05553 0.05196 -0.0560 1.0000 0.0084 -16.750 -1.1915 0.05080 0.04709 -0.0584 1.0000 0.0084 -16.500 -1.1976 0.04683 0.04300 -0.0602 1.0000 0.0084 -16.250 -1.2043 0.04303 0.03908 -0.0617 1.0000 0.0085 -16.000 -1.2086 0.03969 0.03562 -0.0627 1.0000 0.0086 -15.750 -1.2093 0.03690 0.03272 -0.0633 1.0000 0.0087 -15.500 -1.2071 0.03450 0.03023 -0.0636 1.0000 0.0088 -15.250 -1.2022 0.03245 0.02809 -0.0638 1.0000 0.0089 -15.000 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0.0617 0.7962 10.250 1.4700 0.01728 0.01074 -0.0771 0.0598 0.8004 10.500 1.4912 0.01769 0.01117 -0.0761 0.0579 0.8046 10.750 1.5118 0.01809 0.01161 -0.0751 0.0566 0.8086 11.000 1.5323 0.01844 0.01203 -0.0740 0.0554 0.8131 11.250 1.5515 0.01884 0.01247 -0.0727 0.0536 0.8183 11.500 1.5673 0.01928 0.01294 -0.0709 0.0517 0.8236 11.750 1.5835 0.01981 0.01350 -0.0692 0.0494 0.8289 12.000 1.6006 0.02031 0.01406 -0.0676 0.0476 0.8347 12.250 1.6170 0.02089 0.01467 -0.0661 0.0449 0.8404 12.500 1.6316 0.02160 0.01539 -0.0645 0.0416 0.8467 12.750 1.6470 0.02226 0.01610 -0.0629 0.0391 0.8540 13.000 1.6601 0.02309 0.01697 -0.0612 0.0360 0.8621 13.250 1.6733 0.02392 0.01784 -0.0595 0.0335 0.8719 13.500 1.6846 0.02487 0.01885 -0.0577 0.0309 0.8837 13.750 1.6962 0.02580 0.01985 -0.0559 0.0291 0.8989 14.000 1.7060 0.02669 0.02086 -0.0539 0.0276 0.9266 14.250 1.7126 0.02767 0.02193 -0.0514 0.0264 1.0000 14.500 1.7229 0.02887 0.02318 -0.0499 0.0251 1.0000 14.750 1.7341 0.03000 0.02437 -0.0486 0.0243 1.0000 15.000 1.7439 0.03126 0.02569 -0.0472 0.0233 1.0000 15.250 1.7522 0.03268 0.02715 -0.0458 0.0222 1.0000 15.500 1.7589 0.03425 0.02878 -0.0444 0.0213 1.0000 15.750 1.7664 0.03576 0.03037 -0.0431 0.0206 1.0000 16.000 1.7733 0.03737 0.03206 -0.0419 0.0199 1.0000 16.250 1.7786 0.03916 0.03392 -0.0408 0.0192 1.0000 16.500 1.7820 0.04116 0.03599 -0.0396 0.0184 1.0000 16.750 1.7835 0.04338 0.03829 -0.0385 0.0177 1.0000 17.000 1.7853 0.04565 0.04063 -0.0377 0.0171 1.0000 17.250 1.7867 0.04804 0.04311 -0.0369 0.0166 1.0000 17.500 1.7862 0.05071 0.04588 -0.0364 0.0160 1.0000 17.750 1.7833 0.05376 0.04902 -0.0360 0.0154 1.0000 18.000 1.7785 0.05718 0.05253 -0.0359 0.0148 1.0000 18.250 1.7721 0.06095 0.05640 -0.0361 0.0144 1.0000 18.500 1.7653 0.06492 0.06049 -0.0366 0.0140 1.0000 18.750 1.7567 0.06932 0.06501 -0.0375 0.0136 1.0000 19.000 1.7450 0.07436 0.07019 -0.0388 0.0133 1.0000 19.250 1.7295 0.08020 0.07617 -0.0408 0.0130 1.0000