XFOIL Version 6.96 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 = 0.000 Re = 0.500 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.750 -1.0147 0.11643 0.11243 -0.0314 1.0000 0.0186 -19.500 -1.0319 0.10935 0.10523 -0.0353 1.0000 0.0187 -19.250 -1.0482 0.10253 0.09828 -0.0390 1.0000 0.0188 -19.000 -1.0634 0.09597 0.09159 -0.0426 1.0000 0.0189 -18.750 -1.0768 0.08985 0.08535 -0.0458 1.0000 0.0190 -18.500 -1.0886 0.08409 0.07946 -0.0488 1.0000 0.0191 -18.250 -1.0982 0.07877 0.07401 -0.0515 1.0000 0.0193 -18.000 -1.1060 0.07388 0.06900 -0.0540 1.0000 0.0194 -17.750 -1.1118 0.06941 0.06442 -0.0561 1.0000 0.0196 -17.500 -1.1157 0.06531 0.06020 -0.0579 1.0000 0.0197 -17.250 -1.1177 0.06160 0.05637 -0.0595 1.0000 0.0199 -17.000 -1.1180 0.05825 0.05291 -0.0608 1.0000 0.0201 -16.750 -1.1185 0.05498 0.04955 -0.0619 1.0000 0.0203 -16.500 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0.02363 -0.0531 0.0798 0.7830 12.500 1.4967 0.03185 0.02475 -0.0522 0.0788 0.7862 12.750 1.5090 0.03301 0.02597 -0.0511 0.0779 0.7896 13.000 1.5197 0.03433 0.02733 -0.0501 0.0770 0.7934 13.250 1.5320 0.03554 0.02861 -0.0492 0.0763 0.7972 13.500 1.5455 0.03665 0.02982 -0.0484 0.0755 0.8008 13.750 1.5582 0.03782 0.03109 -0.0475 0.0746 0.8048 14.000 1.5702 0.03908 0.03243 -0.0467 0.0737 0.8089 14.250 1.5817 0.04039 0.03383 -0.0460 0.0728 0.8131 14.750 1.6019 0.04332 0.03690 -0.0445 0.0710 0.8211 15.000 1.6104 0.04496 0.03860 -0.0438 0.0702 0.8257 15.250 1.6174 0.04678 0.04048 -0.0431 0.0694 0.8305 15.750 1.6337 0.05030 0.04420 -0.0419 0.0677 0.8400 16.000 1.6428 0.05202 0.04603 -0.0415 0.0668 0.8454 16.250 1.6506 0.05391 0.04802 -0.0411 0.0657 0.8510 16.500 1.6575 0.05588 0.05008 -0.0408 0.0647 0.8567 16.750 1.6631 0.05805 0.05235 -0.0405 0.0638 0.8635 17.000 1.6676 0.06039 0.05477 -0.0403 0.0629 0.8703 17.250 1.6703 0.06295 0.05741 -0.0402 0.0621 0.8781 17.500 1.6724 0.06563 0.06019 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