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 = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.750 -1.1510 0.09368 0.09004 -0.0430 1.0000 0.0186 -19.500 -1.1630 0.08800 0.08426 -0.0459 1.0000 0.0187 -19.250 -1.1736 0.08257 0.07873 -0.0487 1.0000 0.0189 -19.000 -1.1825 0.07752 0.07357 -0.0512 1.0000 0.0190 -18.750 -1.1898 0.07285 0.06880 -0.0534 1.0000 0.0191 -18.500 -1.1954 0.06849 0.06435 -0.0554 1.0000 0.0192 -18.250 -1.1993 0.06447 0.06023 -0.0572 1.0000 0.0194 -18.000 -1.2013 0.06082 0.05649 -0.0586 1.0000 0.0195 -17.750 -1.2018 0.05748 0.05305 -0.0599 1.0000 0.0197 -17.500 -1.2007 0.05440 0.04989 -0.0609 1.0000 0.0198 -17.250 -1.1979 0.05161 0.04700 -0.0618 1.0000 0.0200 -17.000 -1.1935 0.04909 0.04441 -0.0625 1.0000 0.0201 -16.750 -1.1902 0.04656 0.04180 -0.0630 1.0000 0.0202 -16.500 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0.01877 -0.0609 0.0807 0.7886 12.500 1.5745 0.02587 0.01965 -0.0599 0.0798 0.7920 12.750 1.5883 0.02683 0.02065 -0.0587 0.0789 0.7954 13.000 1.5992 0.02803 0.02187 -0.0574 0.0777 0.7984 13.250 1.6074 0.02945 0.02336 -0.0559 0.0766 0.8026 13.500 1.6240 0.03025 0.02424 -0.0552 0.0762 0.8066 13.750 1.6400 0.03110 0.02515 -0.0544 0.0755 0.8108 14.000 1.6551 0.03204 0.02615 -0.0536 0.0748 0.8149 14.250 1.6696 0.03302 0.02720 -0.0529 0.0739 0.8193 14.500 1.6830 0.03411 0.02835 -0.0520 0.0730 0.8243 14.750 1.6962 0.03522 0.02952 -0.0512 0.0721 0.8295 15.000 1.7074 0.03651 0.03084 -0.0504 0.0712 0.8344 15.250 1.7166 0.03796 0.03235 -0.0494 0.0702 0.8397 15.500 1.7204 0.03991 0.03436 -0.0482 0.0690 0.8452 15.750 1.7310 0.04133 0.03585 -0.0475 0.0683 0.8509 16.000 1.7438 0.04256 0.03718 -0.0470 0.0677 0.8570 16.250 1.7559 0.04387 0.03857 -0.0464 0.0668 0.8638 16.500 1.7665 0.04533 0.04010 -0.0459 0.0658 0.8705 16.750 1.7770 0.04680 0.04165 -0.0453 0.0648 0.8784 17.000 1.7852 0.04856 0.04346 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