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 = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -18.750 -1.0041 0.09878 0.09489 -0.0460 1.0000 0.0248 -18.500 -1.0236 0.09139 0.08732 -0.0504 1.0000 0.0249 -18.250 -1.0387 0.08501 0.08079 -0.0540 1.0000 0.0250 -18.000 -1.0509 0.07935 0.07498 -0.0571 1.0000 0.0252 -17.750 -1.0601 0.07434 0.06983 -0.0597 1.0000 0.0253 -17.500 -1.0672 0.06981 0.06517 -0.0618 1.0000 0.0255 -17.250 -1.0728 0.06569 0.06091 -0.0636 1.0000 0.0257 -17.000 -1.0769 0.06190 0.05699 -0.0651 1.0000 0.0259 -16.750 -1.0797 0.05841 0.05336 -0.0664 1.0000 0.0262 -16.500 -1.0813 0.05521 0.05003 -0.0673 1.0000 0.0264 -16.250 -1.0817 0.05227 0.04695 -0.0681 1.0000 0.0267 -16.000 -1.0810 0.04959 0.04413 -0.0686 1.0000 0.0269 -15.750 -1.0796 0.04715 0.04154 -0.0689 1.0000 0.0271 -15.500 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0.3351 0.7422 3.750 0.8012 0.01265 0.00586 -0.0812 0.2860 0.7442 4.000 0.8242 0.01317 0.00619 -0.0805 0.2484 0.7463 4.250 0.8478 0.01364 0.00651 -0.0799 0.2197 0.7484 4.500 0.8719 0.01406 0.00681 -0.0794 0.1980 0.7506 4.750 0.8964 0.01446 0.00712 -0.0790 0.1816 0.7527 5.000 0.9210 0.01484 0.00742 -0.0786 0.1689 0.7549 5.250 0.9448 0.01527 0.00776 -0.0782 0.1582 0.7570 5.500 0.9702 0.01558 0.00804 -0.0779 0.1508 0.7593 5.750 0.9929 0.01603 0.00842 -0.0772 0.1435 0.7614 6.000 1.0173 0.01627 0.00870 -0.0767 0.1391 0.7635 6.250 1.0400 0.01664 0.00906 -0.0760 0.1346 0.7654 6.500 1.0597 0.01718 0.00956 -0.0747 0.1300 0.7675 6.750 1.0832 0.01746 0.00989 -0.0741 0.1276 0.7696 7.000 1.1056 0.01781 0.01027 -0.0733 0.1249 0.7719 7.250 1.1265 0.01822 0.01068 -0.0723 0.1223 0.7743 7.500 1.1449 0.01873 0.01117 -0.0709 0.1197 0.7769 7.750 1.1567 0.01939 0.01182 -0.0684 0.1172 0.7793 8.000 1.1759 0.01973 0.01223 -0.0671 0.1159 0.7818 8.250 1.1950 0.02019 0.01274 -0.0659 0.1144 0.7841 8.500 1.2138 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-0.0450 0.0891 0.8483 13.500 1.5290 0.03765 0.03122 -0.0442 0.0881 0.8528 13.750 1.5400 0.03888 0.03249 -0.0431 0.0870 0.8575 14.000 1.5517 0.04022 0.03384 -0.0418 0.0854 0.8625 14.250 1.5631 0.04150 0.03523 -0.0411 0.0848 0.8681 14.500 1.5730 0.04283 0.03670 -0.0402 0.0840 0.8734 14.750 1.5827 0.04423 0.03822 -0.0394 0.0830 0.8795 15.000 1.5921 0.04570 0.03979 -0.0387 0.0820 0.8860 15.250 1.6006 0.04720 0.04139 -0.0379 0.0809 0.8929 15.500 1.6093 0.04878 0.04304 -0.0373 0.0799 0.9010 15.750 1.6164 0.05042 0.04475 -0.0366 0.0789 0.9102 16.000 1.6229 0.05209 0.04647 -0.0357 0.0778 0.9218 16.250 1.6277 0.05369 0.04813 -0.0342 0.0765 0.9404 16.500 1.6293 0.05564 0.05024 -0.0334 0.0759 1.0000 16.750 1.6354 0.05796 0.05269 -0.0336 0.0749 1.0000 17.000 1.6412 0.06031 0.05513 -0.0338 0.0738 1.0000 17.250 1.6466 0.06270 0.05761 -0.0341 0.0728 1.0000 17.500 1.6513 0.06520 0.06017 -0.0344 0.0717 1.0000 17.750 1.6555 0.06776 0.06278 -0.0347 0.0708 1.0000 18.000 1.6588 0.07041 0.06546 -0.0350 0.0698 1.0000 18.250 1.6602 0.07314 0.06820 -0.0350 0.0685 1.0000 18.500 1.6593 0.07662 0.07185 -0.0359 0.0677 1.0000 18.750 1.6582 0.08012 0.07549 -0.0368 0.0667 1.0000 19.000 1.6570 0.08365 0.07912 -0.0378 0.0656 1.0000 19.250 1.6557 0.08724 0.08280 -0.0389 0.0646 1.0000