XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY LS(1)-0417MOD 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 ------ -------- --------- --------- -------- -------- -------- -19.750 -1.0578 0.10687 0.10342 -0.0384 1.0000 0.0116 -19.500 -1.0839 0.09852 0.09494 -0.0429 1.0000 0.0116 -19.250 -1.1076 0.09073 0.08701 -0.0471 1.0000 0.0116 -19.000 -1.1279 0.08362 0.07975 -0.0509 1.0000 0.0117 -18.750 -1.1436 0.07740 0.07341 -0.0542 1.0000 0.0117 -18.500 -1.1568 0.07164 0.06754 -0.0572 1.0000 0.0117 -18.250 -1.1657 0.06668 0.06246 -0.0597 1.0000 0.0117 -18.000 -1.1727 0.06213 0.05780 -0.0619 1.0000 0.0118 -17.750 -1.1777 0.05798 0.05356 -0.0638 1.0000 0.0118 -17.500 -1.1837 0.05390 0.04937 -0.0655 1.0000 0.0119 -17.250 -1.1869 0.05033 0.04572 -0.0668 1.0000 0.0120 -17.000 -1.1872 0.04725 0.04255 -0.0677 1.0000 0.0121 -16.750 -1.1856 0.04454 0.03977 -0.0683 1.0000 0.0122 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0.00365 -0.0961 0.2615 0.6738 3.250 0.8377 0.01023 0.00376 -0.0963 0.2569 0.6759 3.500 0.8668 0.01033 0.00387 -0.0965 0.2536 0.6777 3.750 0.8959 0.01045 0.00398 -0.0967 0.2498 0.6793 4.000 0.9246 0.01058 0.00410 -0.0969 0.2454 0.6808 4.250 0.9530 0.01075 0.00424 -0.0970 0.2406 0.6824 4.500 0.9817 0.01088 0.00437 -0.0972 0.2372 0.6839 4.750 1.0103 0.01101 0.00450 -0.0973 0.2333 0.6857 5.000 1.0384 0.01118 0.00465 -0.0974 0.2288 0.6875 5.250 1.0663 0.01137 0.00481 -0.0974 0.2242 0.6892 5.500 1.0946 0.01151 0.00495 -0.0975 0.2210 0.6907 5.750 1.1225 0.01168 0.00511 -0.0975 0.2167 0.6921 6.000 1.1499 0.01189 0.00529 -0.0975 0.2120 0.6933 6.250 1.1774 0.01207 0.00546 -0.0975 0.2080 0.6945 6.500 1.2050 0.01224 0.00565 -0.0975 0.2039 0.6966 6.750 1.2319 0.01246 0.00585 -0.0974 0.1990 0.6986 7.000 1.2587 0.01267 0.00606 -0.0973 0.1948 0.7006 7.250 1.2856 0.01287 0.00627 -0.0972 0.1907 0.7025 7.500 1.3119 0.01311 0.00650 -0.0969 0.1858 0.7046 7.750 1.3380 0.01336 0.00675 -0.0967 0.1814 0.7069 8.000 1.3640 0.01359 0.00698 -0.0965 0.1768 0.7093 8.250 1.3892 0.01389 0.00726 -0.0961 0.1717 0.7115 8.500 1.4147 0.01414 0.00752 -0.0958 0.1673 0.7134 8.750 1.4395 0.01443 0.00779 -0.0954 0.1623 0.7151 9.000 1.4638 0.01475 0.00810 -0.0949 0.1577 0.7168 9.250 1.4882 0.01503 0.00840 -0.0944 0.1531 0.7194 9.750 1.5349 0.01571 0.00908 -0.0931 0.1437 0.7245 10.000 1.5569 0.01609 0.00946 -0.0923 0.1386 0.7271 10.250 1.5789 0.01645 0.00983 -0.0914 0.1344 0.7297 10.500 1.5998 0.01685 0.01023 -0.0904 0.1297 0.7323 10.750 1.6198 0.01726 0.01065 -0.0892 0.1256 0.7348 11.000 1.6390 0.01769 0.01108 -0.0879 0.1211 0.7370 11.250 1.6544 0.01816 0.01155 -0.0859 0.1174 0.7395 11.500 1.6710 0.01860 0.01203 -0.0841 0.1138 0.7425 11.750 1.6868 0.01918 0.01262 -0.0824 0.1097 0.7456 12.000 1.7036 0.01972 0.01319 -0.0809 0.1062 0.7490 12.250 1.7186 0.02038 0.01385 -0.0792 0.1019 0.7524 12.500 1.7336 0.02105 0.01454 -0.0775 0.0977 0.7556 13.000 1.7605 0.02262 0.01614 -0.0741 0.0892 0.7610 13.250 1.7723 0.02354 0.01708 -0.0724 0.0852 0.7642 13.500 1.7842 0.02448 0.01804 -0.0707 0.0810 0.7674 13.750 1.7950 0.02552 0.01911 -0.0691 0.0776 0.7708 14.000 1.8061 0.02658 0.02020 -0.0676 0.0746 0.7740 14.250 1.8157 0.02776 0.02142 -0.0660 0.0721 0.7770 14.500 1.8261 0.02893 0.02263 -0.0646 0.0697 0.7796 14.750 1.8341 0.03032 0.02407 -0.0632 0.0672 0.7828 15.000 1.8422 0.03173 0.02553 -0.0618 0.0650 0.7862 15.250 1.8492 0.03327 0.02713 -0.0605 0.0628 0.7896 15.500 1.8532 0.03512 0.02902 -0.0592 0.0606 0.7930 15.750 1.8591 0.03686 0.03083 -0.0581 0.0587 0.7962 16.000 1.8623 0.03887 0.03290 -0.0570 0.0569 0.7992 16.250 1.8631 0.04117 0.03527 -0.0560 0.0553 0.8026 16.500 1.8650 0.04347 0.03766 -0.0552 0.0538 0.8065 16.750 1.8642 0.04612 0.04039 -0.0546 0.0523 0.8105 17.000 1.8601 0.04924 0.04359 -0.0540 0.0509 0.8142 17.250 1.8560 0.05252 0.04697 -0.0538 0.0497 0.8176 17.500 1.8504 0.05614 0.05069 -0.0539 0.0485 0.8209 17.750 1.8403 0.06050 0.05517 -0.0544 0.0474 0.8245 18.000 1.8260 0.06565 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