XFOIL Version 6.96 Calculated polar for: NASA NLF1015 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 ------ -------- --------- --------- -------- -------- -------- -11.250 -0.2369 0.04034 0.03693 -0.1754 0.9076 0.0098 -11.000 -0.2412 0.03506 0.03126 -0.1820 0.8919 0.0097 -10.750 -0.2435 0.03220 0.02814 -0.1837 0.8776 0.0098 -10.500 -0.2518 0.02961 0.02525 -0.1829 0.8643 0.0098 -10.250 -0.2588 0.02790 0.02332 -0.1806 0.8527 0.0098 -10.000 -0.2606 0.02645 0.02167 -0.1784 0.8438 0.0099 -9.750 -0.2550 0.02489 0.01986 -0.1770 0.8366 0.0099 -9.500 -0.2458 0.02351 0.01827 -0.1756 0.8306 0.0100 -9.250 -0.2323 0.02241 0.01700 -0.1746 0.8249 0.0102 -9.000 -0.2164 0.02123 0.01560 -0.1738 0.8203 0.0103 -8.750 -0.1998 0.02020 0.01441 -0.1729 0.8156 0.0104 -8.500 -0.1812 0.01927 0.01333 -0.1722 0.8110 0.0106 -8.250 -0.1607 0.01841 0.01230 -0.1717 0.8069 0.0108 -8.000 -0.1399 0.01763 0.01139 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0.6866 1.750 0.9274 0.00899 0.00341 -0.1842 0.7078 0.6893 2.000 0.9551 0.00904 0.00346 -0.1844 0.7043 0.6923 2.250 0.9824 0.00911 0.00353 -0.1845 0.7007 0.6952 2.750 1.0330 0.00919 0.00370 -0.1838 0.6917 0.7003 3.000 1.0603 0.00926 0.00376 -0.1839 0.6876 0.7030 3.250 1.0842 0.00931 0.00388 -0.1833 0.6827 0.7061 3.500 1.1089 0.00937 0.00397 -0.1828 0.6773 0.7091 3.750 1.1343 0.00945 0.00406 -0.1825 0.6719 0.7121 4.000 1.1569 0.00950 0.00419 -0.1816 0.6649 0.7148 4.250 1.1803 0.00958 0.00428 -0.1808 0.6581 0.7177 4.500 1.2020 0.00966 0.00443 -0.1797 0.6500 0.7211 4.750 1.2228 0.00975 0.00456 -0.1784 0.6412 0.7249 5.000 1.2409 0.00986 0.00468 -0.1765 0.6300 0.7286 5.250 1.2552 0.01001 0.00484 -0.1738 0.6126 0.7317 5.500 1.2639 0.01032 0.00507 -0.1700 0.5829 0.7354 5.750 1.2667 0.01088 0.00547 -0.1652 0.5442 0.7397 6.250 1.2607 0.01272 0.00688 -0.1538 0.4549 0.7484 6.500 1.2592 0.01377 0.00774 -0.1488 0.4132 0.7534 6.750 1.2572 0.01496 0.00873 -0.1440 0.3703 0.7586 7.000 1.2508 0.01646 0.00996 -0.1387 0.3188 0.7634 7.250 1.2497 0.01785 0.01115 -0.1345 0.2748 0.7687 7.500 1.2483 0.01940 0.01245 -0.1306 0.2290 0.7742 7.750 1.2543 0.02059 0.01354 -0.1278 0.1989 0.7793 8.000 1.2612 0.02181 0.01465 -0.1253 0.1713 0.7853 8.500 1.2763 0.02429 0.01694 -0.1206 0.1210 0.7980 9.000 1.2874 0.02723 0.01964 -0.1159 0.0678 0.8123 9.500 1.2989 0.03036 0.02259 -0.1115 0.0259 0.8281 9.750 1.3051 0.03197 0.02414 -0.1095 0.0119 0.8367 10.000 1.3163 0.03317 0.02539 -0.1080 0.0098 0.8467 10.250 1.3281 0.03432 0.02663 -0.1066 0.0087 0.8580 10.500 1.3397 0.03547 0.02788 -0.1051 0.0081 0.8718 10.750 1.3503 0.03663 0.02915 -0.1035 0.0077 0.8909 11.000 1.3559 0.03765 0.03032 -0.1009 0.0073 0.9419 11.250 1.3654 0.03897 0.03172 -0.0995 0.0070 1.0000 11.500 1.3757 0.04053 0.03335 -0.0984 0.0067 1.0000 11.750 1.3849 0.04222 0.03511 -0.0973 0.0064 1.0000 12.000 1.3959 0.04377 0.03673 -0.0963 0.0063 1.0000 12.250 1.4059 0.04541 0.03846 -0.0954 0.0061 1.0000 12.500 1.4155 0.04712 0.04024 -0.0944 0.0059 1.0000 12.750 1.4242 0.04896 0.04217 -0.0935 0.0057 1.0000 13.000 1.4326 0.05086 0.04415 -0.0926 0.0056 1.0000 13.250 1.4400 0.05290 0.04627 -0.0918 0.0054 1.0000 13.500 1.4465 0.05507 0.04852 -0.0909 0.0053 1.0000 13.750 1.4521 0.05737 0.05091 -0.0901 0.0051 1.0000 14.000 1.4570 0.05979 0.05343 -0.0894 0.0050 1.0000 14.250 1.4614 0.06234 0.05606 -0.0887 0.0049 1.0000 14.500 1.4647 0.06507 0.05888 -0.0881 0.0049 1.0000 14.750 1.4673 0.06794 0.06185 -0.0876 0.0048 1.0000 15.000 1.4688 0.07099 0.06501 -0.0872 0.0047 1.0000 15.250 1.4689 0.07428 0.06841 -0.0869 0.0046 1.0000 15.500 1.4682 0.07779 0.07202 -0.0867 0.0045 1.0000 15.750 1.4660 0.08155 0.07589 -0.0866 0.0045 1.0000 16.000 1.4631 0.08550 0.07995 -0.0867 0.0044 1.0000