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 = 1.000 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -15.500 -0.0628 0.12874 0.12689 -0.1138 0.9436 0.0067 -15.250 -0.0515 0.12531 0.12343 -0.1174 0.9411 0.0068 -15.000 -0.1974 0.13909 0.13717 -0.1013 0.9755 0.0065 -14.500 -0.1938 0.13225 0.13035 -0.1025 0.9641 0.0065 -14.250 -0.1903 0.12845 0.12655 -0.1038 0.9578 0.0065 -14.000 -0.1857 0.12341 0.12149 -0.1066 0.9525 0.0065 -13.000 -0.3581 0.05066 0.04828 -0.1555 0.9333 0.0069 -12.750 -0.3572 0.03891 0.03609 -0.1749 0.9228 0.0068 -12.500 -0.3436 0.03347 0.03030 -0.1853 0.9057 0.0069 -12.250 -0.3401 0.03052 0.02708 -0.1884 0.8860 0.0069 -12.000 -0.3467 0.02793 0.02423 -0.1884 0.8685 0.0069 -11.750 -0.3403 0.02685 0.02304 -0.1878 0.8562 0.0070 -11.500 -0.3398 0.02542 0.02144 -0.1865 0.8448 0.0071 -11.250 -0.3373 0.02428 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0.6469 0.6744 3.500 1.1201 0.00729 0.00266 -0.1850 0.6394 0.6773 3.750 1.1422 0.00739 0.00275 -0.1839 0.6301 0.6801 4.250 1.1818 0.00766 0.00299 -0.1809 0.6018 0.6852 4.500 1.1933 0.00801 0.00322 -0.1777 0.5710 0.6881 4.750 1.1949 0.00873 0.00369 -0.1725 0.5168 0.6912 5.000 1.1979 0.00952 0.00423 -0.1678 0.4678 0.6946 5.500 1.2072 0.01112 0.00540 -0.1594 0.3771 0.7012 5.750 1.2134 0.01191 0.00602 -0.1557 0.3372 0.7047 6.000 1.2182 0.01281 0.00672 -0.1519 0.2963 0.7083 6.250 1.2228 0.01376 0.00748 -0.1481 0.2569 0.7119 6.500 1.2272 0.01477 0.00831 -0.1445 0.2173 0.7153 6.750 1.2307 0.01590 0.00924 -0.1408 0.1771 0.7191 7.000 1.2367 0.01696 0.01015 -0.1377 0.1452 0.7233 7.250 1.2454 0.01794 0.01102 -0.1351 0.1196 0.7276 7.500 1.2535 0.01900 0.01196 -0.1325 0.0937 0.7318 7.750 1.2631 0.02003 0.01291 -0.1302 0.0723 0.7365 8.000 1.2702 0.02125 0.01401 -0.1276 0.0484 0.7412 8.250 1.2784 0.02245 0.01511 -0.1253 0.0294 0.7456 8.500 1.2860 0.02374 0.01633 -0.1230 0.0119 0.7504 8.750 1.2988 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