XFOIL Version 6.94 Calculated polar for: EPPLER 395 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.7179 0.00867 0.00297 -0.1722 0.7226 0.5370 0.500 0.7746 0.00870 0.00311 -0.1724 0.7129 0.6005 1.000 0.8289 0.00864 0.00326 -0.1720 0.7024 0.6690 1.500 0.8836 0.00867 0.00345 -0.1717 0.6922 0.7446 2.000 0.9324 0.00858 0.00361 -0.1699 0.6806 0.8294 2.500 0.9749 0.00843 0.00365 -0.1665 0.6690 1.0000 3.000 1.0301 0.00858 0.00376 -0.1665 0.6556 1.0000 3.500 1.0838 0.00877 0.00396 -0.1661 0.6411 1.0000 4.000 1.1369 0.00900 0.00415 -0.1656 0.6255 1.0000 4.500 1.1885 0.00925 0.00438 -0.1648 0.6080 1.0000 5.000 1.2381 0.00952 0.00466 -0.1636 0.5886 1.0000 5.500 1.2854 0.00985 0.00499 -0.1619 0.5656 1.0000 6.000 1.3304 0.01022 0.00539 -0.1598 0.5386 1.0000 6.500 1.3705 0.01076 0.00588 -0.1567 0.5049 1.0000 7.000 1.4047 0.01145 0.00649 -0.1526 0.4634 1.0000 7.500 1.4254 0.01241 0.00727 -0.1460 0.4115 1.0000 8.000 1.4368 0.01382 0.00840 -0.1381 0.3484 1.0000 8.500 1.4459 0.01558 0.00988 -0.1304 0.2855 1.0000 9.000 1.4522 0.01772 0.01175 -0.1230 0.2259 1.0000 9.500 1.4569 0.02024 0.01399 -0.1161 0.1713 1.0000 10.000 1.4622 0.02306 0.01659 -0.1101 0.1255 1.0000 10.500 1.4686 0.02613 0.01949 -0.1048 0.0889 1.0000 11.000 1.4748 0.02952 0.02276 -0.1002 0.0598 1.0000 12.000 1.4866 0.03728 0.03045 -0.0928 0.0266 1.0000 12.500 1.4921 0.04167 0.03494 -0.0898 0.0187 1.0000 13.000 1.4934 0.04683 0.04021 -0.0873 0.0137 1.0000 13.500 1.4956 0.05229 0.04585 -0.0854 0.0105 1.0000 14.000 1.4944 0.05853 0.05228 -0.0842 0.0087 1.0000 14.500 1.4953 0.06491 0.05887 -0.0837 0.0075 1.0000 15.000 1.4849 0.07330 0.06748 -0.0840 0.0069 1.0000 15.500 1.4817 0.08118 0.07563 -0.0848 0.0063 1.0000 16.000 1.4762 0.08975 0.08444 -0.0864 0.0059 1.0000 16.500 1.4694 0.09891 0.09381 -0.0888 0.0056 1.0000 17.000 1.4601 0.10863 0.10374 -0.0918 0.0053 1.0000 17.500 1.4514 0.11837 0.11372 -0.0951 0.0051 1.0000 18.000 1.4461 0.12781 0.12345 -0.0989 0.0050 1.0000 18.500 1.4388 0.13767 0.13361 -0.1035 0.0049 1.0000 19.000 1.4294 0.14807 0.14432 -0.1090 0.0048 1.0000 19.500 1.4173 0.15925 0.15581 -0.1155 0.0048 1.0000 20.000 1.4020 0.17139 0.16826 -0.1232 0.0048 1.0000 20.500 1.3839 0.18461 0.18179 -0.1321 0.0048 1.0000 21.000 1.3623 0.19934 0.19683 -0.1424 0.0048 1.0000 21.500 1.3385 0.21544 0.21320 -0.1538 0.0050 1.0000 22.000 1.3092 0.23460 0.23260 -0.1669 0.0051 1.0000