XFOIL Version 6.94 Calculated polar for: EPPLER 66 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 0.5839 0.00680 0.00249 -0.1432 0.8370 0.7839 0.500 0.6334 0.00670 0.00250 -0.1412 0.8183 0.8555 1.500 0.7343 0.00660 0.00241 -0.1378 0.7746 1.0000 2.000 0.7881 0.00676 0.00246 -0.1373 0.7487 1.0000 2.500 0.8409 0.00698 0.00259 -0.1366 0.7205 1.0000 3.000 0.8919 0.00725 0.00275 -0.1355 0.6888 1.0000 3.500 0.9409 0.00759 0.00297 -0.1340 0.6520 1.0000 4.000 0.9878 0.00800 0.00325 -0.1321 0.6107 1.0000 4.500 1.0289 0.00860 0.00361 -0.1290 0.5486 1.0000 5.000 1.0585 0.00969 0.00414 -0.1238 0.4372 1.0000 5.500 1.0943 0.01070 0.00479 -0.1202 0.3584 1.0000 6.000 1.1290 0.01184 0.00554 -0.1164 0.2784 1.0000 6.500 1.1510 0.01367 0.00662 -0.1108 0.1586 1.0000 7.000 1.1805 0.01507 0.00766 -0.1064 0.0963 1.0000 7.500 1.2112 0.01636 0.00873 -0.1022 0.0573 1.0000 8.000 1.2432 0.01759 0.00988 -0.0984 0.0342 1.0000 8.500 1.2712 0.01906 0.01125 -0.0941 0.0143 1.0000 9.000 1.3021 0.02035 0.01260 -0.0903 0.0095 1.0000 9.500 1.3327 0.02165 0.01403 -0.0866 0.0072 1.0000 10.000 1.3584 0.02330 0.01578 -0.0825 0.0025 1.0000 10.500 1.3816 0.02519 0.01782 -0.0783 0.0017 1.0000 11.000 1.4031 0.02728 0.02013 -0.0742 0.0016 1.0000 11.500 1.4209 0.02975 0.02284 -0.0701 0.0015 1.0000 12.000 1.4349 0.03267 0.02601 -0.0660 0.0014 1.0000 12.500 1.4448 0.03613 0.02975 -0.0623 0.0014 1.0000 13.000 1.4482 0.04044 0.03437 -0.0588 0.0014 1.0000 13.500 1.4471 0.04555 0.03979 -0.0560 0.0014 1.0000 14.000 1.4407 0.05177 0.04633 -0.0543 0.0014 1.0000 14.500 1.4290 0.05932 0.05423 -0.0538 0.0014 1.0000 15.000 1.4123 0.06847 0.06372 -0.0550 0.0014 1.0000 15.500 1.3923 0.07916 0.07477 -0.0581 0.0014 1.0000 16.000 1.3688 0.09162 0.08758 -0.0631 0.0014 1.0000 16.500 1.3417 0.10593 0.10225 -0.0700 0.0014 1.0000 17.000 1.3137 0.12146 0.11811 -0.0785 0.0014 1.0000 17.500 1.2862 0.13770 0.13466 -0.0881 0.0014 1.0000 18.000 1.2595 0.15436 0.15161 -0.0983 0.0015 1.0000 18.500 1.2349 0.17113 0.16862 -0.1088 0.0015 1.0000 19.000 1.2114 0.18831 0.18602 -0.1196 0.0015 1.0000 19.500 1.1860 0.20722 0.20511 -0.1311 0.0016 1.0000