XFOIL Version 6.94 Calculated polar for: EPPLER 407 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.3374 0.01851 0.01318 -0.1091 0.8499 0.6774 0.500 0.4185 0.01779 0.01243 -0.1141 0.8472 0.6838 1.000 0.5037 0.01681 0.01146 -0.1197 0.8455 0.6891 1.500 0.5887 0.01580 0.01053 -0.1250 0.8431 0.6937 2.500 0.7281 0.01429 0.00916 -0.1301 0.8254 0.7056 3.000 0.7806 0.01368 0.00866 -0.1291 0.8101 0.7103 3.500 0.8622 0.01276 0.00788 -0.1338 0.7980 0.7150 4.000 0.9157 0.01229 0.00752 -0.1329 0.7740 0.7214 4.500 0.9790 0.01192 0.00715 -0.1341 0.7404 0.7273 5.000 1.0234 0.01185 0.00706 -0.1314 0.6866 0.7322 5.500 1.0593 0.01239 0.00731 -0.1272 0.6077 0.7379 6.000 1.0776 0.01353 0.00803 -0.1200 0.5199 0.7451 6.500 1.0904 0.01489 0.00903 -0.1124 0.4333 0.7510 7.000 1.1042 0.01644 0.01024 -0.1054 0.3490 0.7570 7.500 1.1241 0.01813 0.01160 -0.1001 0.2686 0.7639 8.000 1.1465 0.01994 0.01308 -0.0955 0.1946 0.7709 9.000 1.1928 0.02400 0.01666 -0.0872 0.0866 0.7856 9.500 1.2143 0.02633 0.01890 -0.0832 0.0567 0.7927 10.000 1.2331 0.02891 0.02151 -0.0789 0.0412 0.8011 10.500 1.2520 0.03168 0.02437 -0.0751 0.0322 0.8104 11.000 1.2698 0.03449 0.02734 -0.0714 0.0259 0.8201 11.500 1.2841 0.03790 0.03092 -0.0680 0.0212 0.8304 12.000 1.2984 0.04134 0.03463 -0.0647 0.0170 0.8424 12.500 1.3080 0.04556 0.03910 -0.0617 0.0143 0.8564 13.000 1.3150 0.05013 0.04393 -0.0589 0.0122 0.8753 13.500 1.3167 0.05433 0.04848 -0.0554 0.0107 0.9180 15.000 1.3172 0.07520 0.07041 -0.0546 0.0085 1.0000 15.500 1.3091 0.08412 0.07962 -0.0565 0.0082 1.0000 16.000 1.2923 0.09502 0.09090 -0.0597 0.0080 1.0000 16.500 1.2663 0.10807 0.10439 -0.0655 0.0079 1.0000 17.000 1.2382 0.12236 0.11909 -0.0735 0.0079 1.0000 17.500 1.2103 0.13781 0.13494 -0.0835 0.0080 1.0000 18.000 1.1171 0.17430 0.17228 -0.1088 0.0091 1.0000 18.500 1.0650 0.20342 0.20148 -0.1262 0.0102 1.0000 19.000 0.8672 0.25758 0.25596 -0.1471 0.0220 0.8328 19.500 0.8707 0.26836 0.26676 -0.1515 0.0212 0.8381 20.000 0.8770 0.27863 0.27706 -0.1559 0.0193 0.8443 20.500 0.8905 0.28565 0.28413 -0.1578 0.0185 0.8525 21.000 0.8899 0.29836 0.29685 -0.1633 0.0174 0.8571 21.500 0.8975 0.30780 0.30632 -0.1668 0.0162 0.8653 22.000 0.9065 0.31464 0.31322 -0.1683 0.0155 0.8758 22.500 0.9084 0.32639 0.32497 -0.1728 0.0149 0.8840 23.000 0.9116 0.33491 0.33353 -0.1750 0.0138 0.8963 23.500 0.9126 0.34024 0.33889 -0.1755 0.0129 0.9161