XFOIL Version 6.94 Calculated polar for: EPPLER E662 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 0.5302 0.01338 0.00893 -0.1237 0.7865 0.8316 0.500 0.5714 0.01357 0.00913 -0.1206 0.7791 0.8455 1.000 0.6234 0.01331 0.00885 -0.1195 0.7737 0.8516 1.500 0.6898 0.01301 0.00847 -0.1217 0.7694 0.8561 2.000 0.7371 0.01292 0.00841 -0.1205 0.7609 0.8613 2.500 0.8001 0.01257 0.00805 -0.1225 0.7539 0.8645 3.000 0.8694 0.01205 0.00748 -0.1254 0.7483 0.8671 3.500 0.9068 0.01173 0.00727 -0.1219 0.7367 0.8703 4.000 0.9709 0.01122 0.00674 -0.1237 0.7282 0.8732 4.500 1.0076 0.01094 0.00657 -0.1201 0.7126 0.8766 5.000 1.0485 0.01068 0.00637 -0.1173 0.6940 0.8805 5.500 1.0883 0.01048 0.00612 -0.1143 0.6691 0.8835 6.000 1.1188 0.01067 0.00625 -0.1096 0.6338 0.8863 6.500 1.1400 0.01111 0.00655 -0.1032 0.5920 0.8906 7.000 1.1526 0.01196 0.00723 -0.0957 0.5471 0.8951 7.500 1.1604 0.01317 0.00828 -0.0879 0.4980 0.8998 8.000 1.1639 0.01487 0.00975 -0.0802 0.4449 0.9045 8.500 1.1723 0.01661 0.01132 -0.0738 0.3951 0.9092 9.000 1.1818 0.01842 0.01298 -0.0678 0.3477 0.9154 9.500 1.1920 0.02053 0.01487 -0.0625 0.2972 0.9211 10.000 1.2078 0.02262 0.01679 -0.0584 0.2488 0.9265 10.500 1.2199 0.02485 0.01881 -0.0539 0.2020 0.9345 11.000 1.2361 0.02708 0.02090 -0.0503 0.1614 0.9442 11.500 1.2559 0.02953 0.02324 -0.0478 0.1258 0.9632 12.000 1.2781 0.03213 0.02573 -0.0462 0.0969 1.0000 12.500 1.2990 0.03503 0.02856 -0.0446 0.0735 1.0000 13.000 1.3187 0.03811 0.03159 -0.0430 0.0546 1.0000 13.500 1.3357 0.04152 0.03497 -0.0415 0.0389 1.0000 14.000 1.3480 0.04549 0.03891 -0.0400 0.0241 1.0000 14.500 1.3554 0.05015 0.04359 -0.0385 0.0126 1.0000 15.000 1.3594 0.05542 0.04898 -0.0373 0.0071 1.0000 15.500 1.3659 0.06072 0.05446 -0.0368 0.0059 1.0000 16.000 1.3714 0.06640 0.06036 -0.0368 0.0054 1.0000 16.500 1.3705 0.07319 0.06737 -0.0373 0.0049 1.0000 17.000 1.3637 0.08118 0.07561 -0.0386 0.0047 1.0000 17.500 1.3557 0.08976 0.08444 -0.0408 0.0046 1.0000 18.000 1.3454 0.09913 0.09408 -0.0438 0.0045 1.0000 18.500 1.3328 0.10918 0.10439 -0.0478 0.0044 1.0000 19.000 1.3188 0.11976 0.11523 -0.0525 0.0044 1.0000 19.500 1.3051 0.13042 0.12613 -0.0578 0.0043 1.0000 20.000 1.2927 0.14089 0.13683 -0.0634 0.0043 1.0000 20.500 1.2815 0.15116 0.14730 -0.0693 0.0043 1.0000 21.000 1.2729 0.16090 0.15723 -0.0751 0.0043 1.0000 21.500 1.2658 0.17023 0.16674 -0.0810 0.0043 1.0000 22.000 1.2598 0.17936 0.17605 -0.0869 0.0043 1.0000 22.500 1.2537 0.18854 0.18542 -0.0930 0.0043 1.0000 23.000 1.2465 0.19803 0.19510 -0.0995 0.0043 1.0000 23.500 1.2373 0.20812 0.20538 -0.1065 0.0043 1.0000 24.000 1.2251 0.21914 0.21662 -0.1141 0.0044 1.0000 24.500 1.2093 0.23146 0.22915 -0.1225 0.0045 1.0000 25.000 1.1861 0.24653 0.24442 -0.1322 0.0046 1.0000