XFOIL Version 6.94 Calculated polar for: EPPLER 549 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 0.3244 0.01340 0.00672 -0.0671 0.6339 0.6586 0.500 0.3808 0.01340 0.00658 -0.0677 0.6233 0.6622 1.000 0.4354 0.01330 0.00645 -0.0680 0.6122 0.6659 1.500 0.4933 0.01340 0.00639 -0.0690 0.6027 0.6694 2.000 0.5463 0.01330 0.00637 -0.0689 0.5929 0.6731 3.000 0.6566 0.01351 0.00662 -0.0695 0.5755 0.6807 3.500 0.7110 0.01360 0.00677 -0.0698 0.5667 0.6850 4.000 0.7689 0.01386 0.00695 -0.0707 0.5582 0.6895 4.500 0.8192 0.01388 0.00714 -0.0701 0.5482 0.6937 5.000 0.8730 0.01400 0.00733 -0.0700 0.5384 0.6989 5.500 0.9232 0.01417 0.00763 -0.0694 0.5269 0.7045 6.000 0.9762 0.01430 0.00777 -0.0693 0.5149 0.7101 6.500 1.0223 0.01438 0.00806 -0.0678 0.5018 0.7157 7.000 1.0711 0.01456 0.00831 -0.0668 0.4882 0.7223 7.500 1.1133 0.01467 0.00856 -0.0646 0.4718 0.7298 8.000 1.1509 0.01481 0.00890 -0.0615 0.4534 0.7379 8.500 1.1837 0.01505 0.00928 -0.0575 0.4318 0.7477 9.000 1.2033 0.01535 0.00969 -0.0511 0.4069 0.7577 9.500 1.2158 0.01605 0.01042 -0.0440 0.3759 0.7706 10.000 1.2192 0.01716 0.01156 -0.0360 0.3372 0.7860 10.500 1.2132 0.01894 0.01330 -0.0277 0.2963 0.8069 11.000 1.2049 0.02118 0.01560 -0.0201 0.2661 0.8409 11.500 1.2096 0.02361 0.01827 -0.0159 0.2310 1.0000 12.000 1.2061 0.02721 0.02171 -0.0121 0.1949 1.0000 12.500 1.1995 0.03138 0.02579 -0.0087 0.1657 1.0000 13.000 1.1931 0.03596 0.03028 -0.0061 0.1395 1.0000 13.500 1.1893 0.04073 0.03501 -0.0043 0.1165 1.0000 14.000 1.1830 0.04612 0.04035 -0.0031 0.0975 1.0000 14.500 1.1811 0.05147 0.04572 -0.0025 0.0807 1.0000 15.000 1.1757 0.05751 0.05174 -0.0025 0.0679 1.0000 15.500 1.1741 0.06348 0.05779 -0.0030 0.0569 1.0000 16.000 1.1706 0.07000 0.06437 -0.0040 0.0483 1.0000 16.500 1.1658 0.07700 0.07143 -0.0055 0.0415 1.0000 17.000 1.1608 0.08432 0.07883 -0.0074 0.0359 1.0000 17.500 1.1576 0.09174 0.08638 -0.0098 0.0310 1.0000 18.000 1.1530 0.09958 0.09433 -0.0127 0.0271 1.0000 18.500 1.1492 0.10761 0.10252 -0.0160 0.0235 1.0000 19.000 1.1451 0.11587 0.11096 -0.0197 0.0205 1.0000 19.500 1.1403 0.12430 0.11956 -0.0238 0.0180 1.0000 20.000 1.1348 0.13293 0.12833 -0.0282 0.0159 1.0000 20.500 1.1306 0.14140 0.13692 -0.0329 0.0142 1.0000 21.000 1.1266 0.14979 0.14551 -0.0376 0.0128 1.0000 21.500 1.1251 0.15770 0.15352 -0.0424 0.0116 1.0000 22.500 1.1146 0.17519 0.17142 -0.0535 0.0098 1.0000 23.000 1.1130 0.18315 0.17945 -0.0589 0.0091 1.0000 24.000 1.0868 0.20479 0.20159 -0.0733 0.0084 1.0000 24.500 1.0573 0.22054 0.21766 -0.0832 0.0084 1.0000