XFOIL Version 6.94 Calculated polar for: EPPLER 542 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 0.3963 0.02025 0.01426 -0.0580 0.6927 0.8255 0.500 0.4532 0.01997 0.01390 -0.0589 0.6851 0.8310 1.000 0.5136 0.01975 0.01374 -0.0603 0.6766 0.8347 1.500 0.5589 0.01951 0.01348 -0.0590 0.6682 0.8398 2.000 0.5123 0.01913 0.01313 -0.0404 0.6598 0.8488 2.500 0.5873 0.01877 0.01280 -0.0447 0.6498 0.8506 3.000 0.6441 0.01854 0.01257 -0.0456 0.6407 0.8525 3.500 0.6723 0.01820 0.01233 -0.0410 0.6301 0.8548 4.000 0.6985 0.01792 0.01202 -0.0360 0.6204 0.8575 4.500 0.6953 0.01739 0.01162 -0.0254 0.6098 0.8613 5.000 0.7276 0.01695 0.01115 -0.0220 0.5988 0.8632 5.500 0.7533 0.01646 0.01080 -0.0173 0.5856 0.8654 6.000 0.7951 0.01622 0.01063 -0.0158 0.5716 0.8677 6.500 0.8375 0.01577 0.01022 -0.0138 0.5553 0.8689 7.000 0.8701 0.01538 0.00996 -0.0099 0.5356 0.8705 7.500 0.8997 0.01512 0.00980 -0.0054 0.5120 0.8723 8.000 0.9168 0.01491 0.00972 0.0014 0.4825 0.8747 8.500 0.9288 0.01507 0.00986 0.0089 0.4441 0.8782 9.000 0.9361 0.01578 0.01045 0.0163 0.3927 0.8809 9.500 0.9343 0.01713 0.01157 0.0242 0.3393 0.8839 10.000 0.9329 0.01892 0.01319 0.0309 0.2872 0.8865 10.500 0.9301 0.02098 0.01511 0.0371 0.2418 0.8894 11.000 0.9279 0.02331 0.01735 0.0427 0.2019 0.8928 11.500 0.9279 0.02592 0.01988 0.0473 0.1668 0.8971 12.000 0.9302 0.02880 0.02269 0.0507 0.1346 0.9005 12.500 0.9331 0.03198 0.02581 0.0535 0.1068 0.9040 13.000 0.9360 0.03550 0.02927 0.0557 0.0838 0.9072 13.500 0.9377 0.03920 0.03296 0.0577 0.0658 0.9114 14.000 0.9396 0.04319 0.03700 0.0592 0.0522 0.9161 14.500 0.9430 0.04742 0.04131 0.0600 0.0422 0.9212 15.000 0.9460 0.05212 0.04610 0.0601 0.0356 0.9265 15.500 0.9497 0.05719 0.05131 0.0594 0.0301 0.9333 16.000 0.9569 0.06255 0.05688 0.0577 0.0250 0.9432 16.500 0.9648 0.06904 0.06359 0.0540 0.0210 0.9564 18.500 0.9639 0.09880 0.09407 0.0391 0.0116 1.0000 19.500 0.9529 0.11631 0.11204 0.0298 0.0095 1.0000 20.000 0.9504 0.12461 0.12049 0.0250 0.0091 1.0000 20.500 0.9431 0.13376 0.12985 0.0198 0.0087 1.0000 21.000 0.9218 0.14638 0.14284 0.0121 0.0084 1.0000 21.500 0.8986 0.15978 0.15656 0.0039 0.0083 1.0000 22.000 0.8763 0.17334 0.17036 -0.0044 0.0085 1.0000 22.500 0.8404 0.19113 0.18838 -0.0145 0.0087 1.0000