XFOIL Version 6.94 Calculated polar for: EPPLER 545 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 0.3089 0.01080 0.00484 -0.0659 0.6372 0.6714 0.500 0.3646 0.01076 0.00477 -0.0663 0.6275 0.6744 1.000 0.4213 0.01079 0.00470 -0.0669 0.6178 0.6771 1.500 0.4768 0.01074 0.00465 -0.0673 0.6084 0.6797 2.000 0.5331 0.01082 0.00464 -0.0679 0.5985 0.6819 2.500 0.5870 0.01071 0.00457 -0.0680 0.5869 0.6844 3.000 0.6409 0.01075 0.00459 -0.0679 0.5745 0.6874 3.500 0.6934 0.01074 0.00467 -0.0676 0.5618 0.6904 4.000 0.7471 0.01086 0.00481 -0.0676 0.5512 0.6933 4.500 0.7997 0.01092 0.00495 -0.0673 0.5392 0.6965 5.000 0.8518 0.01106 0.00513 -0.0670 0.5269 0.6998 5.500 0.9020 0.01122 0.00531 -0.0663 0.5117 0.7027 6.000 0.9508 0.01128 0.00550 -0.0654 0.4955 0.7062 6.500 0.9936 0.01148 0.00574 -0.0632 0.4696 0.7099 7.000 1.0317 0.01180 0.00609 -0.0601 0.4343 0.7144 7.500 1.0531 0.01242 0.00655 -0.0540 0.3834 0.7192 8.000 1.0635 0.01347 0.00735 -0.0462 0.3249 0.7233 8.500 1.0733 0.01469 0.00843 -0.0388 0.2779 0.7285 9.000 1.0876 0.01592 0.00961 -0.0328 0.2441 0.7339 9.500 1.1006 0.01736 0.01099 -0.0270 0.2183 0.7399 10.000 1.1143 0.01894 0.01258 -0.0220 0.1988 0.7460 11.000 1.1310 0.02323 0.01675 -0.0123 0.1261 0.7620 11.500 1.1339 0.02603 0.01962 -0.0078 0.1080 0.7715 12.000 1.1402 0.02894 0.02258 -0.0044 0.0893 0.7826 12.500 1.1457 0.03209 0.02577 -0.0013 0.0704 0.7959 13.000 1.1489 0.03564 0.02940 0.0014 0.0559 0.8142 13.500 1.1513 0.03938 0.03329 0.0038 0.0447 0.8458 14.000 1.1626 0.04337 0.03762 0.0037 0.0355 1.0000 14.500 1.1660 0.04817 0.04247 0.0042 0.0296 1.0000 15.000 1.1702 0.05314 0.04754 0.0043 0.0248 1.0000 15.500 1.1729 0.05854 0.05305 0.0040 0.0207 1.0000 16.000 1.1735 0.06446 0.05910 0.0031 0.0174 1.0000 16.500 1.1698 0.07125 0.06603 0.0018 0.0147 1.0000 17.000 1.1614 0.07909 0.07403 -0.0004 0.0127 1.0000 17.500 1.1555 0.08700 0.08212 -0.0031 0.0112 1.0000 18.000 1.1456 0.09591 0.09124 -0.0065 0.0097 1.0000 18.500 1.1359 0.10514 0.10064 -0.0107 0.0088 1.0000 19.000 1.1243 0.11494 0.11064 -0.0153 0.0082 1.0000 19.500 1.1170 0.12424 0.12015 -0.0201 0.0075 1.0000 20.000 1.1083 0.13390 0.12998 -0.0254 0.0070 1.0000 20.500 1.1000 0.14337 0.13958 -0.0308 0.0067 1.0000 21.000 1.0935 0.15243 0.14880 -0.0361 0.0064 1.0000 21.500 1.0867 0.16186 0.15845 -0.0419 0.0061 1.0000 22.000 1.0791 0.17145 0.16824 -0.0480 0.0059 1.0000 22.500 1.0696 0.18149 0.17850 -0.0544 0.0057 1.0000 23.000 1.0581 0.19213 0.18934 -0.0614 0.0056 1.0000 23.500 1.0409 0.20451 0.20194 -0.0693 0.0055 1.0000 24.000 0.7634 0.29384 0.29244 -0.1198 0.0123 1.0000 24.500 0.7728 0.30238 0.30099 -0.1238 0.0113 1.0000 25.000 0.7804 0.31233 0.31094 -0.1281 0.0104 1.0000 25.500 0.7885 0.32193 0.32055 -0.1322 0.0092 1.0000 26.000 0.7983 0.32923 0.32786 -0.1357 0.0087 1.0000 26.500 0.8057 0.33910 0.33774 -0.1396 0.0085 1.0000 27.000 0.8114 0.35120 0.34983 -0.1440 0.0072 1.0000 27.500 0.8193 0.35940 0.35805 -0.1476 0.0067 1.0000 28.000 0.8281 0.36702 0.36569 -0.1508 0.0065 1.0000