XFOIL Version 6.94 Calculated polar for: EPPLER 546 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 0.2387 0.01090 0.00527 -0.0432 0.6363 0.7333 0.500 0.2940 0.01084 0.00511 -0.0435 0.6231 0.7356 1.000 0.3493 0.01078 0.00500 -0.0439 0.6104 0.7380 1.500 0.4052 0.01084 0.00493 -0.0445 0.5977 0.7409 2.000 0.4603 0.01079 0.00488 -0.0448 0.5856 0.7432 2.500 0.5159 0.01088 0.00490 -0.0454 0.5734 0.7450 3.000 0.5696 0.01077 0.00480 -0.0454 0.5602 0.7473 3.500 0.6232 0.01081 0.00487 -0.0454 0.5471 0.7494 4.000 0.6766 0.01093 0.00498 -0.0453 0.5338 0.7517 4.500 0.7289 0.01099 0.00514 -0.0450 0.5200 0.7547 5.000 0.7809 0.01113 0.00531 -0.0447 0.5052 0.7574 5.500 0.8320 0.01131 0.00548 -0.0442 0.4891 0.7602 6.000 0.8819 0.01142 0.00568 -0.0435 0.4719 0.7630 6.500 0.9302 0.01162 0.00594 -0.0425 0.4516 0.7654 7.000 0.9743 0.01180 0.00617 -0.0408 0.4272 0.7686 7.500 1.0154 0.01211 0.00654 -0.0384 0.3977 0.7718 8.000 1.0480 0.01264 0.00702 -0.0345 0.3557 0.7758 8.500 1.0643 0.01346 0.00767 -0.0277 0.3009 0.7805 9.000 1.0753 0.01472 0.00871 -0.0205 0.2467 0.7847 9.500 1.0848 0.01609 0.00993 -0.0137 0.2009 0.7895 10.000 1.0920 0.01768 0.01144 -0.0071 0.1630 0.7949 10.500 1.0972 0.01961 0.01327 -0.0010 0.1304 0.8010 11.000 1.1017 0.02187 0.01547 0.0043 0.1032 0.8072 11.500 1.1045 0.02441 0.01803 0.0090 0.0805 0.8149 12.000 1.1072 0.02737 0.02099 0.0130 0.0618 0.8239 12.500 1.1086 0.03064 0.02433 0.0164 0.0480 0.8346 14.000 1.1164 0.04248 0.03680 0.0215 0.0261 0.9350 15.000 1.1252 0.05274 0.04730 0.0201 0.0180 1.0000 15.500 1.1244 0.05861 0.05330 0.0192 0.0149 1.0000 16.000 1.1190 0.06540 0.06024 0.0178 0.0123 1.0000 16.500 1.1117 0.07284 0.06785 0.0156 0.0106 1.0000 17.000 1.1055 0.08057 0.07575 0.0130 0.0092 1.0000 17.500 1.0946 0.08937 0.08476 0.0095 0.0082 1.0000 18.000 1.0854 0.09838 0.09397 0.0055 0.0074 1.0000 18.500 1.0725 0.10827 0.10401 0.0008 0.0069 1.0000 19.000 1.0607 0.11824 0.11419 -0.0042 0.0065 1.0000 19.500 1.0520 0.12800 0.12418 -0.0095 0.0061 1.0000 20.000 1.0441 0.13771 0.13409 -0.0149 0.0058 1.0000 20.500 1.0362 0.14745 0.14401 -0.0207 0.0056 1.0000 21.000 1.0287 0.15723 0.15396 -0.0267 0.0054 1.0000 21.500 1.0211 0.16704 0.16392 -0.0328 0.0052 1.0000 22.000 1.0122 0.17714 0.17417 -0.0392 0.0050 1.0000 22.500 0.9983 0.18862 0.18586 -0.0465 0.0050 1.0000 23.000 0.9800 0.20192 0.19940 -0.0547 0.0050 1.0000