XFOIL Version 6.94 Calculated polar for: EPPLER 584 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 ------- -------- --------- --------- -------- ------- ------- 0.500 0.6354 0.01651 0.01063 -0.1032 0.6636 0.9203 1.000 0.6910 0.01651 0.01053 -0.1039 0.6576 0.9279 1.500 0.7350 0.01679 0.01070 -0.1023 0.6515 0.9349 2.000 0.7914 0.01666 0.01064 -0.1033 0.6450 0.9394 2.500 0.8371 0.01667 0.01066 -0.1021 0.6379 0.9444 3.000 0.8793 0.01680 0.01070 -0.1000 0.6316 0.9485 3.500 0.8987 0.01688 0.01090 -0.0936 0.6240 0.9528 4.000 0.9470 0.01670 0.01076 -0.0929 0.6159 0.9549 4.500 1.0075 0.01656 0.01055 -0.0945 0.6092 0.9564 5.000 1.0248 0.01645 0.01064 -0.0876 0.5995 0.9595 5.500 1.0651 0.01611 0.01032 -0.0851 0.5914 0.9621 6.000 1.0784 0.01594 0.01024 -0.0772 0.5828 0.9649 6.500 1.0709 0.01554 0.00991 -0.0649 0.5732 0.9696 7.000 1.1132 0.01521 0.00963 -0.0630 0.5631 0.9715 7.500 1.1397 0.01502 0.00955 -0.0581 0.5503 0.9748 8.000 1.1731 0.01504 0.00966 -0.0547 0.5363 0.9788 8.500 1.2106 0.01506 0.00970 -0.0522 0.5211 0.9826 9.000 1.2328 0.01537 0.01013 -0.0474 0.5027 0.9880 9.500 1.2690 0.01595 0.01080 -0.0458 0.4806 0.9932 10.000 1.2863 0.01678 0.01167 -0.0412 0.4575 1.0000 10.500 1.2863 0.01791 0.01281 -0.0338 0.4345 1.0000 11.000 1.2957 0.01972 0.01464 -0.0293 0.4078 1.0000 11.500 1.3026 0.02210 0.01704 -0.0253 0.3786 1.0000 12.000 1.3058 0.02502 0.01993 -0.0215 0.3481 1.0000 12.500 1.3066 0.02841 0.02326 -0.0180 0.3183 1.0000 13.000 1.3071 0.03216 0.02699 -0.0151 0.2875 1.0000 13.500 1.3035 0.03649 0.03123 -0.0124 0.2576 1.0000 14.000 1.3029 0.04086 0.03559 -0.0104 0.2288 1.0000 14.500 1.3014 0.04555 0.04024 -0.0087 0.2010 1.0000 15.000 1.2976 0.05072 0.04536 -0.0073 0.1757 1.0000 15.500 1.2968 0.05595 0.05060 -0.0066 0.1515 1.0000 16.000 1.2929 0.06179 0.05643 -0.0063 0.1289 1.0000 16.500 1.2873 0.06815 0.06279 -0.0064 0.1088 1.0000 17.000 1.2816 0.07486 0.06951 -0.0071 0.0896 1.0000 17.500 1.2741 0.08211 0.07678 -0.0084 0.0739 1.0000 18.000 1.2647 0.08994 0.08465 -0.0102 0.0603 1.0000 18.500 1.2566 0.09787 0.09268 -0.0124 0.0495 1.0000 19.000 1.2465 0.10637 0.10126 -0.0153 0.0411 1.0000 19.500 1.2360 0.11510 0.11007 -0.0186 0.0347 1.0000 20.000 1.2283 0.12355 0.11868 -0.0222 0.0292 1.0000 20.500 1.2208 0.13207 0.12734 -0.0263 0.0247 1.0000 21.000 1.2134 0.14066 0.13607 -0.0307 0.0205 1.0000 21.500 1.2053 0.14937 0.14493 -0.0355 0.0174 1.0000 22.000 1.1972 0.15813 0.15383 -0.0405 0.0146 1.0000 22.500 1.1894 0.16675 0.16254 -0.0457 0.0127 1.0000 23.000 1.1832 0.17517 0.17112 -0.0510 0.0109 1.0000 24.000 1.1767 0.19052 0.18666 -0.0612 0.0083 1.0000 24.500 1.1711 0.19893 0.19531 -0.0668 0.0073 1.0000 25.000 1.1688 0.20647 0.20294 -0.0722 0.0065 1.0000 25.500 1.1667 0.21382 0.21043 -0.0774 0.0060 1.0000 26.000 1.1561 0.22351 0.22039 -0.0840 0.0056 1.0000 26.500 1.1422 0.23424 0.23136 -0.0912 0.0054 1.0000 27.000 1.1146 0.24937 0.24678 -0.1001 0.0056 1.0000 29.000 1.0623 0.31531 0.31261 -0.1244 0.0130 1.0000 29.500 1.0635 0.32435 0.32159 -0.1303 0.0119 1.0000 30.000 1.0733 0.32903 0.32627 -0.1342 0.0111 1.0000 30.500 1.0840 0.33267 0.32994 -0.1378 0.0107 1.0000 31.000 1.0932 0.33770 0.33497 -0.1416 0.0104 1.0000 31.500 1.1002 0.34420 0.34145 -0.1463 0.0095 1.0000 32.000 1.1092 0.34811 0.34538 -0.1502 0.0086 1.0000 33.000 1.1257 0.35771 0.35498 -0.1584 0.0074 1.0000 33.500 1.1338 0.36180 0.35909 -0.1624 0.0068 1.0000 34.000 1.1418 0.36486 0.36217 -0.1663 0.0064 1.0000 35.000 1.1565 0.37447 0.37178 -0.1746 0.0057 1.0000 35.500 1.1634 0.37833 0.37565 -0.1787 0.0051 1.0000 36.000 1.1701 0.38160 0.37894 -0.1828 0.0048 1.0000 37.000 1.1825 0.38760 0.38499 -0.1906 0.0045 1.0000 37.500 1.1881 0.39249 0.38987 -0.1952 0.0043 1.0000 38.000 1.1939 0.39635 0.39374 -0.1994 0.0042 1.0000 38.500 1.1994 0.39976 0.39717 -0.2035 0.0037 1.0000 39.000 1.2042 0.40256 0.39999 -0.2078 0.0035 1.0000 39.500 1.2086 0.40498 0.40243 -0.2120 0.0033 1.0000