XFOIL Version 6.94 Calculated polar for: EPPLER 582 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 0.5148 0.01086 0.00600 -0.1052 0.7078 0.8584 0.500 0.5632 0.01083 0.00587 -0.1034 0.7004 0.8655 1.000 0.6157 0.01084 0.00585 -0.1031 0.6924 0.8699 1.500 0.6713 0.01077 0.00572 -0.1036 0.6836 0.8737 2.000 0.7324 0.01085 0.00568 -0.1054 0.6749 0.8764 2.500 0.7851 0.01073 0.00561 -0.1054 0.6658 0.8785 3.000 0.8384 0.01061 0.00545 -0.1053 0.6558 0.8808 3.500 0.8901 0.01058 0.00547 -0.1051 0.6449 0.8826 4.000 0.9425 0.01055 0.00541 -0.1048 0.6331 0.8851 4.500 0.9924 0.01057 0.00551 -0.1042 0.6200 0.8872 5.000 1.0441 0.01064 0.00556 -0.1040 0.6058 0.8892 5.500 1.0911 0.01071 0.00567 -0.1028 0.5888 0.8914 6.000 1.1357 0.01086 0.00588 -0.1012 0.5697 0.8942 6.500 1.1767 0.01108 0.00613 -0.0989 0.5479 0.8964 7.000 1.2061 0.01130 0.00636 -0.0942 0.5233 0.8995 7.500 1.2306 0.01178 0.00682 -0.0887 0.4945 0.9033 8.000 1.2529 0.01245 0.00748 -0.0832 0.4599 0.9078 8.500 1.2700 0.01344 0.00839 -0.0771 0.4231 0.9128 9.000 1.2789 0.01473 0.00959 -0.0701 0.3845 0.9197 9.500 1.2880 0.01625 0.01106 -0.0637 0.3457 0.9278 10.000 1.2942 0.01812 0.01288 -0.0574 0.3072 0.9391 10.500 1.3041 0.02037 0.01507 -0.0528 0.2679 0.9679 11.000 1.3156 0.02313 0.01771 -0.0495 0.2290 1.0000 11.500 1.3276 0.02608 0.02055 -0.0465 0.1945 1.0000 12.000 1.3393 0.02919 0.02358 -0.0438 0.1623 1.0000 12.500 1.3458 0.03282 0.02708 -0.0410 0.1295 1.0000 13.000 1.3493 0.03690 0.03101 -0.0384 0.0990 1.0000 13.500 1.3535 0.04117 0.03517 -0.0363 0.0719 1.0000 14.000 1.3568 0.04581 0.03973 -0.0346 0.0505 1.0000 14.500 1.3608 0.05064 0.04457 -0.0334 0.0367 1.0000 15.000 1.3654 0.05573 0.04972 -0.0327 0.0269 1.0000 15.500 1.3681 0.06130 0.05539 -0.0325 0.0198 1.0000 16.000 1.3671 0.06766 0.06185 -0.0327 0.0130 1.0000 16.500 1.3542 0.07597 0.07028 -0.0337 0.0070 1.0000 17.000 1.3424 0.08474 0.07927 -0.0355 0.0045 1.0000 17.500 1.3256 0.09475 0.08953 -0.0384 0.0037 1.0000 18.000 1.3158 0.10411 0.09916 -0.0417 0.0035 1.0000 18.500 1.3019 0.11449 0.10981 -0.0460 0.0031 1.0000 19.000 1.2878 0.12513 0.12069 -0.0509 0.0029 1.0000 19.500 1.2732 0.13603 0.13183 -0.0565 0.0027 1.0000 20.000 1.2601 0.14670 0.14272 -0.0623 0.0027 1.0000 20.500 1.2471 0.15738 0.15359 -0.0684 0.0026 1.0000 21.000 1.2369 0.16745 0.16384 -0.0745 0.0025 1.0000 21.500 1.2300 0.17674 0.17329 -0.0803 0.0024 1.0000 22.000 1.2240 0.18570 0.18242 -0.0859 0.0024 1.0000 22.500 1.2180 0.19502 0.19193 -0.0921 0.0024 1.0000 23.000 1.2112 0.20465 0.20176 -0.0986 0.0024 1.0000 23.500 1.2019 0.21511 0.21243 -0.1057 0.0024 1.0000 24.000 1.1914 0.22629 0.22382 -0.1133 0.0024 1.0000 24.500 1.1830 0.23740 0.23511 -0.1208 0.0024 1.0000