XFOIL Version 6.96 Calculated polar for: EPPLER 857 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -18.750 -0.7443 0.11360 0.11049 -0.0577 1.0000 0.0059 -18.500 -0.7631 0.10696 0.10375 -0.0608 1.0000 0.0058 -18.250 -0.7845 0.10001 0.09670 -0.0641 1.0000 0.0059 -18.000 -0.7986 0.09328 0.08985 -0.0684 0.9996 0.0058 -17.750 -0.8094 0.08630 0.08275 -0.0740 0.9987 0.0059 -17.500 -0.8153 0.08046 0.07681 -0.0786 0.9977 0.0060 -17.250 -0.8201 0.07497 0.07122 -0.0831 0.9966 0.0060 -17.000 -0.8202 0.07020 0.06635 -0.0873 0.9955 0.0059 -16.750 -0.8185 0.06562 0.06168 -0.0917 0.9945 0.0059 -16.500 -0.8148 0.06138 0.05738 -0.0959 0.9936 0.0063 -16.250 -0.8140 0.05746 0.05337 -0.0991 0.9927 0.0062 -16.000 -0.8148 0.05400 0.04983 -0.1010 0.9912 0.0062 -15.750 -0.8126 0.05067 0.04644 -0.1034 0.9897 0.0062 -15.500 -0.8071 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0.9954 0.01059 0.00405 -0.1055 0.3664 0.4882 4.750 1.0196 0.01069 0.00418 -0.1049 0.3643 0.4935 5.000 1.0433 0.01082 0.00431 -0.1042 0.3619 0.4983 5.500 1.0884 0.01111 0.00463 -0.1025 0.3572 0.5080 5.750 1.1111 0.01125 0.00480 -0.1017 0.3552 0.5141 6.000 1.1351 0.01136 0.00494 -0.1011 0.3539 0.5203 6.250 1.1588 0.01147 0.00510 -0.1005 0.3524 0.5267 6.500 1.1821 0.01159 0.00526 -0.0998 0.3503 0.5344 6.750 1.2049 0.01174 0.00543 -0.0991 0.3482 0.5410 7.000 1.2269 0.01189 0.00563 -0.0982 0.3460 0.5488 7.250 1.2481 0.01206 0.00583 -0.0972 0.3438 0.5575 7.500 1.2690 0.01225 0.00606 -0.0962 0.3418 0.5670 7.750 1.2892 0.01246 0.00630 -0.0951 0.3396 0.5778 8.000 1.3111 0.01262 0.00652 -0.0942 0.3380 0.5894 8.500 1.3541 0.01294 0.00698 -0.0925 0.3343 0.6146 8.750 1.3748 0.01313 0.00723 -0.0915 0.3322 0.6294 9.000 1.3949 0.01333 0.00750 -0.0904 0.3300 0.6465 9.250 1.4132 0.01357 0.00782 -0.0891 0.3273 0.6693 9.500 1.4304 0.01385 0.00817 -0.0876 0.3247 0.6956 9.750 1.4483 0.01408 0.00851 -0.0862 0.3220 0.7289 10.000 1.4666 0.01424 0.00883 -0.0849 0.3186 0.7754 10.250 1.4943 0.01422 0.00922 -0.0856 0.3141 1.0000 10.500 1.5102 0.01465 0.00964 -0.0841 0.3097 1.0000 10.750 1.5264 0.01508 0.01008 -0.0827 0.3058 1.0000 11.000 1.5448 0.01544 0.01047 -0.0816 0.3021 1.0000 11.250 1.5611 0.01589 0.01094 -0.0803 0.2981 1.0000 11.500 1.5749 0.01645 0.01150 -0.0787 0.2939 1.0000 11.750 1.5894 0.01701 0.01207 -0.0772 0.2902 1.0000 12.000 1.6051 0.01753 0.01262 -0.0760 0.2855 1.0000 12.250 1.6190 0.01815 0.01327 -0.0745 0.2816 1.0000 12.500 1.6284 0.01901 0.01412 -0.0726 0.2760 1.0000 12.750 1.6424 0.01967 0.01482 -0.0714 0.2716 1.0000 13.000 1.6525 0.02057 0.01573 -0.0697 0.2655 1.0000 13.250 1.6602 0.02163 0.01679 -0.0679 0.2596 1.0000 13.500 1.6696 0.02265 0.01784 -0.0664 0.2532 1.0000 13.750 1.6723 0.02413 0.01930 -0.0644 0.2456 1.0000 14.000 1.6769 0.02554 0.02073 -0.0628 0.2369 1.0000 14.250 1.6779 0.02727 0.02246 -0.0609 0.2294 1.0000 14.500 1.6782 0.02914 0.02434 -0.0592 0.2206 1.0000 14.750 1.6738 0.03144 0.02663 -0.0574 0.2105 1.0000 15.000 1.6641 0.03427 0.02944 -0.0556 0.1994 1.0000 15.250 1.6570 0.03704 0.03223 -0.0541 0.1908 1.0000 15.500 1.6456 0.04030 0.03549 -0.0526 0.1808 1.0000 15.750 1.6314 0.04392 0.03913 -0.0513 0.1716 1.0000 16.000 1.6142 0.04801 0.04322 -0.0502 0.1621 1.0000 16.250 1.6008 0.05186 0.04710 -0.0495 0.1541 1.0000 16.750 1.5724 0.06006 0.05537 -0.0485 0.1405 1.0000 17.000 1.5559 0.06459 0.05992 -0.0482 0.1344 1.0000 17.250 1.5467 0.06839 0.06378 -0.0482 0.1294 1.0000 17.500 1.5265 0.07355 0.06895 -0.0483 0.1219 1.0000 17.750 1.5181 0.07744 0.07289 -0.0486 0.1169 1.0000 18.000 1.5047 0.08197 0.07745 -0.0490 0.1117 1.0000 18.250 1.4922 0.08652 0.08203 -0.0495 0.1057 1.0000