XFOIL Version 6.94 Calculated polar for: EPPLER 553 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 ------- -------- --------- --------- -------- ------- ------- 0.000 0.4424 0.01242 0.00530 -0.1034 0.6375 0.5406 0.500 0.4996 0.01252 0.00521 -0.1038 0.6243 0.5468 1.000 0.5527 0.01245 0.00524 -0.1034 0.6110 0.5532 1.500 0.6088 0.01262 0.00535 -0.1036 0.5990 0.5609 2.000 0.6629 0.01267 0.00538 -0.1034 0.5859 0.5680 2.500 0.7176 0.01281 0.00557 -0.1034 0.5743 0.5751 3.000 0.7714 0.01294 0.00573 -0.1032 0.5620 0.5838 3.500 0.8252 0.01313 0.00597 -0.1030 0.5502 0.5925 4.000 0.8782 0.01330 0.00619 -0.1026 0.5379 0.6019 4.500 0.9305 0.01352 0.00650 -0.1021 0.5258 0.6117 5.000 0.9829 0.01375 0.00678 -0.1016 0.5137 0.6231 5.500 1.0323 0.01397 0.00718 -0.1006 0.5010 0.6355 6.000 1.0843 0.01431 0.00751 -0.1001 0.4876 0.6495 6.500 1.1289 0.01445 0.00791 -0.0981 0.4740 0.6647 7.000 1.1748 0.01475 0.00835 -0.0965 0.4596 0.6831 7.500 1.2187 0.01506 0.00876 -0.0944 0.4439 0.7050 8.000 1.2558 0.01530 0.00927 -0.0911 0.4278 0.7312 8.500 1.2885 0.01561 0.00982 -0.0870 0.4104 0.7644 9.000 1.3100 0.01591 0.01037 -0.0807 0.3924 0.8172 9.500 1.3253 0.01619 0.01094 -0.0734 0.3723 1.0000 10.000 1.3463 0.01712 0.01188 -0.0683 0.3477 1.0000 10.500 1.3597 0.01844 0.01317 -0.0626 0.3187 1.0000 11.000 1.3645 0.02038 0.01503 -0.0564 0.2870 1.0000 11.500 1.3654 0.02293 0.01750 -0.0507 0.2522 1.0000 12.000 1.3610 0.02624 0.02072 -0.0455 0.2192 1.0000 12.500 1.3516 0.03039 0.02477 -0.0410 0.1883 1.0000 13.000 1.3452 0.03487 0.02920 -0.0378 0.1606 1.0000 13.500 1.3345 0.04022 0.03450 -0.0353 0.1369 1.0000 14.000 1.3282 0.04568 0.03998 -0.0338 0.1164 1.0000 14.500 1.3193 0.05189 0.04619 -0.0330 0.0989 1.0000 15.000 1.3101 0.05861 0.05292 -0.0330 0.0844 1.0000 15.500 1.3037 0.06545 0.05983 -0.0335 0.0715 1.0000 16.000 1.2952 0.07294 0.06738 -0.0348 0.0611 1.0000 16.500 1.2859 0.08092 0.07541 -0.0366 0.0528 1.0000 17.000 1.2779 0.08909 0.08366 -0.0389 0.0459 1.0000 17.500 1.2716 0.09729 0.09197 -0.0415 0.0402 1.0000 18.000 1.2661 0.10559 0.10038 -0.0445 0.0355 1.0000 18.500 1.2630 0.11372 0.10865 -0.0478 0.0314 1.0000 19.000 1.2603 0.12182 0.11690 -0.0514 0.0280 1.0000 19.500 1.2580 0.12985 0.12506 -0.0551 0.0251 1.0000 20.000 1.2576 0.13738 0.13263 -0.0589 0.0227 1.0000 20.500 1.2564 0.14541 0.14087 -0.0633 0.0204 1.0000 21.000 1.2548 0.15330 0.14888 -0.0677 0.0185 1.0000 21.500 1.2537 0.16115 0.15685 -0.0725 0.0168 1.0000 22.500 1.2512 0.17672 0.17271 -0.0825 0.0140 1.0000 23.000 1.2497 0.18439 0.18052 -0.0876 0.0130 1.0000 23.500 1.2419 0.19370 0.19009 -0.0940 0.0121 1.0000 24.000 1.2425 0.20099 0.19744 -0.0993 0.0112 1.0000 25.000 1.0511 0.28158 0.27891 -0.1402 0.0137 1.0000 27.000 1.0831 0.31442 0.31165 -0.1589 0.0150 1.0000 27.500 1.0951 0.31910 0.31633 -0.1627 0.0139 1.0000 28.500 1.1154 0.33094 0.32815 -0.1711 0.0126 1.0000 29.000 1.1261 0.33595 0.33317 -0.1751 0.0117 1.0000 29.500 1.1373 0.33998 0.33721 -0.1788 0.0112 1.0000 30.500 1.1564 0.35080 0.34801 -0.1873 0.0104 1.0000 31.000 1.1660 0.35571 0.35293 -0.1914 0.0097 1.0000 31.500 1.1756 0.35994 0.35717 -0.1953 0.0092 1.0000 32.000 1.1850 0.36353 0.36077 -0.1992 0.0089 1.0000 32.500 1.1950 0.36674 0.36400 -0.2026 0.0087 1.0000 33.000 1.2032 0.37127 0.36853 -0.2067 0.0086 1.0000 33.500 1.2105 0.37697 0.37422 -0.2115 0.0083 1.0000 34.000 1.2186 0.38127 0.37853 -0.2157 0.0078 1.0000 34.500 1.2263 0.38505 0.38232 -0.2197 0.0074 1.0000 35.000 1.2336 0.38858 0.38587 -0.2238 0.0071 1.0000 35.500 1.2406 0.39175 0.38906 -0.2278 0.0068 1.0000 36.000 1.2471 0.39456 0.39188 -0.2317 0.0067 1.0000 37.000 1.2594 0.40127 0.39862 -0.2396 0.0065 1.0000 37.500 1.2648 0.40506 0.40241 -0.2441 0.0065 1.0000 38.000 1.2705 0.40873 0.40610 -0.2484 0.0063 1.0000 38.500 1.2757 0.41189 0.40928 -0.2525 0.0061 1.0000 39.000 1.2806 0.41463 0.41204 -0.2567 0.0058 1.0000 39.500 1.2849 0.41706 0.41449 -0.2609 0.0056 1.0000