EPPLER 558 AIRFOIL (e558-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 558 AIRFOIL (e558-il) Reynolds number: 50,000 Max Cl/Cd: 29.7 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e558-il-50000-n5.txt Download as CSV file: xf-e558-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 558 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.750 -0.4025 0.12124 0.11358 -0.0435 1.0000 0.0608
-12.500 -0.4198 0.11307 0.10548 -0.0469 1.0000 0.0622
-12.250 -0.4493 0.10221 0.09471 -0.0517 1.0000 0.0628
-12.000 -0.4308 0.10332 0.09588 -0.0499 1.0000 0.0656
-11.750 -0.4455 0.09671 0.08934 -0.0524 1.0000 0.0679
-11.500 -0.4946 0.08349 0.07613 -0.0589 1.0000 0.0672
-11.250 -0.5326 0.07394 0.06648 -0.0636 1.0000 0.0672
-11.000 -0.5552 0.06780 0.06026 -0.0661 1.0000 0.0679
-10.750 -0.5735 0.06283 0.05519 -0.0677 1.0000 0.0691
-10.500 -0.5923 0.05828 0.05051 -0.0687 1.0000 0.0709
-10.250 -0.6103 0.05436 0.04644 -0.0690 1.0000 0.0725
-10.000 -0.6182 0.05236 0.04462 -0.0676 1.0000 0.0750
-9.750 -0.6333 0.04988 0.04216 -0.0664 1.0000 0.0769
-9.500 -0.6454 0.04715 0.03936 -0.0656 1.0000 0.0799
-9.250 -0.6549 0.04462 0.03673 -0.0642 1.0000 0.0840
-9.000 -0.6586 0.04285 0.03508 -0.0627 1.0000 0.0888
-8.750 -0.6620 0.04067 0.03278 -0.0614 1.0000 0.0958
-8.500 -0.6610 0.03904 0.03123 -0.0600 1.0000 0.1038
-8.250 -0.6577 0.03745 0.02968 -0.0589 1.0000 0.1144
-8.000 -0.6514 0.03604 0.02829 -0.0580 1.0000 0.1270
-7.750 -0.6424 0.03492 0.02718 -0.0572 1.0000 0.1417
-7.500 -0.6238 0.03398 0.02616 -0.0579 0.9974 0.1610
-7.250 -0.5850 0.03356 0.02572 -0.0616 0.9873 0.1853
-7.000 -0.5472 0.03316 0.02516 -0.0648 0.9768 0.2084
-6.750 -0.5103 0.03282 0.02464 -0.0675 0.9660 0.2286
-6.500 -0.4738 0.03266 0.02436 -0.0697 0.9551 0.2458
-6.250 -0.4357 0.03239 0.02393 -0.0721 0.9449 0.2624
-6.000 -0.4020 0.03193 0.02325 -0.0739 0.9328 0.2778
-5.750 -0.3682 0.03187 0.02313 -0.0749 0.9211 0.2904
-5.500 -0.3311 0.03159 0.02273 -0.0767 0.9111 0.3033
-5.250 -0.2952 0.03111 0.02207 -0.0785 0.9002 0.3167
-5.000 -0.2617 0.03053 0.02124 -0.0801 0.8884 0.3302
-4.750 -0.2202 0.03017 0.02077 -0.0823 0.8806 0.3423
-4.500 -0.1892 0.02982 0.02031 -0.0829 0.8679 0.3530
-4.250 -0.1519 0.02921 0.01946 -0.0850 0.8578 0.3660
-4.000 -0.1145 0.02887 0.01904 -0.0864 0.8478 0.3768
-3.750 -0.0825 0.02845 0.01848 -0.0872 0.8357 0.3872
-3.500 -0.0412 0.02783 0.01759 -0.0898 0.8273 0.4007
-3.250 -0.0133 0.02765 0.01741 -0.0894 0.8138 0.4087
-3.000 0.0193 0.02723 0.01679 -0.0904 0.8020 0.4205
-2.750 0.0556 0.02687 0.01636 -0.0915 0.7923 0.4306
-2.500 0.0838 0.02662 0.01598 -0.0916 0.7790 0.4409
-2.250 0.1152 0.02638 0.01562 -0.0921 0.7676 0.4517
-2.000 0.1486 0.02608 0.01523 -0.0928 0.7570 0.4619
-1.750 0.1761 0.02594 0.01493 -0.0928 0.7440 0.4736
-1.500 0.2059 0.02578 0.01474 -0.0928 0.7332 0.4832
-1.250 0.2364 0.02561 0.01444 -0.0932 0.7221 0.4948
-1.000 0.2623 0.02558 0.01437 -0.0927 0.7099 0.5056
-0.750 0.2941 0.02542 0.01411 -0.0931 0.7003 0.5171
-0.500 0.3203 0.02543 0.01405 -0.0929 0.6884 0.5297
-0.250 0.3471 0.02543 0.01403 -0.0925 0.6777 0.5410
0.000 0.3766 0.02537 0.01391 -0.0925 0.6678 0.5536
0.250 0.4015 0.02547 0.01398 -0.0921 0.6568 0.5671
0.500 0.4325 0.02539 0.01384 -0.0923 0.6483 0.5816
0.750 0.4545 0.02557 0.01407 -0.0913 0.6370 0.5948
1.000 0.4840 0.02555 0.01402 -0.0913 0.6288 0.6108
1.250 0.5065 0.02574 0.01428 -0.0904 0.6184 0.6270
1.500 0.5344 0.02576 0.01430 -0.0901 0.6103 0.6452
1.750 0.5564 0.02596 0.01459 -0.0891 0.6004 0.6645
2.000 0.5836 0.02599 0.01464 -0.0886 0.5930 0.6872
2.250 0.6034 0.02626 0.01504 -0.0872 0.5834 0.7114
2.500 0.6302 0.02619 0.01503 -0.0864 0.5769 0.7410
2.750 0.6451 0.02652 0.01556 -0.0841 0.5673 0.7750
3.000 0.6692 0.02645 0.01559 -0.0828 0.5607 0.8237
3.250 0.6940 0.02667 0.01601 -0.0823 0.5520 0.9101
3.500 0.7300 0.02690 0.01612 -0.0843 0.5443 1.0000
3.750 0.7552 0.02749 0.01663 -0.0845 0.5365 1.0000
4.000 0.7806 0.02803 0.01711 -0.0845 0.5289 1.0000
4.250 0.8112 0.02837 0.01731 -0.0851 0.5229 1.0000
4.500 0.8283 0.02924 0.01823 -0.0841 0.5143 1.0000
4.750 0.8577 0.02961 0.01853 -0.0844 0.5083 1.0000
5.000 0.8764 0.03046 0.01941 -0.0835 0.5009 1.0000
5.250 0.8989 0.03112 0.02007 -0.0830 0.4939 1.0000
5.500 0.9314 0.03136 0.02024 -0.0835 0.4889 1.0000
5.750 0.9400 0.03269 0.02171 -0.0815 0.4806 1.0000
6.000 0.9658 0.03321 0.02222 -0.0813 0.4747 1.0000
6.250 0.9875 0.03397 0.02301 -0.0807 0.4686 1.0000
6.500 0.9992 0.03518 0.02434 -0.0790 0.4612 1.0000
6.750 1.0281 0.03556 0.02471 -0.0791 0.4560 1.0000
7.000 1.0374 0.03694 0.02622 -0.0772 0.4491 1.0000
7.250 1.0518 0.03802 0.02739 -0.0758 0.4423 1.0000
7.500 1.0850 0.03819 0.02754 -0.0763 0.4377 1.0000
7.750 1.0766 0.04049 0.03003 -0.0728 0.4297 1.0000
8.000 1.0959 0.04130 0.03092 -0.0719 0.4237 1.0000
8.250 1.1314 0.04132 0.03094 -0.0725 0.4192 1.0000
8.500 1.0962 0.04497 0.03481 -0.0668 0.4102 1.0000
8.750 1.1274 0.04508 0.03496 -0.0669 0.4051 1.0000
9.000 1.1037 0.04848 0.03848 -0.0629 0.3974 1.0000
9.250 1.0988 0.05096 0.04106 -0.0609 0.3901 1.0000
9.500 1.1506 0.04946 0.03962 -0.0617 0.3865 1.0000
9.750 1.0615 0.05920 0.04947 -0.0578 0.3731 1.0000
10.000 1.1069 0.05743 0.04779 -0.0572 0.3702 1.0000
10.250 1.0353 0.06807 0.05850 -0.0575 0.3553 1.0000
10.500 1.0769 0.06607 0.05659 -0.0561 0.3529 1.0000
11.000 1.0516 0.07548 0.06615 -0.0567 0.3348 1.0000
11.500 0.9843 0.09192 0.08271 -0.0609 0.3090 1.0000
11.750 1.0103 0.09153 0.08243 -0.0597 0.3049 1.0000
12.250 1.0085 0.09875 0.08982 -0.0608 0.2888 1.0000
12.750 1.0079 0.10590 0.09714 -0.0621 0.2730 1.0000
13.250 1.0094 0.11280 0.10421 -0.0635 0.2573 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 558 AIRFOIL (e558-il)