EPPLER 557 AIRFOIL (e557-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: EPPLER 557 AIRFOIL (e557-il) Reynolds number: 50,000 Max Cl/Cd: 25.42 at α=11° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e557-il-50000.txt Download as CSV file: xf-e557-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 557 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.3113 0.10925 0.10307 -0.0414 1.0000 0.1765
-10.000 -0.5184 0.08513 0.07869 -0.0638 1.0000 0.1141
-9.750 -0.5345 0.08079 0.07439 -0.0635 1.0000 0.1132
-9.500 -0.5557 0.07685 0.07053 -0.0627 1.0000 0.1122
-9.250 -0.5846 0.07344 0.06722 -0.0612 1.0000 0.1107
-9.000 -0.6165 0.06959 0.06339 -0.0608 1.0000 0.1083
-8.750 -0.6485 0.06517 0.05885 -0.0611 1.0000 0.1060
-8.500 -0.6732 0.06054 0.05393 -0.0616 1.0000 0.1042
-8.250 -0.6811 0.05667 0.04978 -0.0617 1.0000 0.1047
-8.000 -0.6805 0.05313 0.04595 -0.0618 1.0000 0.1064
-7.750 -0.6746 0.04964 0.04207 -0.0622 1.0000 0.1089
-7.500 -0.6631 0.04613 0.03801 -0.0629 1.0000 0.1119
-7.250 -0.6477 0.04307 0.03466 -0.0628 1.0000 0.1167
-7.000 -0.6307 0.04097 0.03246 -0.0621 1.0000 0.1244
-6.750 -0.6128 0.03867 0.03003 -0.0614 1.0000 0.1334
-6.500 -0.5942 0.03676 0.02799 -0.0606 1.0000 0.1456
-6.250 -0.5772 0.03507 0.02643 -0.0592 1.0000 0.1619
-6.000 -0.5600 0.03341 0.02498 -0.0580 1.0000 0.1848
-5.750 -0.5404 0.03139 0.02343 -0.0580 1.0000 0.2254
-5.500 -0.5222 0.03102 0.02436 -0.0564 1.0000 0.3204
-5.250 -0.5177 0.03512 0.02886 -0.0480 1.0000 0.3989
-5.000 -0.5132 0.03864 0.03235 -0.0402 1.0000 0.4422
-4.750 -0.5126 0.04177 0.03551 -0.0310 1.0000 0.4646
-4.500 -0.5064 0.04372 0.03736 -0.0250 1.0000 0.4926
-4.250 -0.5038 0.04560 0.03921 -0.0175 1.0000 0.5125
-4.000 -0.4993 0.04687 0.04041 -0.0113 1.0000 0.5341
-3.750 -0.4900 0.04727 0.04066 -0.0081 1.0000 0.5592
-3.500 -0.4859 0.04792 0.04124 -0.0024 1.0000 0.5779
-3.250 -0.4806 0.04826 0.04151 0.0025 1.0000 0.5971
-3.000 -0.4740 0.04833 0.04149 0.0065 1.0000 0.6166
-2.750 -0.4658 0.04817 0.04122 0.0095 1.0000 0.6358
-2.500 -0.4564 0.04785 0.04079 0.0119 1.0000 0.6531
-2.250 -0.4460 0.04741 0.04024 0.0138 1.0000 0.6678
-2.000 -0.4084 0.04767 0.04027 0.0105 0.9891 0.6838
-1.750 -0.3698 0.04762 0.03999 0.0063 0.9776 0.6958
-1.500 -0.3284 0.04746 0.03958 0.0007 0.9656 0.7057
-1.250 -0.2967 0.04728 0.03925 -0.0018 0.9548 0.7129
-1.000 -0.2575 0.04719 0.03896 -0.0067 0.9439 0.7199
-0.750 -0.2143 0.04725 0.03882 -0.0121 0.9328 0.7263
-0.500 -0.1892 0.04700 0.03846 -0.0139 0.9223 0.7314
-0.250 -0.1519 0.04705 0.03835 -0.0184 0.9114 0.7372
0.000 -0.1111 0.04724 0.03840 -0.0229 0.9007 0.7426
0.250 -0.0844 0.04727 0.03836 -0.0248 0.8903 0.7474
0.500 -0.0503 0.04751 0.03847 -0.0284 0.8794 0.7533
0.750 -0.0080 0.04789 0.03875 -0.0329 0.8693 0.7589
1.000 0.0137 0.04807 0.03889 -0.0339 0.8587 0.7637
1.250 0.0468 0.04852 0.03926 -0.0372 0.8483 0.7700
1.500 0.0868 0.04898 0.03966 -0.0406 0.8385 0.7764
1.750 0.1046 0.04939 0.04005 -0.0413 0.8277 0.7818
2.000 0.1424 0.05008 0.04068 -0.0453 0.8176 0.7886
2.250 0.1694 0.05053 0.04112 -0.0464 0.8076 0.7953
2.500 0.1926 0.05128 0.04185 -0.0480 0.7972 0.8023
2.750 0.2307 0.05187 0.04243 -0.0508 0.7875 0.8099
3.000 0.2486 0.05269 0.04326 -0.0516 0.7770 0.8173
3.250 0.2725 0.05349 0.04408 -0.0527 0.7672 0.8255
3.500 0.3099 0.05420 0.04480 -0.0554 0.7573 0.8358
3.750 0.3181 0.05522 0.04587 -0.0547 0.7472 0.8437
4.000 0.3519 0.05599 0.04667 -0.0568 0.7374 0.8549
4.250 0.3682 0.05699 0.04773 -0.0571 0.7271 0.8661
4.500 0.3843 0.05802 0.04883 -0.0571 0.7175 0.8784
4.750 0.4205 0.05855 0.04943 -0.0588 0.7075 0.8953
5.000 0.4218 0.05999 0.05096 -0.0579 0.6972 0.9111
5.250 0.4489 0.06085 0.05196 -0.0592 0.6868 0.9374
5.500 0.4985 0.06158 0.05281 -0.0639 0.6742 1.0000
5.750 0.5097 0.06403 0.05530 -0.0664 0.6625 1.0000
6.000 0.5419 0.06598 0.05730 -0.0704 0.6497 1.0000
6.250 0.5817 0.06771 0.05907 -0.0748 0.6367 1.0000
6.500 0.6304 0.06898 0.06039 -0.0793 0.6234 1.0000
6.750 0.6616 0.07081 0.06226 -0.0824 0.6101 1.0000
7.000 0.6748 0.07344 0.06492 -0.0842 0.5967 1.0000
7.250 0.6954 0.07571 0.06722 -0.0860 0.5833 1.0000
7.500 0.7211 0.07759 0.06915 -0.0876 0.5697 1.0000
7.750 0.7522 0.07901 0.07060 -0.0890 0.5562 1.0000
8.000 0.7920 0.07957 0.07125 -0.0900 0.5424 1.0000
8.250 0.8038 0.08197 0.07370 -0.0903 0.5287 1.0000
8.500 0.8059 0.08513 0.07690 -0.0904 0.5147 1.0000
8.750 0.8133 0.08801 0.07984 -0.0905 0.5012 1.0000
9.000 0.8282 0.09021 0.08211 -0.0906 0.4872 1.0000
9.250 0.8478 0.09204 0.08403 -0.0906 0.4736 1.0000
9.500 0.8734 0.09315 0.08524 -0.0903 0.4594 1.0000
9.750 0.9098 0.09297 0.08518 -0.0895 0.4453 1.0000
10.000 0.9046 0.09713 0.08942 -0.0897 0.4317 1.0000
11.000 1.3699 0.05389 0.04715 -0.0813 0.3496 1.0000
11.250 1.2849 0.06296 0.05641 -0.0752 0.3487 1.0000
11.500 0.9176 0.11947 0.11222 -0.0920 0.3569 1.0000
11.750 1.3861 0.05784 0.05126 -0.0734 0.2995 1.0000
12.000 1.3230 0.06555 0.05920 -0.0695 0.2982 1.0000
12.250 1.2065 0.08321 0.07693 -0.0710 0.3024 1.0000
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Polar data table (+)
Polar graphs
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