EPPLER 556 AIRFOIL (e556-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: EPPLER 556 AIRFOIL (e556-il) Reynolds number: 50,000 Max Cl/Cd: 25.5 at α=11.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e556-il-50000.txt Download as CSV file: xf-e556-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 556 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
------ -------- --------- --------- -------- -------- --------
-11.250 -0.4240 0.11459 0.10752 -0.0482 1.0000 0.1401
-11.000 -0.4207 0.10992 0.10287 -0.0485 1.0000 0.1374
-10.750 -0.5266 0.08844 0.08164 -0.0646 1.0000 0.1104
-10.500 -0.5325 0.08359 0.07684 -0.0649 1.0000 0.1093
-10.250 -0.5524 0.07834 0.07164 -0.0659 1.0000 0.1080
-10.000 -0.5807 0.07352 0.06687 -0.0665 1.0000 0.1062
-9.750 -0.6163 0.06967 0.06306 -0.0658 1.0000 0.1042
-9.500 -0.7052 0.06535 0.05826 -0.0655 1.0000 0.0964
-9.250 -0.7163 0.06190 0.05458 -0.0642 1.0000 0.0963
-9.000 -0.7207 0.05806 0.05064 -0.0629 1.0000 0.0972
-8.750 -0.7171 0.05482 0.04731 -0.0616 1.0000 0.0987
-8.500 -0.7132 0.05181 0.04414 -0.0602 1.0000 0.1006
-8.250 -0.7081 0.04875 0.04081 -0.0591 1.0000 0.1019
-8.000 -0.6997 0.04597 0.03774 -0.0581 1.0000 0.1051
-7.750 -0.6898 0.04329 0.03448 -0.0575 1.0000 0.1099
-7.500 -0.6743 0.04089 0.03221 -0.0559 1.0000 0.1163
-7.250 -0.6584 0.03874 0.02972 -0.0548 1.0000 0.1253
-7.000 -0.6424 0.03693 0.02802 -0.0530 1.0000 0.1375
-6.750 -0.6268 0.03518 0.02630 -0.0512 1.0000 0.1535
-6.500 -0.6133 0.03346 0.02494 -0.0489 1.0000 0.1748
-6.250 -0.5991 0.03131 0.02319 -0.0479 1.0000 0.2111
-6.000 -0.5852 0.02973 0.02317 -0.0466 1.0000 0.3132
-5.750 -0.5790 0.03515 0.02905 -0.0366 1.0000 0.4161
-5.500 -0.5718 0.03918 0.03299 -0.0284 1.0000 0.4583
-5.250 -0.5673 0.04309 0.03690 -0.0183 1.0000 0.4795
-5.000 -0.5597 0.04524 0.03893 -0.0119 1.0000 0.5072
-4.750 -0.5537 0.04761 0.04124 -0.0036 1.0000 0.5271
-4.500 -0.5464 0.04876 0.04228 0.0021 1.0000 0.5509
-4.250 -0.5396 0.04940 0.04281 0.0072 1.0000 0.5732
-4.000 -0.5326 0.05008 0.04339 0.0130 1.0000 0.5933
-3.750 -0.5255 0.05005 0.04326 0.0168 1.0000 0.6142
-3.500 -0.5178 0.04942 0.04248 0.0184 1.0000 0.6349
-3.250 -0.5105 0.04913 0.04209 0.0220 1.0000 0.6517
-3.000 -0.5017 0.04854 0.04139 0.0241 1.0000 0.6673
-2.750 -0.4926 0.04799 0.04074 0.0266 1.0000 0.6803
-2.500 -0.4825 0.04729 0.03991 0.0283 1.0000 0.6911
-2.250 -0.4687 0.04632 0.03880 0.0275 1.0000 0.7009
-2.000 -0.4519 0.04570 0.03806 0.0273 0.9988 0.7093
-1.750 -0.4062 0.04572 0.03783 0.0212 0.9864 0.7182
-1.500 -0.3644 0.04565 0.03754 0.0159 0.9746 0.7257
-1.250 -0.3237 0.04541 0.03709 0.0103 0.9629 0.7320
-1.000 -0.2915 0.04520 0.03674 0.0074 0.9517 0.7374
-0.750 -0.2544 0.04503 0.03641 0.0027 0.9401 0.7429
-0.500 -0.2172 0.04497 0.03620 -0.0017 0.9291 0.7480
-0.250 -0.1784 0.04508 0.03618 -0.0056 0.9180 0.7529
0.000 -0.1466 0.04502 0.03601 -0.0089 0.9066 0.7580
0.250 -0.1132 0.04507 0.03596 -0.0125 0.8952 0.7631
0.500 -0.0801 0.04524 0.03605 -0.0151 0.8845 0.7679
0.750 -0.0409 0.04550 0.03622 -0.0192 0.8735 0.7738
1.000 -0.0139 0.04568 0.03635 -0.0216 0.8620 0.7789
1.250 0.0145 0.04596 0.03658 -0.0233 0.8514 0.7842
1.500 0.0602 0.04640 0.03694 -0.0280 0.8406 0.7907
1.750 0.0779 0.04671 0.03725 -0.0286 0.8294 0.7962
2.000 0.1047 0.04716 0.03768 -0.0301 0.8185 0.8025
2.250 0.1476 0.04769 0.03817 -0.0343 0.8080 0.8097
2.500 0.1653 0.04818 0.03868 -0.0344 0.7970 0.8160
2.750 0.1925 0.04888 0.03937 -0.0365 0.7859 0.8235
3.000 0.2275 0.04934 0.03985 -0.0384 0.7762 0.8318
3.250 0.2499 0.05012 0.04066 -0.0399 0.7647 0.8398
3.500 0.2666 0.05087 0.04145 -0.0398 0.7543 0.8479
3.750 0.3025 0.05144 0.04205 -0.0420 0.7441 0.8584
4.000 0.3206 0.05228 0.04294 -0.0424 0.7329 0.8689
4.250 0.3353 0.05323 0.04397 -0.0423 0.7223 0.8797
4.500 0.3674 0.05377 0.04458 -0.0437 0.7120 0.8938
4.750 0.3830 0.05473 0.04563 -0.0438 0.7012 0.9090
5.000 0.3988 0.05583 0.04685 -0.0442 0.6900 0.9282
5.500 0.4830 0.05763 0.04888 -0.0522 0.6654 1.0000
5.750 0.5119 0.05938 0.05070 -0.0559 0.6525 1.0000
6.000 0.5381 0.06132 0.05268 -0.0592 0.6390 1.0000
6.250 0.5678 0.06321 0.05463 -0.0627 0.6256 1.0000
6.500 0.6015 0.06490 0.05638 -0.0661 0.6118 1.0000
6.750 0.6365 0.06646 0.05800 -0.0693 0.5981 1.0000
7.000 0.6734 0.06776 0.05935 -0.0720 0.5841 1.0000
7.250 0.7162 0.06848 0.06015 -0.0744 0.5704 1.0000
7.500 0.7440 0.06989 0.06162 -0.0757 0.5563 1.0000
7.750 0.7570 0.07216 0.06393 -0.0762 0.5416 1.0000
8.000 0.7679 0.07462 0.06645 -0.0765 0.5272 1.0000
8.250 0.7821 0.07678 0.06867 -0.0767 0.5125 1.0000
8.500 0.7959 0.07897 0.07092 -0.0768 0.4979 1.0000
8.750 0.8101 0.08110 0.07313 -0.0768 0.4833 1.0000
9.000 0.8245 0.08321 0.07530 -0.0766 0.4687 1.0000
9.250 0.8390 0.08530 0.07750 -0.0764 0.4542 1.0000
9.500 0.8546 0.08718 0.07947 -0.0761 0.4394 1.0000
9.750 0.8704 0.08909 0.08147 -0.0757 0.4250 1.0000
10.000 0.8877 0.09070 0.08318 -0.0751 0.4103 1.0000
10.250 0.9081 0.09192 0.08454 -0.0744 0.3960 1.0000
10.500 0.9279 0.09298 0.08572 -0.0734 0.3812 1.0000
10.750 0.9561 0.09280 0.08569 -0.0719 0.3667 1.0000
11.000 1.1530 0.06679 0.06027 -0.0613 0.3415 1.0000
11.250 0.9027 0.10827 0.10114 -0.0760 0.3440 1.0000
11.500 1.3519 0.05301 0.04619 -0.0614 0.2701 1.0000
11.750 1.1137 0.08089 0.07451 -0.0599 0.3037 1.0000
12.000 0.9516 0.11350 0.10673 -0.0740 0.3050 1.0000
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Polar data table (+)
Polar graphs
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