EPPLER 546 AIRFOIL (e546-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 546 AIRFOIL (e546-il) Reynolds number: 1,000,000 Max Cl/Cd: 96.64 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e546-il-1000000-n5.txt Download as CSV file: xf-e546-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 546 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
------ -------- --------- --------- -------- -------- --------
-15.750 -0.8155 0.07304 0.07056 -0.0582 1.0000 0.0025
-15.500 -0.8328 0.06634 0.06373 -0.0618 1.0000 0.0025
-15.250 -0.8580 0.05934 0.05653 -0.0646 1.0000 0.0025
-15.000 -0.8713 0.05471 0.05175 -0.0659 1.0000 0.0025
-14.750 -0.8870 0.05021 0.04707 -0.0664 1.0000 0.0025
-14.500 -0.8948 0.04697 0.04370 -0.0663 1.0000 0.0026
-14.250 -0.8988 0.04433 0.04094 -0.0658 1.0000 0.0026
-14.000 -0.9104 0.04104 0.03747 -0.0648 1.0000 0.0026
-13.750 -0.9124 0.03880 0.03510 -0.0637 1.0000 0.0026
-13.500 -0.9075 0.03730 0.03354 -0.0628 1.0000 0.0027
-13.250 -0.9161 0.03455 0.03059 -0.0608 1.0000 0.0026
-13.000 -0.9095 0.03328 0.02924 -0.0596 1.0000 0.0027
-12.750 -0.9070 0.03166 0.02752 -0.0579 1.0000 0.0027
-12.500 -0.9025 0.03027 0.02602 -0.0562 1.0000 0.0027
-12.250 -0.8990 0.02884 0.02448 -0.0543 1.0000 0.0028
-12.000 -0.8970 0.02732 0.02283 -0.0519 1.0000 0.0027
-11.750 -0.8727 0.02623 0.02165 -0.0533 0.9987 0.0028
-11.500 -0.8498 0.02480 0.02011 -0.0547 0.9960 0.0029
-11.250 -0.8320 0.02359 0.01879 -0.0546 0.9902 0.0029
-11.000 -0.8190 0.02260 0.01772 -0.0534 0.9747 0.0029
-10.750 -0.7894 0.02136 0.01638 -0.0556 0.9592 0.0030
-10.500 -0.7439 0.02033 0.01524 -0.0608 0.9486 0.0030
-10.250 -0.6917 0.01914 0.01394 -0.0675 0.9398 0.0031
-10.000 -0.6346 0.01804 0.01270 -0.0751 0.9246 0.0031
-9.750 -0.5936 0.01723 0.01171 -0.0792 0.8969 0.0032
-9.500 -0.5734 0.01676 0.01108 -0.0787 0.8655 0.0032
-9.250 -0.5582 0.01631 0.01050 -0.0772 0.8444 0.0033
-9.000 -0.5469 0.01565 0.00972 -0.0750 0.8255 0.0035
-8.750 -0.5327 0.01517 0.00914 -0.0732 0.8102 0.0036
-8.500 -0.5195 0.01464 0.00852 -0.0712 0.7964 0.0036
-8.250 -0.5048 0.01422 0.00802 -0.0694 0.7834 0.0038
-8.000 -0.4897 0.01383 0.00758 -0.0677 0.7715 0.0040
-7.750 -0.4758 0.01345 0.00712 -0.0656 0.7606 0.0041
-7.250 -0.4499 0.01281 0.00635 -0.0608 0.7396 0.0044
-7.000 -0.4330 0.01252 0.00599 -0.0591 0.7305 0.0046
-6.750 -0.4141 0.01223 0.00564 -0.0578 0.7216 0.0048
-6.500 -0.3944 0.01198 0.00533 -0.0565 0.7130 0.0049
-6.250 -0.3751 0.01166 0.00497 -0.0552 0.7041 0.0057
-6.000 -0.3538 0.01143 0.00470 -0.0543 0.6959 0.0064
-5.750 -0.3322 0.01119 0.00442 -0.0534 0.6878 0.0076
-5.500 -0.3101 0.01096 0.00417 -0.0526 0.6805 0.0100
-5.250 -0.2874 0.01075 0.00393 -0.0518 0.6723 0.0134
-5.000 -0.2646 0.01053 0.00371 -0.0512 0.6648 0.0192
-4.750 -0.2418 0.01028 0.00348 -0.0505 0.6571 0.0298
-4.500 -0.2194 0.00999 0.00326 -0.0498 0.6506 0.0488
-4.250 -0.1965 0.00969 0.00305 -0.0492 0.6434 0.0739
-3.750 -0.1546 0.00875 0.00253 -0.0476 0.6294 0.1943
-3.500 -0.1341 0.00816 0.00223 -0.0468 0.6226 0.2830
-3.250 -0.1167 0.00711 0.00177 -0.0458 0.6167 0.4436
-3.000 -0.0958 0.00635 0.00152 -0.0450 0.6099 0.5918
-2.750 -0.0680 0.00636 0.00152 -0.0451 0.6032 0.6170
-2.500 -0.0396 0.00639 0.00154 -0.0453 0.5970 0.6337
-2.250 -0.0113 0.00647 0.00157 -0.0455 0.5906 0.6468
-2.000 0.0171 0.00656 0.00165 -0.0456 0.5848 0.6575
-1.750 0.0456 0.00663 0.00167 -0.0458 0.5781 0.6638
-1.500 0.0737 0.00667 0.00165 -0.0460 0.5720 0.6654
-1.250 0.1025 0.00668 0.00163 -0.0463 0.5666 0.6663
-1.000 0.1309 0.00671 0.00161 -0.0466 0.5603 0.6673
-0.750 0.1593 0.00675 0.00160 -0.0468 0.5546 0.6684
-0.500 0.1880 0.00677 0.00159 -0.0471 0.5491 0.6693
-0.250 0.2163 0.00681 0.00159 -0.0473 0.5432 0.6702
0.000 0.2446 0.00685 0.00159 -0.0476 0.5379 0.6711
0.250 0.2732 0.00688 0.00159 -0.0479 0.5324 0.6720
0.500 0.3014 0.00692 0.00160 -0.0481 0.5268 0.6729
0.750 0.3297 0.00697 0.00162 -0.0484 0.5216 0.6738
1.000 0.3581 0.00701 0.00164 -0.0486 0.5162 0.6748
1.250 0.3861 0.00707 0.00167 -0.0488 0.5104 0.6758
1.500 0.4144 0.00712 0.00170 -0.0491 0.5051 0.6765
1.750 0.4426 0.00718 0.00173 -0.0493 0.4990 0.6772
2.000 0.4701 0.00724 0.00177 -0.0494 0.4922 0.6781
2.250 0.4981 0.00728 0.00181 -0.0496 0.4851 0.6791
2.750 0.5534 0.00742 0.00192 -0.0499 0.4715 0.6809
3.000 0.5808 0.00749 0.00199 -0.0500 0.4645 0.6819
3.250 0.6083 0.00757 0.00206 -0.0501 0.4579 0.6828
3.500 0.6357 0.00765 0.00214 -0.0502 0.4499 0.6838
3.750 0.6628 0.00775 0.00223 -0.0503 0.4420 0.6848
4.000 0.6896 0.00786 0.00232 -0.0503 0.4325 0.6859
4.250 0.7166 0.00796 0.00241 -0.0503 0.4229 0.6871
4.500 0.7429 0.00809 0.00253 -0.0502 0.4132 0.6882
4.750 0.7694 0.00822 0.00264 -0.0502 0.4022 0.6893
5.000 0.7954 0.00836 0.00276 -0.0500 0.3900 0.6903
5.250 0.8209 0.00853 0.00290 -0.0498 0.3764 0.6912
5.500 0.8456 0.00875 0.00306 -0.0495 0.3592 0.6921
5.750 0.8689 0.00903 0.00325 -0.0489 0.3375 0.6930
6.000 0.8920 0.00931 0.00346 -0.0483 0.3152 0.6940
6.250 0.9126 0.00970 0.00374 -0.0472 0.2851 0.6952
6.500 0.9335 0.01008 0.00402 -0.0463 0.2611 0.6965
6.750 0.9541 0.01046 0.00431 -0.0452 0.2370 0.6976
7.000 0.9736 0.01087 0.00464 -0.0440 0.2139 0.6989
7.250 0.9929 0.01128 0.00496 -0.0428 0.1919 0.7001
7.500 1.0102 0.01175 0.00533 -0.0412 0.1677 0.7015
7.750 1.0254 0.01229 0.00574 -0.0393 0.1425 0.7030
8.000 1.0413 0.01269 0.00609 -0.0374 0.1269 0.7045
8.250 1.0558 0.01307 0.00643 -0.0353 0.1139 0.7060
8.500 1.0696 0.01352 0.00682 -0.0330 0.1009 0.7074
8.750 1.0826 0.01402 0.00727 -0.0308 0.0873 0.7086
9.000 1.0943 0.01457 0.00776 -0.0283 0.0737 0.7098
9.250 1.1061 0.01512 0.00828 -0.0260 0.0632 0.7115
9.500 1.1159 0.01575 0.00888 -0.0235 0.0516 0.7131
9.750 1.1278 0.01633 0.00945 -0.0213 0.0452 0.7147
10.000 1.1368 0.01704 0.01014 -0.0189 0.0366 0.7165
10.250 1.1469 0.01774 0.01084 -0.0168 0.0310 0.7183
10.500 1.1558 0.01853 0.01162 -0.0146 0.0254 0.7200
10.750 1.1642 0.01940 0.01248 -0.0125 0.0204 0.7219
11.000 1.1736 0.02026 0.01336 -0.0107 0.0173 0.7236
11.250 1.1829 0.02117 0.01429 -0.0090 0.0151 0.7252
11.500 1.1931 0.02207 0.01523 -0.0074 0.0134 0.7271
11.750 1.2020 0.02308 0.01629 -0.0059 0.0118 0.7295
12.000 1.2108 0.02415 0.01740 -0.0045 0.0103 0.7318
12.250 1.2198 0.02526 0.01857 -0.0032 0.0094 0.7343
12.500 1.2282 0.02645 0.01980 -0.0020 0.0082 0.7367
12.750 1.2360 0.02773 0.02114 -0.0009 0.0073 0.7389
13.000 1.2442 0.02904 0.02249 0.0001 0.0066 0.7410
13.250 1.2501 0.03055 0.02406 0.0012 0.0055 0.7434
13.500 1.2566 0.03207 0.02563 0.0020 0.0047 0.7460
13.750 1.2613 0.03378 0.02741 0.0029 0.0038 0.7488
14.000 1.2666 0.03551 0.02921 0.0036 0.0033 0.7517
14.250 1.2709 0.03738 0.03114 0.0042 0.0028 0.7546
14.500 1.2763 0.03922 0.03305 0.0047 0.0026 0.7573
14.750 1.2810 0.04117 0.03508 0.0051 0.0024 0.7604
15.000 1.2837 0.04338 0.03736 0.0054 0.0020 0.7639
15.250 1.2860 0.04569 0.03976 0.0055 0.0017 0.7676
15.500 1.2889 0.04804 0.04219 0.0056 0.0016 0.7712
16.000 1.2941 0.05297 0.04731 0.0053 0.0014 0.7794
16.250 1.2935 0.05589 0.05031 0.0050 0.0011 0.7844
16.500 1.2933 0.05888 0.05340 0.0045 0.0009 0.7894
16.750 1.2953 0.06166 0.05628 0.0040 0.0010 0.7953
17.000 1.2956 0.06471 0.05944 0.0033 0.0010 0.8016
17.250 1.2950 0.06797 0.06280 0.0025 0.0008 0.8082
17.500 1.2941 0.07131 0.06626 0.0015 0.0009 0.8160
17.750 1.2919 0.07494 0.07001 0.0004 0.0008 0.8242
18.000 1.2909 0.07843 0.07363 -0.0007 0.0008 0.8350
18.250 1.2876 0.08228 0.07762 -0.0021 0.0007 0.8482
18.500 1.2828 0.08638 0.08187 -0.0035 0.0007 0.8681
18.750 1.2811 0.09048 0.08631 -0.0056 0.0006 0.9557
19.000 1.2765 0.09488 0.09081 -0.0076 0.0007 1.0000
19.250 1.2688 0.09987 0.09590 -0.0098 0.0006 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 546 AIRFOIL (e546-il)