EPPLER 547 AIRFOIL (e547-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 547 AIRFOIL (e547-il) Reynolds number: 1,000,000 Max Cl/Cd: 111.7 at α=7.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e547-il-1000000.txt Download as CSV file: xf-e547-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 547 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-15.750 -0.2818 0.14295 0.14137 -0.0380 1.0000 0.0101
-15.500 -0.2826 0.14029 0.13871 -0.0393 1.0000 0.0104
-15.250 -0.4260 0.15226 0.15051 -0.0303 1.0000 0.0093
-11.250 -0.7613 0.03243 0.02848 -0.0619 0.9938 0.0064
-11.000 -0.7455 0.03125 0.02720 -0.0621 0.9750 0.0063
-10.750 -0.7224 0.02737 0.02310 -0.0640 0.9615 0.0059
-10.500 -0.6854 0.02451 0.01995 -0.0681 0.9519 0.0056
-10.250 -0.6381 0.02209 0.01726 -0.0735 0.9442 0.0055
-10.000 -0.5869 0.02026 0.01522 -0.0795 0.9310 0.0054
-9.750 -0.5445 0.01909 0.01385 -0.0837 0.9052 0.0054
-9.500 -0.5243 0.01837 0.01294 -0.0832 0.8774 0.0054
-9.250 -0.5101 0.01772 0.01214 -0.0816 0.8562 0.0053
-9.000 -0.4972 0.01706 0.01136 -0.0796 0.8389 0.0055
-8.750 -0.4843 0.01643 0.01062 -0.0777 0.8239 0.0056
-8.500 -0.4716 0.01580 0.00989 -0.0757 0.8107 0.0056
-8.250 -0.4578 0.01528 0.00928 -0.0738 0.7985 0.0056
-8.000 -0.4448 0.01473 0.00863 -0.0718 0.7872 0.0059
-7.750 -0.4308 0.01422 0.00805 -0.0699 0.7767 0.0060
-7.500 -0.4170 0.01376 0.00752 -0.0679 0.7668 0.0061
-7.250 -0.4033 0.01334 0.00703 -0.0658 0.7565 0.0062
-7.000 -0.3967 0.01272 0.00633 -0.0624 0.7474 0.0066
-6.750 -0.3854 0.01242 0.00596 -0.0597 0.7386 0.0069
-6.500 -0.3685 0.01212 0.00563 -0.0579 0.7304 0.0082
-6.250 -0.3523 0.01177 0.00522 -0.0560 0.7225 0.0095
-6.000 -0.3334 0.01145 0.00487 -0.0546 0.7145 0.0118
-5.750 -0.3157 0.01110 0.00451 -0.0530 0.7067 0.0190
-5.500 -0.2965 0.01074 0.00422 -0.0518 0.6991 0.0319
-5.250 -0.2768 0.01043 0.00395 -0.0506 0.6920 0.0490
-5.000 -0.2561 0.01009 0.00370 -0.0496 0.6850 0.0723
-4.500 -0.2179 0.00919 0.00315 -0.0472 0.6711 0.1678
-4.250 -0.2038 0.00834 0.00274 -0.0454 0.6642 0.2818
-4.000 -0.1962 0.00679 0.00204 -0.0428 0.6579 0.5009
-3.750 -0.1743 0.00637 0.00199 -0.0419 0.6514 0.6142
-3.500 -0.1464 0.00646 0.00201 -0.0419 0.6452 0.6365
-3.250 -0.1179 0.00651 0.00205 -0.0421 0.6389 0.6466
-3.000 -0.0895 0.00659 0.00203 -0.0423 0.6324 0.6533
-2.750 -0.0610 0.00667 0.00210 -0.0424 0.6264 0.6604
-2.500 -0.0325 0.00682 0.00217 -0.0426 0.6203 0.6679
-2.250 -0.0045 0.00696 0.00230 -0.0426 0.6145 0.6751
-2.000 0.0243 0.00718 0.00248 -0.0427 0.6088 0.6836
-1.750 0.0524 0.00730 0.00262 -0.0427 0.6027 0.6894
-1.500 0.0809 0.00734 0.00261 -0.0430 0.5973 0.6910
-1.250 0.1097 0.00734 0.00257 -0.0433 0.5918 0.6922
-1.000 0.1380 0.00736 0.00253 -0.0436 0.5863 0.6932
-0.750 0.1667 0.00737 0.00250 -0.0439 0.5812 0.6942
-0.500 0.1954 0.00739 0.00247 -0.0442 0.5757 0.6953
-0.250 0.2236 0.00743 0.00245 -0.0445 0.5705 0.6963
0.000 0.2523 0.00744 0.00243 -0.0448 0.5657 0.6971
0.250 0.2809 0.00746 0.00242 -0.0452 0.5606 0.6978
0.500 0.3091 0.00751 0.00242 -0.0454 0.5555 0.6984
0.750 0.3374 0.00749 0.00237 -0.0457 0.5510 0.6996
1.000 0.3659 0.00748 0.00236 -0.0461 0.5462 0.7007
1.250 0.3939 0.00751 0.00237 -0.0463 0.5413 0.7018
1.500 0.4219 0.00755 0.00239 -0.0465 0.5368 0.7027
1.750 0.4505 0.00757 0.00241 -0.0468 0.5322 0.7036
2.000 0.4786 0.00761 0.00244 -0.0471 0.5275 0.7046
2.250 0.5060 0.00769 0.00249 -0.0472 0.5225 0.7055
2.500 0.5347 0.00770 0.00252 -0.0476 0.5183 0.7065
2.750 0.5627 0.00775 0.00256 -0.0478 0.5132 0.7074
3.000 0.5898 0.00783 0.00262 -0.0479 0.5079 0.7085
3.250 0.6182 0.00786 0.00266 -0.0482 0.5032 0.7095
3.500 0.6460 0.00791 0.00271 -0.0484 0.4979 0.7105
3.750 0.6728 0.00802 0.00278 -0.0484 0.4923 0.7116
4.000 0.7011 0.00805 0.00284 -0.0488 0.4876 0.7126
4.250 0.7286 0.00812 0.00291 -0.0489 0.4823 0.7135
4.500 0.7551 0.00823 0.00300 -0.0489 0.4768 0.7143
4.750 0.7830 0.00823 0.00304 -0.0492 0.4718 0.7157
5.000 0.8099 0.00829 0.00312 -0.0492 0.4661 0.7169
5.250 0.8362 0.00838 0.00322 -0.0492 0.4602 0.7181
5.500 0.8634 0.00844 0.00331 -0.0493 0.4537 0.7192
5.750 0.8888 0.00857 0.00343 -0.0491 0.4465 0.7204
6.000 0.9160 0.00863 0.00353 -0.0492 0.4399 0.7215
6.500 0.9675 0.00885 0.00377 -0.0489 0.4237 0.7241
6.750 0.9922 0.00900 0.00392 -0.0486 0.4145 0.7254
7.000 1.0169 0.00915 0.00407 -0.0483 0.4034 0.7267
7.250 1.0410 0.00932 0.00423 -0.0479 0.3908 0.7281
7.500 1.0637 0.00955 0.00442 -0.0472 0.3756 0.7293
7.750 1.0842 0.00982 0.00464 -0.0461 0.3567 0.7308
8.000 1.1035 0.01013 0.00489 -0.0449 0.3341 0.7325
8.250 1.1194 0.01056 0.00522 -0.0431 0.3080 0.7341
8.500 1.1307 0.01107 0.00561 -0.0404 0.2790 0.7357
8.750 1.1387 0.01160 0.00602 -0.0371 0.2522 0.7374
9.000 1.1493 0.01212 0.00646 -0.0344 0.2302 0.7392
9.250 1.1588 0.01270 0.00695 -0.0315 0.2104 0.7410
9.500 1.1676 0.01331 0.00748 -0.0287 0.1903 0.7428
9.750 1.1765 0.01393 0.00803 -0.0260 0.1729 0.7445
10.000 1.1846 0.01459 0.00864 -0.0232 0.1571 0.7464
10.250 1.1912 0.01532 0.00933 -0.0204 0.1416 0.7488
10.500 1.1955 0.01620 0.01015 -0.0175 0.1256 0.7512
10.750 1.1997 0.01715 0.01106 -0.0147 0.1107 0.7537
11.000 1.2040 0.01818 0.01204 -0.0122 0.0976 0.7562
11.250 1.2121 0.01909 0.01295 -0.0103 0.0895 0.7586
11.500 1.2157 0.02030 0.01411 -0.0081 0.0780 0.7607
11.750 1.2155 0.02176 0.01553 -0.0056 0.0648 0.7638
12.000 1.2217 0.02295 0.01674 -0.0040 0.0580 0.7667
12.250 1.2252 0.02437 0.01815 -0.0022 0.0507 0.7697
12.500 1.2330 0.02559 0.01941 -0.0010 0.0471 0.7728
12.750 1.2351 0.02724 0.02104 0.0005 0.0400 0.7758
13.000 1.2384 0.02886 0.02269 0.0018 0.0355 0.7794
13.250 1.2435 0.03040 0.02427 0.0029 0.0318 0.7832
13.500 1.2473 0.03210 0.02601 0.0039 0.0288 0.7873
13.750 1.2519 0.03382 0.02778 0.0048 0.0260 0.7914
14.000 1.2542 0.03574 0.02974 0.0056 0.0228 0.7964
14.250 1.2587 0.03756 0.03163 0.0063 0.0210 0.8020
14.500 1.2631 0.03945 0.03357 0.0068 0.0193 0.8079
14.750 1.2650 0.04159 0.03578 0.0073 0.0174 0.8152
15.000 1.2694 0.04357 0.03784 0.0077 0.0162 0.8236
15.250 1.2725 0.04568 0.04004 0.0080 0.0148 0.8345
15.500 1.2740 0.04797 0.04242 0.0084 0.0136 0.8490
15.750 1.2763 0.05006 0.04469 0.0089 0.0127 0.8754
16.000 1.2899 0.05261 0.04753 0.0062 0.0115 0.9904
16.250 1.2890 0.05528 0.05024 0.0064 0.0106 1.0000
16.500 1.2923 0.05779 0.05280 0.0060 0.0095 1.0000
16.750 1.2945 0.06047 0.05553 0.0056 0.0089 1.0000
17.000 1.2937 0.06354 0.05864 0.0051 0.0079 1.0000
17.250 1.2935 0.06661 0.06179 0.0045 0.0073 1.0000
17.500 1.2951 0.06952 0.06475 0.0038 0.0067 1.0000
17.750 1.2926 0.07300 0.06828 0.0030 0.0058 1.0000
18.000 1.2913 0.07641 0.07176 0.0020 0.0053 1.0000
18.250 1.2885 0.08007 0.07548 0.0010 0.0045 1.0000
18.500 1.2858 0.08378 0.07928 -0.0002 0.0044 1.0000
18.750 1.2790 0.08817 0.08375 -0.0017 0.0039 1.0000
19.000 1.2775 0.09190 0.08757 -0.0031 0.0037 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 547 AIRFOIL (e547-il)