EPPLER 548 AIRFOIL (e548-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 548 AIRFOIL (e548-il) Reynolds number: 500,000 Max Cl/Cd: 90.66 at α=8° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e548-il-500000-n5.txt Download as CSV file: xf-e548-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 548 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-15.500 -0.6654 0.10054 0.09781 -0.0358 1.0000 0.0039
-15.250 -0.7013 0.08676 0.08384 -0.0439 1.0000 0.0038
-15.000 -0.7268 0.07694 0.07383 -0.0500 1.0000 0.0037
-14.750 -0.7535 0.06844 0.06511 -0.0547 1.0000 0.0037
-14.500 -0.7728 0.06220 0.05868 -0.0574 1.0000 0.0037
-14.250 -0.7878 0.05723 0.05354 -0.0589 1.0000 0.0037
-14.000 -0.8025 0.05273 0.04883 -0.0596 1.0000 0.0037
-13.750 -0.8124 0.04916 0.04509 -0.0596 1.0000 0.0037
-13.500 -0.8246 0.04545 0.04115 -0.0591 1.0000 0.0037
-13.250 -0.8291 0.04272 0.03826 -0.0583 1.0000 0.0037
-13.000 -0.8310 0.04033 0.03571 -0.0573 1.0000 0.0037
-12.750 -0.8316 0.03810 0.03332 -0.0561 1.0000 0.0038
-12.500 -0.8309 0.03601 0.03107 -0.0547 1.0000 0.0038
-12.250 -0.8269 0.03431 0.02924 -0.0533 1.0000 0.0038
-12.000 -0.8224 0.03264 0.02742 -0.0518 1.0000 0.0039
-11.750 -0.8160 0.03120 0.02588 -0.0502 1.0000 0.0039
-11.500 -0.8090 0.02977 0.02430 -0.0485 1.0000 0.0040
-11.000 -0.6999 0.02539 0.01951 -0.0649 0.9116 0.0042
-10.750 -0.6453 0.02402 0.01776 -0.0726 0.8364 0.0045
-10.500 -0.6326 0.02335 0.01684 -0.0710 0.7950 0.0044
-10.250 -0.6196 0.02271 0.01604 -0.0694 0.7668 0.0047
-10.000 -0.6062 0.02207 0.01524 -0.0678 0.7447 0.0048
-9.750 -0.5927 0.02140 0.01444 -0.0662 0.7255 0.0049
-9.250 -0.5634 0.02019 0.01299 -0.0633 0.6938 0.0052
-9.000 -0.5483 0.01960 0.01229 -0.0619 0.6791 0.0057
-8.750 -0.5329 0.01904 0.01162 -0.0604 0.6674 0.0058
-8.500 -0.5177 0.01849 0.01099 -0.0589 0.6565 0.0059
-8.250 -0.5027 0.01792 0.01034 -0.0574 0.6463 0.0061
-8.000 -0.4877 0.01737 0.00973 -0.0558 0.6371 0.0064
-7.750 -0.4721 0.01690 0.00918 -0.0543 0.6280 0.0069
-7.500 -0.4566 0.01642 0.00864 -0.0526 0.6199 0.0076
-7.250 -0.4414 0.01599 0.00813 -0.0509 0.6114 0.0079
-7.000 -0.4277 0.01556 0.00766 -0.0489 0.6037 0.0085
-6.750 -0.4146 0.01523 0.00727 -0.0466 0.5961 0.0094
-6.500 -0.3997 0.01491 0.00691 -0.0446 0.5893 0.0112
-6.250 -0.3816 0.01459 0.00654 -0.0432 0.5826 0.0127
-6.000 -0.3636 0.01426 0.00618 -0.0417 0.5764 0.0154
-5.750 -0.3447 0.01392 0.00583 -0.0404 0.5709 0.0201
-5.500 -0.3251 0.01359 0.00551 -0.0393 0.5652 0.0279
-5.250 -0.3054 0.01327 0.00521 -0.0381 0.5599 0.0395
-5.000 -0.2851 0.01291 0.00491 -0.0371 0.5551 0.0564
-4.750 -0.2649 0.01253 0.00462 -0.0361 0.5498 0.0812
-4.500 -0.2448 0.01213 0.00433 -0.0350 0.5447 0.1159
-4.250 -0.2260 0.01159 0.00401 -0.0339 0.5404 0.1720
-4.000 -0.2105 0.01070 0.00356 -0.0325 0.5362 0.2786
-3.500 -0.1795 0.00872 0.00280 -0.0295 0.5284 0.5914
-3.250 -0.1518 0.00878 0.00284 -0.0295 0.5248 0.6212
-3.000 -0.1235 0.00888 0.00291 -0.0295 0.5212 0.6404
-2.750 -0.0952 0.00902 0.00300 -0.0296 0.5173 0.6557
-2.500 -0.0670 0.00922 0.00315 -0.0295 0.5134 0.6674
-2.000 -0.0104 0.00955 0.00340 -0.0295 0.5068 0.6820
-1.750 0.0184 0.00954 0.00332 -0.0298 0.5037 0.6833
-1.500 0.0472 0.00954 0.00324 -0.0301 0.5006 0.6846
-1.250 0.0758 0.00955 0.00317 -0.0304 0.4977 0.6858
-1.000 0.1043 0.00956 0.00312 -0.0307 0.4948 0.6868
-0.750 0.1328 0.00957 0.00308 -0.0309 0.4921 0.6876
-0.500 0.1617 0.00957 0.00306 -0.0312 0.4892 0.6885
-0.250 0.1904 0.00958 0.00304 -0.0315 0.4864 0.6894
0.000 0.2190 0.00960 0.00304 -0.0318 0.4838 0.6904
0.250 0.2475 0.00964 0.00304 -0.0320 0.4813 0.6915
0.500 0.2758 0.00968 0.00304 -0.0323 0.4788 0.6925
0.750 0.3043 0.00973 0.00305 -0.0326 0.4764 0.6934
1.000 0.3330 0.00974 0.00307 -0.0329 0.4739 0.6944
1.250 0.3617 0.00977 0.00309 -0.0332 0.4715 0.6955
1.500 0.3903 0.00981 0.00312 -0.0335 0.4691 0.6965
1.750 0.4187 0.00985 0.00314 -0.0338 0.4666 0.6975
2.000 0.4470 0.00991 0.00318 -0.0340 0.4641 0.6986
2.250 0.4752 0.00999 0.00322 -0.0343 0.4617 0.6999
2.500 0.5037 0.01005 0.00328 -0.0346 0.4597 0.7010
2.750 0.5323 0.01009 0.00333 -0.0349 0.4572 0.7021
3.000 0.5606 0.01013 0.00340 -0.0352 0.4546 0.7029
3.250 0.5887 0.01018 0.00346 -0.0355 0.4520 0.7039
3.500 0.6166 0.01024 0.00354 -0.0357 0.4493 0.7048
3.750 0.6443 0.01032 0.00362 -0.0358 0.4466 0.7058
4.000 0.6721 0.01040 0.00372 -0.0360 0.4438 0.7068
4.250 0.7002 0.01044 0.00381 -0.0362 0.4405 0.7078
4.500 0.7279 0.01050 0.00391 -0.0364 0.4367 0.7089
4.750 0.7553 0.01058 0.00401 -0.0365 0.4331 0.7100
5.000 0.7824 0.01069 0.00411 -0.0366 0.4297 0.7112
5.250 0.8101 0.01076 0.00424 -0.0368 0.4262 0.7125
5.500 0.8376 0.01083 0.00436 -0.0370 0.4217 0.7140
5.750 0.8645 0.01092 0.00447 -0.0370 0.4171 0.7154
6.000 0.8910 0.01103 0.00459 -0.0370 0.4127 0.7167
6.250 0.9184 0.01111 0.00473 -0.0372 0.4078 0.7179
6.500 0.9449 0.01121 0.00487 -0.0372 0.4020 0.7190
6.750 0.9707 0.01132 0.00501 -0.0370 0.3965 0.7202
7.000 0.9970 0.01141 0.00517 -0.0370 0.3894 0.7214
7.250 1.0218 0.01155 0.00534 -0.0367 0.3815 0.7227
7.500 1.0471 0.01168 0.00552 -0.0365 0.3717 0.7240
7.750 1.0712 0.01185 0.00572 -0.0361 0.3607 0.7254
8.000 1.0943 0.01207 0.00595 -0.0356 0.3478 0.7269
8.250 1.1159 0.01236 0.00622 -0.0348 0.3313 0.7285
8.500 1.1349 0.01275 0.00656 -0.0336 0.3105 0.7302
8.750 1.1513 0.01326 0.00698 -0.0320 0.2874 0.7320
9.250 1.1759 0.01454 0.00807 -0.0277 0.2386 0.7357
9.500 1.1827 0.01516 0.00863 -0.0245 0.2187 0.7375
9.750 1.1886 0.01584 0.00927 -0.0212 0.2010 0.7394
10.000 1.1943 0.01659 0.00999 -0.0182 0.1841 0.7414
10.250 1.1982 0.01743 0.01080 -0.0150 0.1678 0.7435
10.500 1.2008 0.01838 0.01172 -0.0120 0.1534 0.7459
10.750 1.2028 0.01944 0.01277 -0.0092 0.1414 0.7482
11.000 1.2031 0.02072 0.01403 -0.0066 0.1297 0.7506
11.250 1.2046 0.02207 0.01539 -0.0045 0.1196 0.7529
11.500 1.2058 0.02357 0.01693 -0.0027 0.1106 0.7553
11.750 1.2054 0.02533 0.01871 -0.0011 0.1023 0.7580
12.000 1.2047 0.02725 0.02066 0.0003 0.0935 0.7608
12.250 1.2050 0.02921 0.02266 0.0014 0.0862 0.7638
12.500 1.2030 0.03146 0.02492 0.0024 0.0793 0.7671
12.750 1.2025 0.03366 0.02715 0.0031 0.0722 0.7701
13.000 1.2017 0.03593 0.02947 0.0038 0.0669 0.7735
13.250 1.1999 0.03835 0.03193 0.0044 0.0609 0.7771
13.500 1.1991 0.04075 0.03438 0.0048 0.0558 0.7807
13.750 1.1972 0.04332 0.03697 0.0051 0.0511 0.7843
14.000 1.1972 0.04575 0.03947 0.0053 0.0468 0.7881
14.500 1.1967 0.05095 0.04478 0.0053 0.0392 0.7970
15.000 1.1969 0.05636 0.05031 0.0049 0.0327 0.8068
15.250 1.1968 0.05918 0.05318 0.0045 0.0303 0.8129
15.500 1.1981 0.06191 0.05601 0.0041 0.0277 0.8192
15.750 1.1986 0.06475 0.05892 0.0036 0.0256 0.8269
16.000 1.1995 0.06763 0.06190 0.0030 0.0239 0.8357
16.250 1.1995 0.07061 0.06497 0.0024 0.0215 0.8468
16.500 1.1995 0.07356 0.06804 0.0019 0.0201 0.8628
16.750 1.1996 0.07648 0.07113 0.0013 0.0184 0.8950
17.000 1.2047 0.08005 0.07486 -0.0013 0.0168 1.0000
17.250 1.2059 0.08324 0.07812 -0.0024 0.0159 1.0000
17.500 1.2049 0.08677 0.08169 -0.0036 0.0143 1.0000
17.750 1.2050 0.09024 0.08524 -0.0049 0.0132 1.0000
18.000 1.2048 0.09378 0.08883 -0.0063 0.0121 1.0000
18.250 1.2029 0.09765 0.09276 -0.0078 0.0114 1.0000
18.500 1.2029 0.10124 0.09644 -0.0093 0.0107 1.0000
18.750 1.2020 0.10504 0.10032 -0.0110 0.0101 1.0000
19.000 1.1986 0.10927 0.10460 -0.0129 0.0090 1.0000
19.250 1.1980 0.11310 0.10851 -0.0146 0.0086 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 548 AIRFOIL (e548-il)