EPPLER 547 AIRFOIL (e547-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 547 AIRFOIL (e547-il) Reynolds number: 200,000 Max Cl/Cd: 62 at α=8.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e547-il-200000-n5.txt Download as CSV file: xf-e547-il-200000-n5.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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-13.750 -0.5198 0.09170 0.08781 -0.0573 1.0000 0.0085
-13.500 -0.5417 0.08251 0.07851 -0.0631 1.0000 0.0084
-13.250 -0.5654 0.07481 0.07066 -0.0674 1.0000 0.0083
-13.000 -0.5893 0.06847 0.06416 -0.0701 1.0000 0.0083
-12.750 -0.6081 0.06364 0.05919 -0.0713 1.0000 0.0082
-12.500 -0.6306 0.05881 0.05416 -0.0716 1.0000 0.0082
-12.250 -0.6496 0.05475 0.04989 -0.0710 1.0000 0.0082
-12.000 -0.6604 0.05187 0.04687 -0.0699 1.0000 0.0082
-11.750 -0.6759 0.04858 0.04335 -0.0681 1.0000 0.0082
-11.500 -0.6907 0.04544 0.03996 -0.0657 1.0000 0.0082
-11.250 -0.6990 0.04312 0.03744 -0.0632 1.0000 0.0082
-11.000 -0.7068 0.04091 0.03503 -0.0603 1.0000 0.0083
-10.750 -0.7129 0.03905 0.03298 -0.0570 1.0000 0.0083
-10.500 -0.6963 0.03633 0.02994 -0.0580 0.9897 0.0083
-10.250 -0.6765 0.03391 0.02719 -0.0586 0.9754 0.0085
-10.000 -0.6521 0.03191 0.02505 -0.0600 0.9625 0.0088
-9.750 -0.6217 0.03028 0.02327 -0.0622 0.9520 0.0091
-9.500 -0.5874 0.02866 0.02146 -0.0647 0.9427 0.0092
-9.250 -0.5526 0.02722 0.01987 -0.0672 0.9315 0.0096
-9.000 -0.5170 0.02588 0.01836 -0.0698 0.9198 0.0103
-8.750 -0.4816 0.02471 0.01700 -0.0722 0.9068 0.0112
-8.500 -0.4516 0.02365 0.01589 -0.0741 0.8918 0.0123
-8.250 -0.4245 0.02277 0.01490 -0.0751 0.8765 0.0131
-8.000 -0.4021 0.02193 0.01392 -0.0751 0.8618 0.0147
-7.750 -0.3839 0.02113 0.01304 -0.0743 0.8478 0.0158
-7.500 -0.3676 0.02044 0.01224 -0.0731 0.8345 0.0170
-7.250 -0.3530 0.01977 0.01145 -0.0715 0.8219 0.0186
-7.000 -0.3383 0.01917 0.01079 -0.0699 0.8104 0.0222
-6.750 -0.3248 0.01856 0.01012 -0.0680 0.7997 0.0264
-6.500 -0.3133 0.01795 0.00948 -0.0657 0.7890 0.0324
-6.250 -0.3033 0.01740 0.00892 -0.0632 0.7791 0.0435
-5.750 -0.2847 0.01631 0.00797 -0.0577 0.7606 0.0854
-5.500 -0.2746 0.01569 0.00749 -0.0552 0.7524 0.1252
-5.250 -0.2674 0.01491 0.00699 -0.0523 0.7436 0.1868
-5.000 -0.2648 0.01386 0.00638 -0.0488 0.7359 0.2877
-4.750 -0.2678 0.01247 0.00576 -0.0442 0.7279 0.4465
-4.500 -0.2511 0.01266 0.00662 -0.0412 0.7210 0.6029
-4.250 -0.2264 0.01285 0.00669 -0.0405 0.7134 0.6333
-4.000 -0.2013 0.01303 0.00668 -0.0399 0.7069 0.6527
-3.750 -0.1752 0.01343 0.00699 -0.0392 0.6996 0.6687
-3.500 -0.1484 0.01398 0.00744 -0.0383 0.6934 0.6839
-3.250 -0.1218 0.01463 0.00803 -0.0373 0.6864 0.6983
-3.000 -0.0920 0.01527 0.00863 -0.0365 0.6798 0.7045
-2.750 -0.0657 0.01525 0.00849 -0.0364 0.6737 0.7079
-2.500 -0.0405 0.01508 0.00819 -0.0363 0.6672 0.7113
-2.250 -0.0149 0.01488 0.00782 -0.0363 0.6617 0.7145
-2.000 0.0119 0.01483 0.00771 -0.0363 0.6551 0.7157
-1.750 0.0388 0.01479 0.00758 -0.0362 0.6490 0.7171
-1.500 0.0658 0.01474 0.00743 -0.0363 0.6436 0.7184
-1.250 0.0924 0.01469 0.00731 -0.0363 0.6375 0.7200
-1.000 0.1193 0.01463 0.00716 -0.0363 0.6322 0.7217
-0.750 0.1460 0.01456 0.00701 -0.0364 0.6265 0.7233
-0.500 0.1726 0.01447 0.00686 -0.0365 0.6206 0.7251
-0.250 0.1998 0.01441 0.00668 -0.0368 0.6159 0.7272
0.000 0.2266 0.01432 0.00653 -0.0370 0.6105 0.7292
0.250 0.2535 0.01428 0.00644 -0.0372 0.6049 0.7303
0.500 0.2809 0.01427 0.00637 -0.0373 0.6003 0.7313
0.750 0.3077 0.01427 0.00637 -0.0373 0.5950 0.7323
1.000 0.3345 0.01428 0.00636 -0.0374 0.5899 0.7335
1.250 0.3618 0.01431 0.00634 -0.0375 0.5854 0.7349
1.500 0.3886 0.01432 0.00635 -0.0376 0.5804 0.7362
1.750 0.4154 0.01433 0.00635 -0.0377 0.5753 0.7375
2.000 0.4427 0.01435 0.00633 -0.0379 0.5708 0.7389
2.250 0.4698 0.01438 0.00635 -0.0381 0.5663 0.7404
2.500 0.4965 0.01440 0.00638 -0.0382 0.5611 0.7419
2.750 0.5238 0.01443 0.00639 -0.0384 0.5565 0.7435
3.000 0.5514 0.01448 0.00639 -0.0388 0.5523 0.7453
3.250 0.5772 0.01454 0.00652 -0.0386 0.5470 0.7464
3.500 0.6036 0.01460 0.00661 -0.0386 0.5422 0.7476
4.000 0.6562 0.01475 0.00684 -0.0386 0.5324 0.7500
4.250 0.6823 0.01483 0.00694 -0.0385 0.5272 0.7513
4.500 0.7092 0.01492 0.00702 -0.0386 0.5225 0.7527
4.750 0.7340 0.01500 0.00720 -0.0384 0.5162 0.7542
5.000 0.7598 0.01508 0.00730 -0.0383 0.5103 0.7558
5.250 0.7854 0.01518 0.00744 -0.0382 0.5044 0.7576
5.500 0.8102 0.01527 0.00759 -0.0380 0.4976 0.7596
5.750 0.8359 0.01538 0.00769 -0.0379 0.4916 0.7613
6.000 0.8590 0.01548 0.00793 -0.0373 0.4842 0.7627
6.250 0.8833 0.01560 0.00807 -0.0369 0.4777 0.7641
6.500 0.9062 0.01572 0.00831 -0.0363 0.4700 0.7657
6.750 0.9295 0.01586 0.00847 -0.0357 0.4625 0.7674
7.000 0.9514 0.01599 0.00873 -0.0350 0.4535 0.7692
7.250 0.9733 0.01615 0.00891 -0.0342 0.4448 0.7712
7.500 0.9938 0.01631 0.00915 -0.0332 0.4344 0.7733
7.750 1.0138 0.01649 0.00941 -0.0322 0.4235 0.7754
8.000 1.0322 0.01669 0.00967 -0.0308 0.4118 0.7774
8.250 1.0485 0.01691 0.00995 -0.0290 0.3990 0.7794
8.500 1.0623 0.01716 0.01025 -0.0268 0.3848 0.7818
8.750 1.0735 0.01745 0.01058 -0.0242 0.3695 0.7845
9.000 1.0830 0.01786 0.01101 -0.0215 0.3512 0.7872
9.250 1.0909 0.01841 0.01153 -0.0187 0.3313 0.7901
9.500 1.0958 0.01909 0.01217 -0.0156 0.3109 0.7928
9.750 1.0990 0.01989 0.01295 -0.0124 0.2894 0.7955
10.000 1.0996 0.02089 0.01390 -0.0092 0.2687 0.7986
10.250 1.1010 0.02199 0.01496 -0.0064 0.2487 0.8021
10.500 1.1015 0.02325 0.01618 -0.0038 0.2293 0.8059
10.750 1.1014 0.02461 0.01754 -0.0013 0.2119 0.8094
11.000 1.1009 0.02610 0.01902 0.0010 0.1957 0.8134
11.250 1.1008 0.02771 0.02062 0.0030 0.1799 0.8179
11.500 1.1014 0.02940 0.02230 0.0046 0.1650 0.8227
11.750 1.1019 0.03116 0.02408 0.0061 0.1510 0.8271
12.250 1.1041 0.03495 0.02790 0.0084 0.1260 0.8378
12.500 1.1053 0.03692 0.02991 0.0093 0.1148 0.8436
12.750 1.1064 0.03898 0.03202 0.0101 0.1048 0.8507
13.000 1.1068 0.04117 0.03426 0.0108 0.0956 0.8587
13.500 1.1101 0.04551 0.03877 0.0117 0.0800 0.8825
13.750 1.1105 0.04799 0.04134 0.0119 0.0729 0.9058
14.000 1.1141 0.05025 0.04373 0.0113 0.0659 1.0000
14.250 1.1149 0.05296 0.04646 0.0111 0.0603 1.0000
14.500 1.1172 0.05557 0.04912 0.0108 0.0548 1.0000
14.750 1.1173 0.05848 0.05205 0.0104 0.0502 1.0000
15.000 1.1193 0.06122 0.05486 0.0099 0.0458 1.0000
15.500 1.1199 0.06731 0.06108 0.0087 0.0390 1.0000
15.750 1.1194 0.07055 0.06436 0.0080 0.0358 1.0000
16.000 1.1172 0.07405 0.06791 0.0071 0.0335 1.0000
16.250 1.1184 0.07720 0.07118 0.0062 0.0310 1.0000
16.500 1.1173 0.08073 0.07477 0.0051 0.0289 1.0000
16.750 1.1129 0.08476 0.07885 0.0038 0.0272 1.0000
17.000 1.1134 0.08820 0.08241 0.0027 0.0255 1.0000
17.250 1.1120 0.09194 0.08625 0.0013 0.0236 1.0000
17.500 1.1082 0.09611 0.09049 -0.0003 0.0224 1.0000
17.750 1.1042 0.10040 0.09486 -0.0021 0.0211 1.0000
18.000 1.1023 0.10444 0.09902 -0.0037 0.0197 1.0000
18.250 1.0996 0.10865 0.10336 -0.0056 0.0186 1.0000
18.500 1.0947 0.11327 0.10807 -0.0078 0.0176 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 547 AIRFOIL (e547-il)