EPPLER 542 AIRFOIL (e542-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 542 AIRFOIL (e542-il) Reynolds number: 50,000 Max Cl/Cd: 18.33 at α=10.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e542-il-50000-n5.txt Download as CSV file: xf-e542-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 542 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.000 -0.4451 0.09933 0.09227 -0.0645 1.0000 0.0365
-11.750 -0.4559 0.09430 0.08727 -0.0664 1.0000 0.0362
-11.500 -0.4691 0.08960 0.08259 -0.0678 1.0000 0.0358
-11.250 -0.4892 0.08477 0.07776 -0.0687 1.0000 0.0355
-11.000 -0.5099 0.08067 0.07363 -0.0686 1.0000 0.0352
-10.750 -0.5325 0.07702 0.06994 -0.0673 1.0000 0.0350
-10.500 -0.5527 0.07409 0.06698 -0.0652 1.0000 0.0346
-10.250 -0.5774 0.07141 0.06425 -0.0618 1.0000 0.0345
-10.000 -0.6016 0.06950 0.06231 -0.0575 1.0000 0.0344
-9.750 -0.6278 0.06816 0.06096 -0.0520 1.0000 0.0340
-9.500 -0.6592 0.06741 0.06020 -0.0450 1.0000 0.0341
-9.250 -0.6863 0.06668 0.05943 -0.0385 1.0000 0.0339
-9.000 -0.6967 0.06396 0.05647 -0.0358 0.9944 0.0337
-8.750 -0.6905 0.06000 0.05207 -0.0359 0.9835 0.0338
-8.500 -0.6815 0.05635 0.04795 -0.0355 0.9728 0.0339
-8.250 -0.6688 0.05297 0.04409 -0.0349 0.9632 0.0341
-8.000 -0.6480 0.04964 0.04023 -0.0352 0.9559 0.0346
-7.750 -0.6250 0.04674 0.03682 -0.0351 0.9477 0.0354
-7.500 -0.5900 0.04396 0.03362 -0.0369 0.9429 0.0376
-7.250 -0.5599 0.04212 0.03158 -0.0377 0.9357 0.0407
-7.000 -0.5070 0.03976 0.02889 -0.0413 0.9333 0.0447
-6.750 -0.4344 0.03776 0.02675 -0.0474 0.9338 0.0543
-6.500 -0.3788 0.03636 0.02525 -0.0510 0.9315 0.0682
-6.250 -0.3412 0.03496 0.02393 -0.0529 0.9273 0.0892
-6.000 -0.3226 0.03369 0.02286 -0.0520 0.9185 0.1218
-5.750 -0.3088 0.03160 0.02160 -0.0513 0.9119 0.2138
-4.750 -0.1401 0.03956 0.03056 -0.0464 0.8925 0.7945
-4.500 -0.0869 0.03984 0.03045 -0.0499 0.8872 0.8156
-4.250 -0.0492 0.03954 0.02984 -0.0522 0.8814 0.8283
-4.000 -0.0196 0.03927 0.02929 -0.0533 0.8752 0.8393
-3.750 0.0171 0.03883 0.02860 -0.0556 0.8681 0.8465
-3.500 0.0428 0.03856 0.02812 -0.0561 0.8623 0.8556
-3.250 0.0764 0.03814 0.02749 -0.0580 0.8550 0.8613
-3.000 0.1009 0.03792 0.02710 -0.0583 0.8483 0.8687
-2.750 0.1340 0.03748 0.02647 -0.0601 0.8432 0.8740
-2.500 0.1550 0.03740 0.02627 -0.0598 0.8349 0.8799
-2.000 0.2039 0.03705 0.02568 -0.0603 0.8224 0.8902
-1.750 0.2286 0.03690 0.02542 -0.0607 0.8161 0.8948
-1.500 0.2544 0.03671 0.02511 -0.0611 0.8115 0.8993
-1.250 0.2672 0.03691 0.02528 -0.0594 0.8032 0.9036
-1.000 0.2956 0.03670 0.02499 -0.0603 0.7980 0.9070
-0.750 0.3216 0.03656 0.02477 -0.0608 0.7933 0.9105
-0.500 0.3232 0.03711 0.02531 -0.0570 0.7848 0.9151
-0.250 0.3519 0.03691 0.02506 -0.0580 0.7802 0.9177
0.000 0.3719 0.03703 0.02516 -0.0575 0.7741 0.9207
0.250 0.3861 0.03730 0.02543 -0.0560 0.7673 0.9239
0.500 0.4106 0.03724 0.02534 -0.0561 0.7629 0.9265
0.750 0.4081 0.03797 0.02608 -0.0517 0.7549 0.9302
1.000 0.4305 0.03805 0.02617 -0.0516 0.7493 0.9322
1.250 0.4615 0.03785 0.02598 -0.0528 0.7456 0.9340
1.500 0.4601 0.03870 0.02687 -0.0489 0.7366 0.9371
1.750 0.4809 0.03880 0.02698 -0.0483 0.7315 0.9391
2.000 0.4914 0.03921 0.02742 -0.0461 0.7254 0.9414
2.250 0.4893 0.03994 0.02819 -0.0417 0.7173 0.9442
2.500 0.5195 0.03979 0.02809 -0.0426 0.7133 0.9454
2.750 0.5138 0.04079 0.02916 -0.0383 0.7037 0.9477
3.000 0.5393 0.04078 0.02920 -0.0384 0.6986 0.9492
3.500 0.5529 0.04186 0.03041 -0.0329 0.6834 0.9532
3.750 0.5855 0.04154 0.03016 -0.0338 0.6797 0.9541
4.000 0.5600 0.04294 0.03161 -0.0262 0.6678 0.9574
4.250 0.5956 0.04256 0.03132 -0.0275 0.6639 0.9583
4.500 0.5780 0.04381 0.03263 -0.0215 0.6518 0.9605
4.750 0.6160 0.04334 0.03229 -0.0230 0.6478 0.9611
5.000 0.5975 0.04464 0.03365 -0.0171 0.6354 0.9639
5.250 0.6383 0.04399 0.03313 -0.0188 0.6314 0.9648
5.500 0.6178 0.04539 0.03459 -0.0126 0.6185 0.9679
5.750 0.6487 0.04488 0.03421 -0.0127 0.6134 0.9687
6.000 0.6373 0.04595 0.03536 -0.0078 0.6013 0.9709
6.500 0.6754 0.04604 0.03573 -0.0058 0.5839 0.9740
6.750 0.6729 0.04718 0.03699 -0.0029 0.5707 0.9767
7.000 0.7107 0.04607 0.03605 -0.0033 0.5655 0.9779
7.250 0.7087 0.04698 0.03708 -0.0001 0.5524 0.9804
7.500 0.7065 0.04795 0.03817 0.0030 0.5389 0.9827
8.000 0.7543 0.04698 0.03757 0.0046 0.5198 0.9859
8.250 0.7602 0.04770 0.03845 0.0065 0.5055 0.9887
8.500 0.7687 0.04823 0.03914 0.0082 0.4914 0.9917
8.750 0.7808 0.04852 0.03961 0.0096 0.4776 0.9945
9.000 0.7965 0.04850 0.03978 0.0109 0.4639 0.9973
9.250 0.8147 0.04823 0.03969 0.0122 0.4503 1.0000
9.500 0.8273 0.04746 0.03906 0.0152 0.4382 1.0000
9.750 0.8409 0.04656 0.03829 0.0182 0.4253 1.0000
10.000 0.8406 0.04668 0.03850 0.0219 0.4102 1.0000
10.250 0.8426 0.04659 0.03848 0.0256 0.3948 1.0000
10.500 0.8472 0.04631 0.03825 0.0291 0.3788 1.0000
10.750 0.8485 0.04630 0.03827 0.0327 0.3620 1.0000
11.000 0.8470 0.04658 0.03858 0.0363 0.3448 1.0000
11.250 0.8459 0.04699 0.03900 0.0395 0.3270 1.0000
11.500 0.8478 0.04754 0.03953 0.0422 0.3082 1.0000
11.750 0.8525 0.04818 0.04014 0.0443 0.2881 1.0000
12.000 0.8583 0.04902 0.04090 0.0459 0.2683 1.0000
12.250 0.8588 0.05069 0.04259 0.0471 0.2490 1.0000
12.500 0.8606 0.05241 0.04431 0.0480 0.2302 1.0000
12.750 0.8628 0.05425 0.04612 0.0487 0.2124 1.0000
13.000 0.8648 0.05626 0.04809 0.0492 0.1954 1.0000
13.250 0.8663 0.05845 0.05025 0.0495 0.1795 1.0000
13.500 0.8677 0.06082 0.05262 0.0496 0.1648 1.0000
13.750 0.8688 0.06336 0.05518 0.0495 0.1511 1.0000
14.000 0.8696 0.06606 0.05792 0.0493 0.1385 1.0000
14.250 0.8707 0.06885 0.06077 0.0490 0.1271 1.0000
14.500 0.8723 0.07162 0.06354 0.0486 0.1170 1.0000
14.750 0.8746 0.07435 0.06627 0.0481 0.1079 1.0000
15.000 0.8740 0.07778 0.06989 0.0473 0.0996 1.0000
15.250 0.8762 0.08072 0.07282 0.0467 0.0923 1.0000
15.500 0.8752 0.08436 0.07664 0.0456 0.0858 1.0000
15.750 0.8774 0.08750 0.07983 0.0448 0.0803 1.0000
16.000 0.8734 0.09187 0.08445 0.0433 0.0757 1.0000
16.250 0.8794 0.09438 0.08688 0.0426 0.0708 1.0000
16.500 0.8683 0.10020 0.09306 0.0402 0.0682 1.0000
16.750 0.8575 0.10607 0.09917 0.0376 0.0657 1.0000
17.000 0.8554 0.11027 0.10345 0.0356 0.0625 1.0000
17.250 0.8498 0.11531 0.10857 0.0332 0.0601 1.0000
17.500 0.8279 0.12405 0.11756 0.0285 0.0597 1.0000
17.750 0.8003 0.13474 0.12841 0.0226 0.0598 1.0000
18.000 0.7687 0.14758 0.14129 0.0155 0.0601 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 542 AIRFOIL (e542-il)