EPPLER 542 AIRFOIL (e542-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 542 AIRFOIL (e542-il) Reynolds number: 100,000 Max Cl/Cd: 32.91 at α=9.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e542-il-100000-n5.txt Download as CSV file: xf-e542-il-100000-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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.250 -0.4528 0.09231 0.08725 -0.0641 1.0000 0.0179
-12.000 -0.4690 0.08565 0.08057 -0.0676 1.0000 0.0177
-11.750 -0.4890 0.07981 0.07468 -0.0700 1.0000 0.0176
-11.500 -0.5095 0.07502 0.06983 -0.0710 1.0000 0.0174
-11.250 -0.5338 0.07054 0.06526 -0.0706 1.0000 0.0173
-11.000 -0.5557 0.06699 0.06162 -0.0691 1.0000 0.0172
-10.750 -0.5789 0.06380 0.05833 -0.0664 1.0000 0.0171
-10.500 -0.5992 0.06146 0.05592 -0.0630 1.0000 0.0170
-10.250 -0.6257 0.05944 0.05382 -0.0579 1.0000 0.0169
-10.000 -0.6377 0.05638 0.05055 -0.0561 0.9927 0.0168
-9.750 -0.6400 0.05251 0.04632 -0.0558 0.9775 0.0167
-9.500 -0.6407 0.04877 0.04216 -0.0545 0.9626 0.0167
-9.250 -0.6317 0.04529 0.03824 -0.0540 0.9514 0.0166
-9.000 -0.6141 0.04172 0.03417 -0.0542 0.9437 0.0167
-8.750 -0.5928 0.03868 0.03067 -0.0543 0.9350 0.0169
-8.500 -0.5619 0.03571 0.02721 -0.0556 0.9294 0.0173
-8.250 -0.5201 0.03294 0.02398 -0.0583 0.9261 0.0179
-8.000 -0.4875 0.03137 0.02232 -0.0599 0.9180 0.0193
-7.750 -0.4452 0.02966 0.02033 -0.0626 0.9134 0.0215
-7.500 -0.4089 0.02826 0.01889 -0.0647 0.9070 0.0236
-7.250 -0.3749 0.02693 0.01740 -0.0659 0.8994 0.0258
-7.000 -0.3454 0.02575 0.01618 -0.0669 0.8918 0.0285
-6.750 -0.3212 0.02483 0.01520 -0.0668 0.8819 0.0339
-6.500 -0.2979 0.02390 0.01422 -0.0665 0.8729 0.0411
-6.250 -0.2756 0.02302 0.01336 -0.0660 0.8640 0.0544
-6.000 -0.2546 0.02214 0.01266 -0.0652 0.8554 0.0855
-5.750 -0.2339 0.02117 0.01220 -0.0644 0.8473 0.1717
-5.500 -0.2297 0.01997 0.01183 -0.0611 0.8379 0.3188
-5.250 -0.1629 0.02316 0.01648 -0.0612 0.8349 0.6546
-5.000 -0.1511 0.02351 0.01663 -0.0582 0.8268 0.6892
-4.750 -0.1407 0.02401 0.01698 -0.0547 0.8176 0.7124
-4.500 -0.1199 0.02460 0.01734 -0.0528 0.8113 0.7312
-4.250 -0.0802 0.02591 0.01846 -0.0532 0.8045 0.7487
-4.000 -0.0333 0.02704 0.01937 -0.0550 0.7990 0.7655
-3.750 -0.0147 0.02718 0.01933 -0.0533 0.7930 0.7771
-3.500 0.0217 0.02721 0.01919 -0.0549 0.7867 0.7807
-3.250 0.0504 0.02712 0.01893 -0.0552 0.7811 0.7858
-3.000 0.0613 0.02699 0.01868 -0.0526 0.7748 0.7936
-2.750 0.0932 0.02689 0.01846 -0.0536 0.7688 0.7965
-2.500 0.1219 0.02674 0.01817 -0.0540 0.7638 0.8002
-2.250 0.1229 0.02664 0.01800 -0.0497 0.7578 0.8086
-2.000 0.1547 0.02651 0.01778 -0.0507 0.7524 0.8106
-1.750 0.1851 0.02634 0.01750 -0.0515 0.7477 0.8131
-1.500 0.2108 0.02621 0.01726 -0.0515 0.7431 0.8164
-1.250 0.2206 0.02615 0.01718 -0.0487 0.7371 0.8215
-1.000 0.2351 0.02602 0.01699 -0.0467 0.7323 0.8256
-0.750 0.2664 0.02583 0.01669 -0.0477 0.7287 0.8274
-0.500 0.2887 0.02580 0.01666 -0.0471 0.7230 0.8299
-0.250 0.3089 0.02572 0.01655 -0.0461 0.7180 0.8328
0.000 0.3241 0.02561 0.01639 -0.0442 0.7140 0.8365
0.250 0.3276 0.02559 0.01635 -0.0401 0.7092 0.8408
0.500 0.3524 0.02555 0.01632 -0.0401 0.7040 0.8424
0.750 0.3784 0.02546 0.01621 -0.0401 0.6998 0.8442
1.000 0.4035 0.02532 0.01604 -0.0400 0.6964 0.8460
1.250 0.4151 0.02540 0.01616 -0.0375 0.6908 0.8488
1.500 0.4182 0.02543 0.01620 -0.0333 0.6857 0.8526
1.750 0.4200 0.02533 0.01608 -0.0288 0.6816 0.8559
2.000 0.4465 0.02527 0.01603 -0.0289 0.6776 0.8572
2.250 0.4616 0.02536 0.01619 -0.0271 0.6717 0.8589
2.500 0.4817 0.02530 0.01616 -0.0260 0.6672 0.8603
2.750 0.5071 0.02516 0.01601 -0.0259 0.6637 0.8618
3.000 0.5105 0.02532 0.01625 -0.0219 0.6573 0.8647
3.250 0.5184 0.02527 0.01623 -0.0186 0.6520 0.8669
3.500 0.5350 0.02508 0.01603 -0.0169 0.6480 0.8687
3.750 0.5332 0.02517 0.01621 -0.0119 0.6414 0.8718
4.000 0.5530 0.02513 0.01622 -0.0108 0.6356 0.8729
4.250 0.5823 0.02492 0.01602 -0.0113 0.6315 0.8737
4.500 0.5858 0.02508 0.01629 -0.0074 0.6240 0.8755
4.750 0.6055 0.02497 0.01625 -0.0062 0.6183 0.8768
5.000 0.6315 0.02478 0.01610 -0.0061 0.6134 0.8782
5.250 0.6295 0.02493 0.01635 -0.0012 0.6051 0.8812
5.500 0.6565 0.02465 0.01612 -0.0012 0.5998 0.8823
5.750 0.6552 0.02474 0.01630 0.0035 0.5916 0.8846
6.000 0.6764 0.02452 0.01614 0.0044 0.5850 0.8859
6.250 0.6883 0.02453 0.01626 0.0069 0.5768 0.8873
6.500 0.7107 0.02436 0.01618 0.0077 0.5689 0.8885
6.750 0.7198 0.02441 0.01634 0.0107 0.5598 0.8906
7.000 0.7458 0.02413 0.01614 0.0109 0.5515 0.8919
7.250 0.7437 0.02426 0.01640 0.0157 0.5409 0.8943
7.500 0.7606 0.02409 0.01631 0.0174 0.5312 0.8958
7.750 0.7726 0.02403 0.01633 0.0197 0.5202 0.8977
8.000 0.7761 0.02426 0.01668 0.0231 0.5076 0.8999
8.250 0.7861 0.02444 0.01696 0.0255 0.4942 0.9017
8.500 0.7966 0.02461 0.01724 0.0278 0.4797 0.9037
8.750 0.8072 0.02483 0.01755 0.0300 0.4634 0.9060
9.000 0.8185 0.02508 0.01785 0.0320 0.4452 0.9083
9.250 0.8324 0.02529 0.01808 0.0336 0.4249 0.9103
9.500 0.8412 0.02583 0.01864 0.0355 0.4022 0.9125
9.750 0.8526 0.02633 0.01907 0.0371 0.3776 0.9145
10.000 0.8598 0.02712 0.01983 0.0389 0.3526 0.9166
10.250 0.8659 0.02802 0.02067 0.0408 0.3282 0.9187
10.500 0.8707 0.02910 0.02172 0.0425 0.3043 0.9211
10.750 0.8748 0.03032 0.02289 0.0441 0.2816 0.9237
11.000 0.8784 0.03167 0.02421 0.0456 0.2593 0.9267
11.250 0.8815 0.03313 0.02563 0.0469 0.2388 0.9299
11.500 0.8850 0.03463 0.02713 0.0480 0.2187 0.9329
11.750 0.8883 0.03626 0.02876 0.0489 0.1991 0.9360
12.000 0.8914 0.03801 0.03050 0.0497 0.1808 0.9392
12.250 0.8947 0.03985 0.03232 0.0502 0.1640 0.9425
12.500 0.8982 0.04180 0.03426 0.0505 0.1479 0.9458
12.750 0.9031 0.04379 0.03628 0.0505 0.1327 0.9494
13.000 0.9076 0.04593 0.03845 0.0503 0.1187 0.9535
13.500 0.9172 0.05060 0.04320 0.0491 0.0950 0.9634
13.750 0.9229 0.05299 0.04567 0.0482 0.0850 0.9703
14.250 0.9260 0.05812 0.05087 0.0471 0.0705 1.0000
14.500 0.9299 0.06067 0.05355 0.0465 0.0641 1.0000
14.750 0.9308 0.06357 0.05647 0.0458 0.0592 1.0000
15.000 0.9332 0.06643 0.05946 0.0450 0.0542 1.0000
15.250 0.9338 0.06955 0.06261 0.0441 0.0503 1.0000
15.500 0.9359 0.07258 0.06577 0.0432 0.0464 1.0000
15.750 0.9354 0.07594 0.06913 0.0420 0.0437 1.0000
16.000 0.9378 0.07912 0.07250 0.0410 0.0403 1.0000
16.250 0.9378 0.08259 0.07603 0.0396 0.0380 1.0000
16.500 0.9378 0.08615 0.07971 0.0382 0.0356 1.0000
16.750 0.9377 0.08984 0.08354 0.0367 0.0333 1.0000
17.000 0.9357 0.09380 0.08755 0.0349 0.0315 1.0000
17.250 0.9337 0.09798 0.09191 0.0330 0.0295 1.0000
17.500 0.9310 0.10230 0.09640 0.0309 0.0278 1.0000
17.750 0.9281 0.10665 0.10080 0.0286 0.0265 1.0000
18.000 0.9242 0.11136 0.10566 0.0262 0.0252 1.0000
18.250 0.9191 0.11642 0.11093 0.0235 0.0240 1.0000
18.500 0.9133 0.12166 0.11630 0.0205 0.0229 1.0000
18.750 0.9107 0.12626 0.12096 0.0178 0.0221 1.0000
19.000 0.9057 0.13144 0.12626 0.0148 0.0213 1.0000
19.250 0.8944 0.13827 0.13331 0.0108 0.0209 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 542 AIRFOIL (e542-il)