DU 86-137/25 AIRFOIL (du861372-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: DU 86-137/25 AIRFOIL (du861372-il) Reynolds number: 200,000 Max Cl/Cd: 35.47 at α=7.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-du861372-il-200000-n5.txt Download as CSV file: xf-du861372-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: DU 86-137/25 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
------ -------- --------- --------- -------- -------- --------
-14.000 -0.7175 0.10251 0.09701 -0.0333 0.2664 0.0162
-13.750 -0.7325 0.09703 0.09143 -0.0364 0.2664 0.0162
-13.500 -0.7550 0.09185 0.08611 -0.0393 0.2665 0.0162
-13.250 -0.7726 0.08738 0.08152 -0.0409 0.2664 0.0162
-13.000 -0.7960 0.08335 0.07733 -0.0418 0.2665 0.0162
-12.750 -0.8183 0.07998 0.07380 -0.0418 0.2666 0.0162
-12.500 -0.8327 0.07672 0.07040 -0.0412 0.2666 0.0162
-12.250 -0.8573 0.07426 0.06775 -0.0392 0.2667 0.0162
-12.000 -0.8881 0.07269 0.06597 -0.0355 0.2667 0.0163
-11.750 -0.8335 0.06466 0.05811 -0.0401 0.2663 0.0168
-11.500 -0.8392 0.06154 0.05486 -0.0388 0.2664 0.0171
-11.250 -0.8483 0.05887 0.05204 -0.0367 0.2663 0.0173
-11.000 -0.8583 0.05669 0.04968 -0.0339 0.2662 0.0174
-10.750 -0.8682 0.05478 0.04757 -0.0304 0.2663 0.0177
-10.500 -0.8731 0.05281 0.04536 -0.0274 0.2662 0.0180
-10.250 -0.8771 0.05109 0.04334 -0.0243 0.2661 0.0184
-10.000 -0.8807 0.04988 0.04177 -0.0209 0.2661 0.0188
-9.750 -0.8895 0.04974 0.04106 -0.0165 0.2661 0.0191
-9.500 -0.8956 0.04997 0.04067 -0.0122 0.2660 0.0192
-9.250 -0.8653 0.04491 0.03571 -0.0137 0.2659 0.0197
-9.000 -0.8384 0.04132 0.03214 -0.0146 0.2658 0.0205
-8.750 -0.8246 0.03987 0.03040 -0.0131 0.2657 0.0215
-8.500 -0.8195 0.04003 0.03003 -0.0099 0.2657 0.0222
-8.250 -0.8183 0.04126 0.03051 -0.0061 0.2657 0.0225
-8.000 -0.7715 0.03536 0.02507 -0.0099 0.2656 0.0234
-7.750 -0.7473 0.03365 0.02324 -0.0098 0.2655 0.0244
-7.500 -0.7339 0.03355 0.02272 -0.0078 0.2656 0.0256
-7.250 -0.7019 0.03096 0.02015 -0.0089 0.2655 0.0271
-7.000 -0.6813 0.03012 0.01908 -0.0081 0.2655 0.0292
-6.750 -0.6469 0.02766 0.01666 -0.0096 0.2655 0.0331
-6.500 -0.6058 0.02536 0.01465 -0.0121 0.2656 0.0355
-6.250 -0.5799 0.02444 0.01375 -0.0121 0.2656 0.0365
-6.000 -0.5573 0.02379 0.01306 -0.0115 0.2657 0.0373
-5.750 -0.5358 0.02330 0.01250 -0.0108 0.2658 0.0378
-5.500 -0.5170 0.02246 0.01169 -0.0096 0.2659 0.0382
-5.250 -0.5007 0.02172 0.01099 -0.0082 0.2660 0.0386
-5.000 -0.4841 0.02112 0.01036 -0.0067 0.2662 0.0391
-4.750 -0.4654 0.02067 0.00986 -0.0055 0.2663 0.0394
-4.500 -0.4457 0.02029 0.00938 -0.0045 0.2664 0.0398
-4.250 -0.4247 0.01999 0.00898 -0.0036 0.2666 0.0401
-4.000 -0.4028 0.01977 0.00864 -0.0029 0.2668 0.0403
-3.750 -0.3802 0.01960 0.00837 -0.0022 0.2669 0.0404
-3.500 -0.3571 0.01947 0.00814 -0.0017 0.2671 0.0404
-3.250 -0.3335 0.01937 0.00792 -0.0012 0.2672 0.0403
-3.000 -0.3091 0.01932 0.00778 -0.0008 0.2673 0.0399
-2.750 -0.2847 0.01926 0.00763 -0.0004 0.2674 0.0398
-2.500 -0.2602 0.01918 0.00748 -0.0001 0.2674 0.0398
-2.250 -0.2350 0.01915 0.00739 0.0001 0.2674 0.0392
-2.000 -0.2096 0.01911 0.00731 0.0003 0.2676 0.0382
-1.750 -0.1843 0.01904 0.00720 0.0005 0.2677 0.0374
-1.500 -0.1589 0.01895 0.00710 0.0007 0.2680 0.0366
-1.250 -0.1331 0.01889 0.00701 0.0009 0.2684 0.0356
-1.000 -0.1072 0.01885 0.00695 0.0010 0.2687 0.0348
-0.750 -0.0812 0.01883 0.00691 0.0010 0.2691 0.0339
-0.500 -0.0554 0.01878 0.00683 0.0012 0.2695 0.0333
-0.250 -0.0295 0.01875 0.00677 0.0012 0.2699 0.0327
0.000 -0.0034 0.01873 0.00672 0.0013 0.2703 0.0320
0.250 0.0228 0.01874 0.00672 0.0013 0.2705 0.0313
0.500 0.0493 0.01884 0.00679 0.0013 0.2701 0.0306
0.750 0.0760 0.01904 0.00694 0.0012 0.2692 0.0300
1.000 0.1029 0.01925 0.00711 0.0011 0.2685 0.0296
1.250 0.1297 0.01944 0.00725 0.0010 0.2681 0.0294
1.500 0.1563 0.01947 0.00728 0.0010 0.2683 0.0297
1.750 0.1831 0.01955 0.00733 0.0009 0.2682 0.0322
2.000 0.2104 0.01984 0.00760 0.0007 0.2677 0.0327
2.250 0.2370 0.01995 0.00774 0.0007 0.2668 0.0334
2.750 0.2895 0.01990 0.00790 0.0008 0.2645 0.0710
3.000 0.3165 0.02002 0.00807 0.0007 0.2640 0.0752
3.250 0.3434 0.02014 0.00828 0.0007 0.2633 0.0833
3.500 0.3525 0.01826 0.00850 0.0032 0.2628 0.6884
3.750 0.3776 0.01861 0.00904 0.0038 0.2605 0.7375
4.000 0.4043 0.01982 0.01026 0.0037 0.2553 0.7562
4.250 0.4261 0.01955 0.01029 0.0052 0.2497 0.7792
4.500 0.4480 0.02026 0.01114 0.0066 0.2444 0.8048
4.750 0.4674 0.02060 0.01169 0.0088 0.2403 0.8264
5.000 0.4893 0.02053 0.01182 0.0100 0.2347 0.8323
5.250 0.5132 0.02062 0.01205 0.0105 0.2289 0.8353
5.500 0.5363 0.02015 0.01180 0.0112 0.2206 0.8382
5.750 0.5606 0.01961 0.01150 0.0117 0.2088 0.8409
6.000 0.5847 0.01880 0.01084 0.0123 0.1638 0.8437
6.250 0.6032 0.01827 0.01032 0.0138 0.1392 0.8464
6.500 0.6261 0.01859 0.01054 0.0144 0.1093 0.8489
6.750 0.6485 0.01898 0.01079 0.0149 0.0845 0.8515
7.000 0.6707 0.01938 0.01113 0.0154 0.0679 0.8541
7.250 0.6922 0.01977 0.01153 0.0161 0.0569 0.8571
7.500 0.7121 0.02026 0.01200 0.0169 0.0483 0.8604
7.750 0.7321 0.02064 0.01251 0.0179 0.0428 0.8632
8.000 0.7486 0.02119 0.01308 0.0193 0.0374 0.8660
8.250 0.7661 0.02173 0.01372 0.0206 0.0343 0.8690
8.500 0.7834 0.02232 0.01442 0.0219 0.0316 0.8722
8.750 0.7995 0.02300 0.01516 0.0232 0.0295 0.8760
9.000 0.8106 0.02391 0.01612 0.0251 0.0276 0.8801
9.250 0.8222 0.02460 0.01693 0.0273 0.0262 0.8838
9.500 0.8353 0.02528 0.01774 0.0291 0.0247 0.8878
9.750 0.8485 0.02605 0.01860 0.0307 0.0233 0.8922
10.250 0.8718 0.02790 0.02064 0.0339 0.0216 0.9022
10.500 0.8817 0.02903 0.02186 0.0355 0.0209 0.9086
10.750 0.8887 0.03062 0.02353 0.0372 0.0201 0.9153
11.000 0.9019 0.03188 0.02497 0.0382 0.0197 0.9225
11.250 0.9158 0.03328 0.02657 0.0389 0.0192 0.9315
11.750 0.9493 0.03677 0.03054 0.0382 0.0179 0.9683
12.000 0.9582 0.03834 0.03227 0.0388 0.0172 1.0000
12.250 0.9659 0.04007 0.03414 0.0393 0.0167 1.0000
12.500 0.9721 0.04188 0.03607 0.0398 0.0162 1.0000
12.750 0.9764 0.04385 0.03815 0.0402 0.0158 1.0000
13.000 0.9780 0.04613 0.04056 0.0407 0.0155 1.0000
13.250 0.9775 0.04867 0.04322 0.0410 0.0153 1.0000
13.500 0.9733 0.05170 0.04641 0.0412 0.0151 1.0000
13.750 0.9658 0.05518 0.05008 0.0410 0.0150 1.0000
14.000 0.9536 0.05936 0.05445 0.0404 0.0148 1.0000
14.250 0.9384 0.06416 0.05945 0.0391 0.0147 1.0000
14.500 0.9231 0.06933 0.06483 0.0369 0.0147 1.0000
14.750 0.9053 0.07544 0.07114 0.0337 0.0147 1.0000
15.000 0.8835 0.08329 0.07921 0.0288 0.0147 1.0000
|
Polar data table (+)
Polar graphs
<< Back to DU 86-137/25 AIRFOIL (du861372-il)