DU 86-137/25 AIRFOIL (du861372-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: DU 86-137/25 AIRFOIL (du861372-il) Reynolds number: 200,000 Max Cl/Cd: 32.77 at α=7.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-du861372-il-200000.txt Download as CSV file: xf-du861372-il-200000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-13.500 -0.6830 0.10284 0.09782 -0.0346 0.3565 0.0411
-13.250 -0.6924 0.09677 0.09172 -0.0378 0.3566 0.0410
-13.000 -0.7009 0.09133 0.08625 -0.0404 0.3566 0.0407
-12.750 -0.7255 0.08454 0.07937 -0.0440 0.3568 0.0405
-12.500 -0.7424 0.07937 0.07412 -0.0459 0.3568 0.0402
-12.250 -0.7596 0.07464 0.06930 -0.0471 0.3568 0.0396
-12.000 -0.7853 0.06998 0.06449 -0.0475 0.3570 0.0396
-11.750 -0.8066 0.06608 0.06043 -0.0468 0.3570 0.0395
-11.500 -0.8275 0.06268 0.05685 -0.0452 0.3573 0.0396
-11.250 -0.8537 0.05976 0.05368 -0.0422 0.3573 0.0402
-11.000 -0.8733 0.05728 0.05099 -0.0384 0.3574 0.0403
-10.750 -0.8970 0.05525 0.04862 -0.0333 0.3576 0.0408
-10.500 -0.8695 0.05400 0.04760 -0.0346 0.3578 0.0447
-10.250 -0.8769 0.05172 0.04503 -0.0314 0.3580 0.0457
-10.000 -0.8887 0.04980 0.04265 -0.0271 0.3581 0.0468
-9.750 -0.8640 0.04860 0.04167 -0.0280 0.3583 0.0522
-9.500 -0.8660 0.04718 0.03982 -0.0245 0.3585 0.0543
-9.250 -0.8477 0.04588 0.03869 -0.0247 0.3587 0.0622
-9.000 -0.8408 0.04368 0.03613 -0.0223 0.3589 0.0633
-8.750 -0.8298 0.04144 0.03354 -0.0202 0.3590 0.0639
-8.500 -0.8156 0.03938 0.03114 -0.0184 0.3592 0.0644
-8.250 -0.8005 0.03842 0.02980 -0.0164 0.3594 0.0653
-8.000 -0.7831 0.03536 0.02641 -0.0153 0.3595 0.0664
-7.750 -0.7596 0.03308 0.02390 -0.0147 0.3597 0.0666
-7.500 -0.7362 0.03231 0.02290 -0.0137 0.3601 0.0659
-7.250 -0.7047 0.02957 0.02006 -0.0143 0.3604 0.0647
-7.000 -0.6736 0.02750 0.01788 -0.0147 0.3608 0.0639
-6.750 -0.6421 0.02587 0.01613 -0.0150 0.3611 0.0630
-6.500 -0.6108 0.02450 0.01468 -0.0154 0.3613 0.0621
-6.250 -0.5805 0.02329 0.01345 -0.0156 0.3615 0.0610
-6.000 -0.5539 0.02224 0.01238 -0.0152 0.3616 0.0596
-5.750 -0.5321 0.02135 0.01145 -0.0141 0.3619 0.0582
-5.500 -0.5134 0.02059 0.01067 -0.0127 0.3622 0.0578
-5.250 -0.4960 0.01989 0.00995 -0.0112 0.3624 0.0584
-5.000 -0.4781 0.01927 0.00927 -0.0097 0.3627 0.0586
-4.750 -0.4592 0.01871 0.00862 -0.0083 0.3630 0.0587
-4.500 -0.4388 0.01824 0.00805 -0.0072 0.3633 0.0599
-4.250 -0.4173 0.01784 0.00756 -0.0062 0.3636 0.0628
-4.000 -0.3948 0.01752 0.00714 -0.0054 0.3640 0.0674
-3.750 -0.3724 0.01719 0.00674 -0.0046 0.3644 0.0738
-3.500 -0.3484 0.01698 0.00644 -0.0041 0.3647 0.0793
-3.250 -0.3257 0.01659 0.00617 -0.0035 0.3652 0.0957
-3.000 -0.3006 0.01640 0.00595 -0.0032 0.3659 0.1214
-2.750 -0.2735 0.01638 0.00578 -0.0031 0.3666 0.0905
-2.500 -0.2467 0.01635 0.00562 -0.0030 0.3673 0.0861
-2.250 -0.2201 0.01630 0.00549 -0.0029 0.3682 0.0821
-2.000 -0.1938 0.01624 0.00536 -0.0028 0.3690 0.0787
-1.750 -0.1676 0.01616 0.00525 -0.0026 0.3700 0.0775
-1.500 -0.1421 0.01602 0.00512 -0.0024 0.3710 0.0847
-1.250 -0.1153 0.01601 0.00503 -0.0024 0.3721 0.0779
-1.000 -0.0889 0.01596 0.00495 -0.0023 0.3730 0.0761
-0.750 -0.0626 0.01590 0.00486 -0.0022 0.3738 0.0754
-0.500 -0.0363 0.01584 0.00478 -0.0021 0.3745 0.0763
-0.250 -0.0111 0.01567 0.00471 -0.0019 0.3753 0.0946
0.000 0.0147 0.01552 0.00464 -0.0018 0.3754 0.1353
0.250 0.0360 0.01483 0.00441 -0.0012 0.3749 0.2567
0.500 0.0412 0.01309 0.00455 0.0025 0.3746 0.6731
0.750 0.0678 0.01313 0.00467 0.0027 0.3742 0.6946
1.000 0.0945 0.01321 0.00477 0.0029 0.3747 0.7107
1.250 0.1210 0.01331 0.00491 0.0032 0.3752 0.7269
1.500 0.1473 0.01345 0.00509 0.0035 0.3757 0.7430
1.750 0.1731 0.01362 0.00533 0.0041 0.3759 0.7594
2.000 0.1973 0.01393 0.00575 0.0053 0.3756 0.7803
2.250 0.2219 0.01414 0.00606 0.0063 0.3749 0.7930
2.500 0.2446 0.01443 0.00646 0.0080 0.3743 0.8086
2.750 0.2645 0.01479 0.00698 0.0106 0.3730 0.8266
3.000 0.2868 0.01495 0.00723 0.0122 0.3717 0.8390
3.250 0.3047 0.01520 0.00762 0.0153 0.3690 0.8554
3.500 0.3185 0.01545 0.00804 0.0200 0.3632 0.8755
3.750 0.3387 0.01543 0.00810 0.0220 0.3268 0.8880
4.000 0.3547 0.01584 0.00815 0.0238 0.3195 0.8915
4.250 0.3734 0.01659 0.00866 0.0249 0.3138 0.8950
4.500 0.4082 0.01822 0.00977 0.0230 0.3051 0.8973
4.750 0.4401 0.01953 0.01099 0.0216 0.2955 0.9004
5.000 0.4658 0.02056 0.01209 0.0213 0.2837 0.9036
5.250 0.4905 0.02178 0.01342 0.0213 0.2694 0.9061
5.500 0.5106 0.02206 0.01393 0.0223 0.2566 0.9098
5.750 0.5301 0.02217 0.01426 0.0234 0.2454 0.9140
6.000 0.5485 0.02217 0.01443 0.0245 0.2364 0.9182
6.250 0.5630 0.02094 0.01359 0.0269 0.2213 0.9222
6.500 0.5815 0.01966 0.01263 0.0287 0.1880 0.9269
6.750 0.5921 0.01864 0.01184 0.0314 0.1724 0.9332
7.000 0.6086 0.01862 0.01182 0.0329 0.1462 0.9378
7.250 0.6240 0.01907 0.01219 0.0343 0.1240 0.9427
7.500 0.6394 0.01951 0.01257 0.0357 0.1061 0.9486
7.750 0.6522 0.02014 0.01314 0.0373 0.0934 0.9547
8.000 0.6684 0.02091 0.01390 0.0380 0.0822 0.9608
8.250 0.6861 0.02176 0.01481 0.0385 0.0736 0.9675
8.500 0.7048 0.02334 0.01632 0.0378 0.0660 0.9737
8.750 0.7345 0.02410 0.01720 0.0364 0.0597 0.9790
9.000 0.7596 0.02541 0.01845 0.0350 0.0547 0.9848
9.250 0.7888 0.02685 0.01999 0.0339 0.0508 0.9900
9.500 0.8200 0.02811 0.02135 0.0326 0.0479 0.9974
9.750 0.8351 0.02899 0.02227 0.0338 0.0457 1.0000
10.000 0.8551 0.03039 0.02362 0.0341 0.0431 1.0000
10.250 0.8868 0.03278 0.02622 0.0329 0.0409 1.0000
10.500 0.9100 0.03470 0.02840 0.0328 0.0397 1.0000
10.750 0.9290 0.03695 0.03092 0.0331 0.0386 1.0000
11.000 0.9420 0.03948 0.03374 0.0338 0.0378 1.0000
11.250 0.9493 0.04198 0.03651 0.0350 0.0370 1.0000
11.500 0.9565 0.04378 0.03843 0.0361 0.0358 1.0000
11.750 0.9625 0.04550 0.04024 0.0370 0.0347 1.0000
12.000 0.9633 0.04791 0.04279 0.0381 0.0340 1.0000
12.250 0.9560 0.05118 0.04629 0.0394 0.0338 1.0000
12.500 0.9438 0.05471 0.05006 0.0405 0.0338 1.0000
12.750 0.9232 0.05897 0.05462 0.0412 0.0340 1.0000
13.000 0.8588 0.06819 0.06439 0.0404 0.0356 1.0000
13.250 0.7903 0.08085 0.07741 0.0344 0.0372 1.0000
13.500 0.7237 0.09867 0.09535 0.0232 0.0404 1.0000
13.750 0.7023 0.10840 0.10505 0.0182 0.0417 1.0000
|
Polar data table (+)
Polar graphs
<< Back to DU 86-137/25 AIRFOIL (du861372-il)