OAF117 AIRFOIL (oaf117-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: OAF117 AIRFOIL (oaf117-il) Reynolds number: 100,000 Max Cl/Cd: 49.45 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-oaf117-il-100000-n5.txt Download as CSV file: xf-oaf117-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: OAF117 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
------ -------- --------- --------- -------- -------- --------
-11.250 -0.7028 0.10383 0.09854 0.0109 1.0000 0.0463
-11.000 -0.8126 0.07316 0.06774 -0.0111 1.0000 0.0454
-10.750 -0.8620 0.05879 0.05310 -0.0264 1.0000 0.0446
-10.500 -0.8806 0.05004 0.04380 -0.0364 1.0000 0.0450
-10.250 -0.8817 0.04409 0.03715 -0.0410 1.0000 0.0461
-10.000 -0.8629 0.04224 0.03531 -0.0415 1.0000 0.0472
-9.750 -0.8442 0.04004 0.03295 -0.0426 1.0000 0.0485
-9.500 -0.8248 0.03749 0.03011 -0.0437 1.0000 0.0501
-9.250 -0.8038 0.03492 0.02711 -0.0450 1.0000 0.0522
-9.000 -0.7807 0.03288 0.02483 -0.0458 1.0000 0.0542
-8.750 -0.7560 0.03149 0.02340 -0.0463 1.0000 0.0561
-8.500 -0.7307 0.02991 0.02164 -0.0469 1.0000 0.0583
-8.250 -0.7043 0.02831 0.01969 -0.0475 1.0000 0.0612
-8.000 -0.6782 0.02707 0.01848 -0.0480 1.0000 0.0637
-7.750 -0.6512 0.02592 0.01726 -0.0485 1.0000 0.0666
-7.500 -0.6234 0.02475 0.01587 -0.0489 1.0000 0.0703
-7.250 -0.5959 0.02373 0.01488 -0.0494 1.0000 0.0739
-7.000 -0.5677 0.02279 0.01387 -0.0499 1.0000 0.0783
-6.750 -0.5393 0.02186 0.01288 -0.0504 1.0000 0.0832
-6.500 -0.5105 0.02108 0.01208 -0.0509 1.0000 0.0892
-6.250 -0.4816 0.02029 0.01128 -0.0515 1.0000 0.0959
-6.000 -0.4523 0.01961 0.01054 -0.0520 1.0000 0.1049
-5.750 -0.4230 0.01893 0.00990 -0.0526 1.0000 0.1151
-5.500 -0.3935 0.01830 0.00929 -0.0532 1.0000 0.1278
-5.250 -0.3639 0.01772 0.00875 -0.0538 1.0000 0.1422
-5.000 -0.3342 0.01723 0.00831 -0.0544 1.0000 0.1582
-4.750 -0.3045 0.01682 0.00790 -0.0549 1.0000 0.1758
-4.500 -0.2748 0.01643 0.00754 -0.0555 1.0000 0.1934
-4.250 -0.2453 0.01605 0.00726 -0.0560 1.0000 0.2109
-4.000 -0.2159 0.01574 0.00704 -0.0566 1.0000 0.2296
-3.750 -0.1849 0.01549 0.00687 -0.0575 0.9887 0.2502
-3.500 -0.1456 0.01532 0.00671 -0.0598 0.9531 0.2749
-3.250 -0.1101 0.01518 0.00660 -0.0612 0.9224 0.2966
-3.000 -0.0795 0.01510 0.00651 -0.0613 0.8931 0.3168
-2.750 -0.0518 0.01505 0.00643 -0.0608 0.8655 0.3356
-2.500 -0.0251 0.01501 0.00634 -0.0600 0.8405 0.3544
-2.250 0.0018 0.01497 0.00625 -0.0593 0.8168 0.3731
-2.000 0.0287 0.01493 0.00617 -0.0586 0.7947 0.3916
-1.750 0.0558 0.01490 0.00608 -0.0580 0.7739 0.4105
-1.500 0.0834 0.01488 0.00602 -0.0575 0.7545 0.4297
-1.250 0.1111 0.01484 0.00597 -0.0571 0.7359 0.4494
-1.000 0.1388 0.01482 0.00593 -0.0566 0.7172 0.4712
-0.750 0.1667 0.01480 0.00590 -0.0562 0.6983 0.4942
-0.500 0.1948 0.01479 0.00586 -0.0558 0.6801 0.5167
-0.250 0.2227 0.01477 0.00584 -0.0555 0.6633 0.5375
0.000 0.2509 0.01476 0.00583 -0.0551 0.6477 0.5603
0.250 0.2791 0.01475 0.00584 -0.0549 0.6323 0.5869
0.500 0.3070 0.01472 0.00588 -0.0545 0.6176 0.6178
0.750 0.3346 0.01469 0.00594 -0.0540 0.6039 0.6544
1.000 0.3612 0.01463 0.00598 -0.0533 0.5910 0.6969
1.250 0.3861 0.01454 0.00601 -0.0521 0.5783 0.7500
1.750 0.4345 0.01407 0.00583 -0.0492 0.5516 1.0000
2.000 0.4646 0.01428 0.00591 -0.0495 0.5384 1.0000
2.250 0.4944 0.01451 0.00601 -0.0497 0.5251 1.0000
2.500 0.5241 0.01472 0.00614 -0.0500 0.5107 1.0000
2.750 0.5537 0.01495 0.00629 -0.0502 0.4961 1.0000
3.000 0.5831 0.01518 0.00646 -0.0504 0.4816 1.0000
3.250 0.6123 0.01542 0.00663 -0.0505 0.4675 1.0000
3.500 0.6414 0.01567 0.00682 -0.0506 0.4539 1.0000
3.750 0.6703 0.01594 0.00702 -0.0507 0.4405 1.0000
4.000 0.6993 0.01621 0.00727 -0.0508 0.4263 1.0000
4.250 0.7282 0.01649 0.00754 -0.0509 0.4121 1.0000
4.500 0.7569 0.01678 0.00783 -0.0510 0.3980 1.0000
4.750 0.7854 0.01709 0.00812 -0.0510 0.3837 1.0000
5.000 0.8137 0.01741 0.00842 -0.0510 0.3695 1.0000
5.250 0.8418 0.01775 0.00876 -0.0510 0.3545 1.0000
5.500 0.8698 0.01811 0.00910 -0.0510 0.3391 1.0000
5.750 0.8975 0.01849 0.00949 -0.0510 0.3226 1.0000
6.000 0.9251 0.01889 0.00989 -0.0510 0.3049 1.0000
6.250 0.9523 0.01931 0.01031 -0.0510 0.2862 1.0000
6.500 0.9790 0.01980 0.01077 -0.0509 0.2670 1.0000
6.750 1.0054 0.02033 0.01128 -0.0508 0.2476 1.0000
7.000 1.0316 0.02090 0.01183 -0.0508 0.2276 1.0000
7.250 1.0574 0.02151 0.01244 -0.0507 0.2078 1.0000
7.500 1.0825 0.02221 0.01310 -0.0506 0.1889 1.0000
7.750 1.1068 0.02301 0.01385 -0.0504 0.1712 1.0000
8.000 1.1305 0.02389 0.01468 -0.0502 0.1545 1.0000
8.250 1.1538 0.02479 0.01561 -0.0499 0.1386 1.0000
8.500 1.1763 0.02577 0.01660 -0.0495 0.1238 1.0000
8.750 1.1977 0.02684 0.01767 -0.0491 0.1109 1.0000
9.000 1.2179 0.02802 0.01884 -0.0486 0.1001 1.0000
9.250 1.2375 0.02921 0.02008 -0.0480 0.0908 1.0000
9.500 1.2559 0.03048 0.02142 -0.0472 0.0835 1.0000
9.750 1.2716 0.03193 0.02283 -0.0464 0.0779 1.0000
10.000 1.2881 0.03326 0.02433 -0.0454 0.0727 1.0000
10.250 1.3012 0.03480 0.02589 -0.0443 0.0689 1.0000
10.500 1.3135 0.03638 0.02758 -0.0431 0.0655 1.0000
10.750 1.3243 0.03799 0.02930 -0.0418 0.0624 1.0000
11.000 1.3316 0.03975 0.03111 -0.0404 0.0601 1.0000
11.250 1.3350 0.04167 0.03309 -0.0386 0.0584 1.0000
11.500 1.3389 0.04369 0.03530 -0.0372 0.0565 1.0000
11.750 1.3421 0.04595 0.03770 -0.0362 0.0547 1.0000
12.000 1.3444 0.04841 0.04026 -0.0357 0.0532 1.0000
12.250 1.3462 0.05103 0.04294 -0.0355 0.0519 1.0000
12.500 1.3480 0.05381 0.04576 -0.0353 0.0507 1.0000
12.750 1.3485 0.05695 0.04915 -0.0355 0.0495 1.0000
13.000 1.3481 0.06027 0.05266 -0.0358 0.0484 1.0000
13.250 1.3465 0.06379 0.05635 -0.0364 0.0475 1.0000
13.500 1.3440 0.06750 0.06021 -0.0372 0.0466 1.0000
13.750 1.3409 0.07131 0.06414 -0.0382 0.0457 1.0000
14.000 1.3382 0.07509 0.06801 -0.0392 0.0450 1.0000
14.250 1.3368 0.07866 0.07163 -0.0400 0.0442 1.0000
14.500 1.3306 0.08313 0.07623 -0.0414 0.0436 1.0000
14.750 1.3161 0.08919 0.08256 -0.0442 0.0431 1.0000
15.000 1.2990 0.09586 0.08949 -0.0475 0.0427 1.0000
15.250 1.2795 0.10326 0.09714 -0.0515 0.0424 1.0000
15.500 1.2566 0.11183 0.10595 -0.0565 0.0422 1.0000
15.750 1.2269 0.12249 0.11688 -0.0632 0.0421 1.0000
16.000 1.1775 0.13905 0.13375 -0.0742 0.0424 1.0000
|
Polar data table (+)
Polar graphs
<< Back to OAF117 AIRFOIL (oaf117-il)