OAF117 AIRFOIL (oaf117-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: OAF117 AIRFOIL (oaf117-il) Reynolds number: 500,000 Max Cl/Cd: 84.41 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-oaf117-il-500000.txt Download as CSV file: xf-oaf117-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-14.750 -1.0533 0.08980 0.08684 0.0067 1.0000 0.0231
-14.500 -1.1002 0.07723 0.07405 -0.0018 1.0000 0.0229
-14.250 -1.1306 0.06794 0.06459 -0.0084 1.0000 0.0229
-14.000 -1.1522 0.06018 0.05666 -0.0143 1.0000 0.0230
-13.750 -1.1677 0.05334 0.04966 -0.0199 1.0000 0.0231
-13.500 -1.1781 0.04715 0.04330 -0.0257 1.0000 0.0232
-13.250 -1.1858 0.04142 0.03738 -0.0322 1.0000 0.0234
-13.000 -1.1797 0.03746 0.03319 -0.0372 1.0000 0.0238
-12.750 -1.1677 0.03493 0.03046 -0.0390 1.0000 0.0244
-12.500 -1.1518 0.03289 0.02819 -0.0403 1.0000 0.0249
-12.250 -1.1359 0.03061 0.02571 -0.0412 1.0000 0.0254
-12.000 -1.1164 0.02895 0.02400 -0.0417 1.0000 0.0261
-11.750 -1.0933 0.02792 0.02293 -0.0422 1.0000 0.0267
-11.500 -1.0694 0.02693 0.02186 -0.0427 1.0000 0.0276
-11.250 -1.0449 0.02591 0.02069 -0.0431 1.0000 0.0285
-11.000 -1.0196 0.02497 0.01958 -0.0435 1.0000 0.0292
-10.750 -0.9962 0.02333 0.01785 -0.0440 1.0000 0.0302
-10.500 -0.9703 0.02246 0.01696 -0.0444 1.0000 0.0310
-10.250 -0.9434 0.02175 0.01621 -0.0448 1.0000 0.0321
-10.000 -0.9161 0.02106 0.01541 -0.0451 1.0000 0.0333
-9.750 -0.8882 0.02047 0.01469 -0.0454 1.0000 0.0342
-9.500 -0.8620 0.01914 0.01334 -0.0460 1.0000 0.0355
-9.250 -0.8340 0.01850 0.01268 -0.0464 1.0000 0.0367
-9.000 -0.8056 0.01793 0.01206 -0.0467 1.0000 0.0381
-8.750 -0.7768 0.01746 0.01149 -0.0470 1.0000 0.0395
-8.500 -0.7487 0.01646 0.01047 -0.0476 1.0000 0.0412
-8.250 -0.7197 0.01590 0.00990 -0.0480 1.0000 0.0428
-8.000 -0.6905 0.01543 0.00938 -0.0483 1.0000 0.0446
-7.750 -0.6611 0.01500 0.00888 -0.0487 1.0000 0.0463
-7.500 -0.6318 0.01421 0.00812 -0.0493 1.0000 0.0488
-7.250 -0.6021 0.01377 0.00766 -0.0497 1.0000 0.0511
-7.000 -0.5723 0.01344 0.00727 -0.0500 1.0000 0.0532
-6.750 -0.5423 0.01275 0.00661 -0.0507 1.0000 0.0569
-6.500 -0.5122 0.01239 0.00623 -0.0511 1.0000 0.0601
-6.250 -0.4819 0.01189 0.00573 -0.0517 1.0000 0.0643
-6.000 -0.4516 0.01152 0.00537 -0.0522 1.0000 0.0692
-5.750 -0.4211 0.01108 0.00497 -0.0527 1.0000 0.0759
-5.500 -0.3905 0.01069 0.00463 -0.0533 1.0000 0.0850
-5.250 -0.3598 0.01031 0.00434 -0.0539 1.0000 0.0991
-5.000 -0.3291 0.01000 0.00412 -0.0545 1.0000 0.1168
-4.750 -0.2954 0.00982 0.00398 -0.0557 0.9664 0.1337
-4.500 -0.2680 0.00977 0.00392 -0.0551 0.9258 0.1482
-4.000 -0.2160 0.00967 0.00373 -0.0535 0.8681 0.1768
-3.750 -0.1885 0.00958 0.00360 -0.0531 0.8419 0.1912
-3.500 -0.1601 0.00949 0.00346 -0.0530 0.8166 0.2063
-3.250 -0.1313 0.00941 0.00334 -0.0529 0.7927 0.2231
-3.000 -0.1022 0.00934 0.00323 -0.0529 0.7715 0.2406
-2.750 -0.0729 0.00927 0.00313 -0.0530 0.7504 0.2595
-2.500 -0.0433 0.00920 0.00305 -0.0531 0.7285 0.2794
-2.250 -0.0138 0.00916 0.00298 -0.0532 0.7076 0.3003
-2.000 0.0159 0.00912 0.00294 -0.0533 0.6894 0.3234
-1.750 0.0456 0.00910 0.00290 -0.0535 0.6723 0.3440
-1.500 0.0754 0.00908 0.00286 -0.0536 0.6560 0.3614
-1.250 0.1052 0.00907 0.00283 -0.0538 0.6405 0.3790
-1.000 0.1349 0.00907 0.00281 -0.0540 0.6256 0.3956
-0.750 0.1647 0.00908 0.00279 -0.0541 0.6107 0.4120
0.000 0.2542 0.00909 0.00280 -0.0547 0.5689 0.4646
0.250 0.2840 0.00910 0.00282 -0.0548 0.5556 0.4836
0.500 0.3137 0.00911 0.00285 -0.0550 0.5423 0.5051
0.750 0.3435 0.00911 0.00288 -0.0552 0.5283 0.5277
1.250 0.4029 0.00912 0.00296 -0.0556 0.5005 0.5777
1.500 0.4325 0.00912 0.00302 -0.0557 0.4878 0.6105
1.750 0.4620 0.00911 0.00309 -0.0559 0.4751 0.6510
2.000 0.4913 0.00904 0.00318 -0.0559 0.4626 0.7006
2.250 0.5197 0.00894 0.00328 -0.0558 0.4507 0.7676
2.500 0.5431 0.00870 0.00331 -0.0543 0.4392 0.8767
2.750 0.5681 0.00850 0.00318 -0.0532 0.4273 1.0000
3.000 0.5980 0.00864 0.00327 -0.0535 0.4148 1.0000
3.250 0.6278 0.00880 0.00337 -0.0537 0.4026 1.0000
3.500 0.6575 0.00898 0.00347 -0.0539 0.3900 1.0000
3.750 0.6872 0.00914 0.00359 -0.0542 0.3769 1.0000
4.000 0.7168 0.00930 0.00372 -0.0544 0.3638 1.0000
4.250 0.7463 0.00949 0.00385 -0.0546 0.3501 1.0000
4.500 0.7756 0.00969 0.00400 -0.0547 0.3352 1.0000
4.750 0.8049 0.00991 0.00415 -0.0549 0.3188 1.0000
5.000 0.8341 0.01014 0.00432 -0.0551 0.3020 1.0000
5.250 0.8631 0.01039 0.00450 -0.0553 0.2853 1.0000
5.500 0.8920 0.01066 0.00471 -0.0555 0.2688 1.0000
5.750 0.9209 0.01095 0.00494 -0.0556 0.2523 1.0000
6.000 0.9496 0.01125 0.00518 -0.0558 0.2339 1.0000
6.250 0.9780 0.01162 0.00544 -0.0559 0.2138 1.0000
6.500 1.0063 0.01199 0.00574 -0.0561 0.1948 1.0000
6.750 1.0344 0.01239 0.00606 -0.0562 0.1772 1.0000
7.000 1.0623 0.01282 0.00641 -0.0563 0.1613 1.0000
7.250 1.0900 0.01325 0.00679 -0.0564 0.1457 1.0000
7.500 1.1175 0.01373 0.00719 -0.0565 0.1297 1.0000
7.750 1.1446 0.01426 0.00763 -0.0565 0.1133 1.0000
8.000 1.1715 0.01480 0.00810 -0.0565 0.0968 1.0000
8.250 1.1977 0.01545 0.00865 -0.0565 0.0813 1.0000
8.500 1.2236 0.01614 0.00926 -0.0564 0.0697 1.0000
8.750 1.2493 0.01679 0.00989 -0.0563 0.0619 1.0000
9.000 1.2748 0.01745 0.01056 -0.0561 0.0561 1.0000
9.250 1.2990 0.01828 0.01136 -0.0558 0.0511 1.0000
9.500 1.3244 0.01883 0.01197 -0.0556 0.0480 1.0000
9.750 1.3480 0.01962 0.01275 -0.0552 0.0451 1.0000
10.000 1.3711 0.02045 0.01363 -0.0548 0.0428 1.0000
10.250 1.3946 0.02112 0.01437 -0.0543 0.0408 1.0000
10.500 1.4165 0.02198 0.01524 -0.0538 0.0389 1.0000
10.750 1.4356 0.02311 0.01642 -0.0530 0.0372 1.0000
11.000 1.4574 0.02383 0.01722 -0.0524 0.0359 1.0000
11.250 1.4773 0.02469 0.01815 -0.0516 0.0346 1.0000
11.500 1.4953 0.02569 0.01919 -0.0507 0.0335 1.0000
11.750 1.5072 0.02722 0.02075 -0.0492 0.0325 1.0000
12.000 1.5201 0.02847 0.02210 -0.0478 0.0317 1.0000
12.250 1.5331 0.02958 0.02332 -0.0463 0.0311 1.0000
12.500 1.5423 0.03086 0.02470 -0.0446 0.0304 1.0000
12.750 1.5453 0.03239 0.02633 -0.0424 0.0298 1.0000
13.000 1.5485 0.03428 0.02830 -0.0411 0.0293 1.0000
13.250 1.5517 0.03649 0.03059 -0.0405 0.0287 1.0000
13.500 1.5529 0.03914 0.03332 -0.0404 0.0282 1.0000
13.750 1.5498 0.04246 0.03671 -0.0405 0.0278 1.0000
14.000 1.5452 0.04606 0.04040 -0.0407 0.0273 1.0000
14.250 1.5471 0.04908 0.04355 -0.0413 0.0271 1.0000
14.500 1.5471 0.05238 0.04698 -0.0421 0.0268 1.0000
14.750 1.5457 0.05591 0.05064 -0.0429 0.0264 1.0000
15.000 1.5426 0.05967 0.05452 -0.0438 0.0261 1.0000
15.250 1.5386 0.06357 0.05854 -0.0448 0.0258 1.0000
15.500 1.5336 0.06763 0.06271 -0.0459 0.0256 1.0000
15.750 1.5278 0.07183 0.06701 -0.0471 0.0253 1.0000
16.000 1.5216 0.07613 0.07142 -0.0484 0.0250 1.0000
16.250 1.5150 0.08053 0.07591 -0.0498 0.0248 1.0000
16.500 1.5088 0.08502 0.08049 -0.0513 0.0246 1.0000
|
Polar data table (+)
Polar graphs
<< Back to OAF117 AIRFOIL (oaf117-il)