OAF117 AIRFOIL (oaf117-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: OAF117 AIRFOIL (oaf117-il) Reynolds number: 200,000 Max Cl/Cd: 64.17 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-oaf117-il-200000.txt Download as CSV file: xf-oaf117-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.6110 0.13424 0.13055 0.0305 1.0000 0.0766
-11.000 -0.8846 0.05990 0.05590 -0.0265 1.0000 0.0471
-10.750 -0.9196 0.04892 0.04414 -0.0390 1.0000 0.0465
-10.500 -0.9074 0.04570 0.04084 -0.0402 1.0000 0.0474
-10.250 -0.8898 0.04352 0.03861 -0.0409 1.0000 0.0483
-10.000 -0.8755 0.04012 0.03490 -0.0426 1.0000 0.0493
-9.750 -0.8583 0.03690 0.03129 -0.0441 1.0000 0.0508
-9.500 -0.8383 0.03385 0.02769 -0.0456 1.0000 0.0525
-9.250 -0.8162 0.03082 0.02426 -0.0468 1.0000 0.0540
-9.000 -0.7912 0.02906 0.02248 -0.0472 1.0000 0.0556
-8.750 -0.7650 0.02772 0.02102 -0.0477 1.0000 0.0577
-8.500 -0.7379 0.02628 0.01930 -0.0482 1.0000 0.0601
-8.250 -0.7107 0.02451 0.01724 -0.0488 1.0000 0.0622
-8.000 -0.6835 0.02313 0.01590 -0.0492 1.0000 0.0646
-7.750 -0.6553 0.02217 0.01487 -0.0496 1.0000 0.0676
-7.500 -0.6262 0.02129 0.01374 -0.0500 1.0000 0.0706
-7.250 -0.5983 0.01983 0.01236 -0.0504 1.0000 0.0739
-7.000 -0.5693 0.01907 0.01157 -0.0508 1.0000 0.0779
-6.750 -0.5400 0.01814 0.01053 -0.0512 1.0000 0.0823
-6.500 -0.5109 0.01733 0.00981 -0.0516 1.0000 0.0872
-6.250 -0.4812 0.01664 0.00904 -0.0520 1.0000 0.0934
-6.000 -0.4515 0.01596 0.00844 -0.0526 1.0000 0.1008
-5.750 -0.4216 0.01528 0.00780 -0.0531 1.0000 0.1099
-5.500 -0.3914 0.01467 0.00724 -0.0537 1.0000 0.1222
-5.250 -0.3611 0.01407 0.00671 -0.0543 1.0000 0.1386
-5.000 -0.3307 0.01360 0.00632 -0.0549 1.0000 0.1585
-4.750 -0.3002 0.01323 0.00601 -0.0555 1.0000 0.1790
-4.500 -0.2698 0.01286 0.00573 -0.0562 1.0000 0.1981
-4.250 -0.2393 0.01255 0.00550 -0.0569 1.0000 0.2165
-4.000 -0.2087 0.01228 0.00532 -0.0576 1.0000 0.2354
-3.750 -0.1782 0.01205 0.00516 -0.0584 1.0000 0.2550
-3.500 -0.1429 0.01188 0.00507 -0.0602 0.9848 0.2765
-3.250 -0.1022 0.01177 0.00504 -0.0627 0.9538 0.3006
-3.000 -0.0719 0.01181 0.00506 -0.0627 0.9208 0.3237
-2.500 -0.0233 0.01187 0.00511 -0.0599 0.8633 0.3647
-2.250 0.0018 0.01189 0.00508 -0.0587 0.8389 0.3845
-2.000 0.0286 0.01187 0.00503 -0.0580 0.8146 0.4046
-1.750 0.0557 0.01186 0.00496 -0.0574 0.7920 0.4243
-1.500 0.0832 0.01187 0.00490 -0.0568 0.7711 0.4443
-1.250 0.1114 0.01187 0.00484 -0.0565 0.7511 0.4647
-1.000 0.1398 0.01184 0.00481 -0.0562 0.7319 0.4842
-0.750 0.1684 0.01182 0.00478 -0.0560 0.7137 0.5043
-0.500 0.1972 0.01181 0.00477 -0.0559 0.6968 0.5262
-0.250 0.2260 0.01182 0.00476 -0.0557 0.6808 0.5514
0.000 0.2547 0.01179 0.00478 -0.0555 0.6651 0.5763
0.250 0.2836 0.01177 0.00479 -0.0555 0.6491 0.6045
0.500 0.3123 0.01172 0.00483 -0.0553 0.6336 0.6360
0.750 0.3405 0.01167 0.00488 -0.0550 0.6189 0.6729
1.000 0.3678 0.01158 0.00493 -0.0545 0.6048 0.7183
1.250 0.3926 0.01142 0.00495 -0.0533 0.5912 0.7848
1.750 0.4414 0.01100 0.00466 -0.0506 0.5632 1.0000
2.000 0.4718 0.01117 0.00472 -0.0509 0.5481 1.0000
2.250 0.5019 0.01135 0.00480 -0.0512 0.5336 1.0000
2.500 0.5318 0.01154 0.00489 -0.0515 0.5197 1.0000
2.750 0.5615 0.01175 0.00499 -0.0517 0.5063 1.0000
3.000 0.5911 0.01195 0.00511 -0.0518 0.4929 1.0000
3.250 0.6208 0.01213 0.00525 -0.0520 0.4786 1.0000
3.500 0.6503 0.01233 0.00541 -0.0522 0.4649 1.0000
3.750 0.6796 0.01255 0.00558 -0.0523 0.4513 1.0000
4.000 0.7088 0.01278 0.00574 -0.0524 0.4380 1.0000
4.250 0.7378 0.01303 0.00591 -0.0525 0.4244 1.0000
4.500 0.7670 0.01323 0.00612 -0.0527 0.4099 1.0000
4.750 0.7960 0.01346 0.00633 -0.0528 0.3951 1.0000
5.000 0.8248 0.01370 0.00655 -0.0528 0.3799 1.0000
5.250 0.8535 0.01394 0.00677 -0.0529 0.3638 1.0000
5.500 0.8819 0.01420 0.00699 -0.0530 0.3466 1.0000
5.750 0.9102 0.01448 0.00723 -0.0530 0.3286 1.0000
6.000 0.9382 0.01480 0.00749 -0.0530 0.3101 1.0000
6.250 0.9662 0.01513 0.00779 -0.0531 0.2909 1.0000
6.500 0.9940 0.01549 0.00814 -0.0531 0.2713 1.0000
6.750 1.0214 0.01592 0.00853 -0.0531 0.2516 1.0000
7.000 1.0486 0.01638 0.00894 -0.0532 0.2317 1.0000
7.250 1.0755 0.01688 0.00939 -0.0532 0.2116 1.0000
7.500 1.1019 0.01745 0.00992 -0.0531 0.1913 1.0000
7.750 1.1275 0.01815 0.01053 -0.0531 0.1713 1.0000
8.000 1.1527 0.01892 0.01124 -0.0529 0.1512 1.0000
8.250 1.1771 0.01977 0.01203 -0.0528 0.1310 1.0000
8.500 1.1999 0.02086 0.01298 -0.0525 0.1136 1.0000
8.750 1.2225 0.02193 0.01400 -0.0520 0.0998 1.0000
9.000 1.2441 0.02307 0.01509 -0.0515 0.0894 1.0000
9.250 1.2663 0.02405 0.01614 -0.0509 0.0818 1.0000
9.500 1.2857 0.02538 0.01744 -0.0501 0.0760 1.0000
9.750 1.3062 0.02643 0.01855 -0.0494 0.0711 1.0000
10.000 1.3231 0.02798 0.02005 -0.0484 0.0673 1.0000
10.250 1.3423 0.02911 0.02131 -0.0475 0.0640 1.0000
10.500 1.3598 0.03036 0.02257 -0.0465 0.0612 1.0000
10.750 1.3747 0.03217 0.02433 -0.0452 0.0587 1.0000
11.000 1.3911 0.03341 0.02576 -0.0441 0.0565 1.0000
11.250 1.4062 0.03480 0.02724 -0.0428 0.0545 1.0000
11.500 1.4205 0.03626 0.02873 -0.0416 0.0529 1.0000
11.750 1.4352 0.03821 0.03064 -0.0404 0.0513 1.0000
12.000 1.4460 0.04011 0.03272 -0.0389 0.0501 1.0000
12.250 1.4531 0.04184 0.03466 -0.0371 0.0490 1.0000
12.500 1.4568 0.04366 0.03664 -0.0350 0.0479 1.0000
12.750 1.4602 0.04566 0.03878 -0.0334 0.0469 1.0000
13.000 1.4646 0.04775 0.04097 -0.0323 0.0459 1.0000
13.250 1.4712 0.04992 0.04321 -0.0313 0.0451 1.0000
13.500 1.4814 0.05245 0.04575 -0.0304 0.0442 1.0000
13.750 1.4806 0.05586 0.04935 -0.0296 0.0437 1.0000
14.000 1.4716 0.05937 0.05313 -0.0294 0.0434 1.0000
14.250 1.4608 0.06340 0.05743 -0.0298 0.0431 1.0000
14.500 1.4479 0.06797 0.06226 -0.0309 0.0428 1.0000
14.750 1.4324 0.07314 0.06769 -0.0326 0.0426 1.0000
15.000 1.4141 0.07897 0.07377 -0.0350 0.0425 1.0000
15.250 1.3928 0.08554 0.08059 -0.0381 0.0424 1.0000
15.500 1.3681 0.09298 0.08828 -0.0421 0.0424 1.0000
15.750 1.3394 0.10152 0.09707 -0.0471 0.0424 1.0000
16.000 1.3068 0.11145 0.10724 -0.0533 0.0427 1.0000
16.250 1.2690 0.12343 0.11945 -0.0613 0.0430 1.0000
16.500 1.2255 0.13789 0.13412 -0.0711 0.0435 1.0000
|
Polar data table (+)
Polar graphs
<< Back to OAF117 AIRFOIL (oaf117-il)