HQ 2.0/9 AIRFOIL (hq209-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: HQ 2.0/9 AIRFOIL (hq209-il) Reynolds number: 200,000 Max Cl/Cd: 74.24 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq209-il-200000.txt Download as CSV file: xf-hq209-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 2.0/9 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.4505 0.10425 0.10064 -0.0216 1.0000 0.0367
-9.500 -0.4486 0.10055 0.09697 -0.0235 1.0000 0.0378
-9.250 -0.4487 0.09660 0.09306 -0.0261 1.0000 0.0393
-9.000 -0.3774 0.08228 0.07907 -0.0317 1.0000 0.0436
-8.750 -0.3701 0.07937 0.07617 -0.0310 1.0000 0.0457
-8.500 -0.3707 0.07544 0.07227 -0.0321 1.0000 0.0469
-8.250 -0.3742 0.07118 0.06805 -0.0335 1.0000 0.0485
-7.500 -0.4931 0.06423 0.06076 -0.0490 1.0000 0.0422
-7.250 -0.4778 0.06321 0.05999 -0.0446 1.0000 0.0453
-7.000 -0.4793 0.05989 0.05663 -0.0447 1.0000 0.0463
-6.750 -0.4814 0.05705 0.05374 -0.0440 1.0000 0.0493
-6.500 -0.4852 0.05390 0.05042 -0.0430 1.0000 0.0517
-6.250 -0.4876 0.05367 0.04952 -0.0412 1.0000 0.0542
-6.000 -0.4891 0.04654 0.04254 -0.0407 1.0000 0.0559
-5.750 -0.4813 0.04394 0.03995 -0.0392 1.0000 0.0574
-5.500 -0.4717 0.04153 0.03745 -0.0380 1.0000 0.0597
-5.250 -0.4437 0.03782 0.03300 -0.0407 0.9969 0.0689
-5.000 -0.4058 0.02688 0.02079 -0.0408 0.9947 0.0308
-4.750 -0.3707 0.02445 0.01801 -0.0422 0.9910 0.0290
-4.500 -0.3363 0.02154 0.01463 -0.0436 0.9874 0.0289
-4.250 -0.2993 0.01958 0.01231 -0.0452 0.9842 0.0299
-4.000 -0.2644 0.01764 0.01028 -0.0471 0.9813 0.0356
-3.750 -0.2316 0.01636 0.00888 -0.0478 0.9762 0.0394
-3.500 -0.1961 0.01498 0.00749 -0.0494 0.9721 0.0487
-3.250 -0.1573 0.01397 0.00650 -0.0519 0.9690 0.0777
-3.000 -0.1265 0.01320 0.00591 -0.0529 0.9624 0.1222
-2.750 -0.0971 0.01109 0.00550 -0.0547 0.9581 0.4976
-2.500 -0.0618 0.01081 0.00562 -0.0558 0.9542 0.6441
-2.250 -0.0325 0.01078 0.00560 -0.0558 0.9468 0.6947
-2.000 0.0053 0.01074 0.00558 -0.0573 0.9430 0.7383
-1.750 0.0369 0.01073 0.00556 -0.0576 0.9369 0.7699
-1.500 0.0695 0.01066 0.00550 -0.0580 0.9310 0.7996
-1.250 0.1031 0.01055 0.00539 -0.0585 0.9265 0.8302
-1.000 0.1275 0.01044 0.00532 -0.0571 0.9180 0.8569
-0.750 0.1582 0.01024 0.00515 -0.0568 0.9131 0.8821
-0.500 0.1833 0.01012 0.00503 -0.0558 0.9041 0.9026
-0.250 0.2174 0.00997 0.00485 -0.0565 0.8982 0.9190
0.000 0.2521 0.00981 0.00468 -0.0575 0.8891 0.9372
0.250 0.2916 0.00965 0.00448 -0.0595 0.8795 0.9553
0.500 0.3357 0.00945 0.00423 -0.0624 0.8709 0.9712
0.750 0.3796 0.00929 0.00403 -0.0656 0.8600 0.9889
1.000 0.4098 0.00922 0.00392 -0.0664 0.8476 1.0000
1.250 0.4307 0.00925 0.00390 -0.0652 0.8357 1.0000
1.500 0.4551 0.00928 0.00388 -0.0645 0.8243 1.0000
1.750 0.4807 0.00930 0.00386 -0.0640 0.8129 1.0000
2.000 0.5067 0.00932 0.00384 -0.0634 0.8009 1.0000
2.250 0.5328 0.00933 0.00385 -0.0628 0.7884 1.0000
2.500 0.5587 0.00936 0.00385 -0.0622 0.7745 1.0000
2.750 0.5845 0.00939 0.00388 -0.0616 0.7592 1.0000
3.000 0.6103 0.00942 0.00391 -0.0609 0.7422 1.0000
3.250 0.6361 0.00946 0.00394 -0.0602 0.7238 1.0000
3.500 0.6617 0.00952 0.00401 -0.0595 0.7027 1.0000
3.750 0.6870 0.00960 0.00406 -0.0587 0.6780 1.0000
4.000 0.7116 0.00972 0.00413 -0.0577 0.6461 1.0000
4.250 0.7350 0.00990 0.00419 -0.0565 0.6024 1.0000
4.500 0.7569 0.01023 0.00432 -0.0551 0.5425 1.0000
4.750 0.7776 0.01075 0.00453 -0.0537 0.4733 1.0000
5.000 0.7980 0.01138 0.00487 -0.0524 0.4091 1.0000
5.250 0.8191 0.01202 0.00529 -0.0514 0.3570 1.0000
5.500 0.8405 0.01266 0.00574 -0.0504 0.3121 1.0000
5.750 0.8620 0.01332 0.00624 -0.0495 0.2734 1.0000
6.000 0.8835 0.01398 0.00680 -0.0487 0.2363 1.0000
6.250 0.9045 0.01470 0.00734 -0.0478 0.1828 1.0000
6.500 0.9212 0.01601 0.00819 -0.0464 0.1116 1.0000
6.750 0.9391 0.01723 0.00924 -0.0449 0.0876 1.0000
7.000 0.9581 0.01830 0.01028 -0.0437 0.0752 1.0000
7.250 0.9774 0.01930 0.01126 -0.0426 0.0652 1.0000
7.500 0.9988 0.02005 0.01218 -0.0416 0.0577 1.0000
7.750 1.0184 0.02098 0.01314 -0.0407 0.0493 1.0000
8.000 1.0365 0.02204 0.01421 -0.0396 0.0399 1.0000
8.250 1.0524 0.02355 0.01582 -0.0379 0.0284 1.0000
8.500 1.0698 0.02495 0.01735 -0.0363 0.0227 1.0000
8.750 1.0819 0.02819 0.02065 -0.0343 0.0200 1.0000
9.000 1.1008 0.03004 0.02275 -0.0329 0.0190 1.0000
9.250 1.1181 0.03252 0.02554 -0.0315 0.0182 1.0000
9.500 1.1327 0.03530 0.02865 -0.0298 0.0177 1.0000
9.750 1.1446 0.03736 0.03099 -0.0280 0.0164 1.0000
10.000 1.1539 0.03950 0.03343 -0.0261 0.0155 1.0000
10.250 1.1612 0.04101 0.03508 -0.0243 0.0145 1.0000
10.500 1.1623 0.04351 0.03781 -0.0218 0.0140 1.0000
10.750 1.1564 0.04650 0.04109 -0.0186 0.0139 1.0000
11.000 1.1462 0.04972 0.04461 -0.0157 0.0139 1.0000
11.250 1.1322 0.05335 0.04855 -0.0134 0.0140 1.0000
11.500 1.1145 0.05759 0.05309 -0.0120 0.0142 1.0000
11.750 1.0947 0.06239 0.05817 -0.0118 0.0143 1.0000
12.000 1.0754 0.06749 0.06349 -0.0129 0.0143 1.0000
12.250 1.0484 0.07461 0.07088 -0.0158 0.0147 1.0000
12.500 1.0158 0.08386 0.08040 -0.0215 0.0152 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 2.0/9 AIRFOIL (hq209-il)