Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

HQ 2.0/9 AIRFOIL (hq209-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: HQ 2.0/9 AIRFOIL (hq209-il)
Reynolds number: 50,000
Max Cl/Cd: 37.86 at α=5.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hq209-il-50000-n5.txt
Download as CSV file: xf-hq209-il-50000-n5.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 =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.4649   0.11367   0.10654  -0.0215   1.0000   0.0567
  -9.750  -0.4587   0.10974   0.10265  -0.0221   1.0000   0.0548
  -9.250  -0.4692   0.09952   0.09262  -0.0312   1.0000   0.0452
  -9.000  -0.4640   0.09567   0.08881  -0.0314   1.0000   0.0447
  -8.750  -0.4642   0.09140   0.08462  -0.0334   1.0000   0.0446
  -8.500  -0.4637   0.08736   0.08065  -0.0348   1.0000   0.0442
  -8.250  -0.4647   0.08338   0.07674  -0.0364   1.0000   0.0436
  -8.000  -0.4689   0.07897   0.07241  -0.0390   1.0000   0.0431
  -7.750  -0.4718   0.07460   0.06808  -0.0415   1.0000   0.0425
  -7.500  -0.4749   0.07023   0.06372  -0.0434   1.0000   0.0420
  -7.250  -0.4766   0.06614   0.05958  -0.0447   1.0000   0.0414
  -7.000  -0.4779   0.06191   0.05527  -0.0457   1.0000   0.0408
  -6.750  -0.4770   0.05785   0.05106  -0.0461   1.0000   0.0403
  -6.500  -0.4738   0.05380   0.04680  -0.0462   1.0000   0.0399
  -6.250  -0.4678   0.04991   0.04261  -0.0460   1.0000   0.0395
  -6.000  -0.4591   0.04615   0.03851  -0.0455   1.0000   0.0395
  -5.750  -0.4475   0.04267   0.03463  -0.0448   1.0000   0.0396
  -5.500  -0.4332   0.03933   0.03083  -0.0439   1.0000   0.0401
  -5.250  -0.4165   0.03631   0.02724  -0.0429   1.0000   0.0409
  -5.000  -0.3977   0.03372   0.02412  -0.0418   1.0000   0.0430
  -4.750  -0.3780   0.03144   0.02139  -0.0407   1.0000   0.0467
  -4.500  -0.3586   0.02969   0.01948  -0.0397   1.0000   0.0511
  -4.250  -0.3369   0.02784   0.01731  -0.0383   1.0000   0.0545
  -4.000  -0.3155   0.02623   0.01545  -0.0368   1.0000   0.0601
  -3.750  -0.2950   0.02514   0.01416  -0.0355   1.0000   0.0717
  -3.500  -0.2748   0.02385   0.01285  -0.0341   1.0000   0.0838
  -3.250  -0.2540   0.02276   0.01178  -0.0333   1.0000   0.1053
  -3.000  -0.2322   0.02167   0.01067  -0.0327   1.0000   0.1365
  -2.750  -0.2089   0.02029   0.00968  -0.0327   1.0000   0.2067
  -2.500  -0.1934   0.01844   0.00962  -0.0309   1.0000   0.5437
  -2.250  -0.1823   0.01832   0.00977  -0.0266   1.0000   0.6879
  -2.000  -0.1707   0.01826   0.00978  -0.0225   0.9992   0.7731
  -1.750  -0.1475   0.01810   0.00974  -0.0197   0.9916   0.8785
  -1.500  -0.0837   0.01799   0.00929  -0.0265   0.9871   1.0000
  -1.250  -0.0471   0.01815   0.00908  -0.0295   0.9785   1.0000
  -1.000  -0.0090   0.01837   0.00898  -0.0327   0.9707   1.0000
  -0.750   0.0271   0.01859   0.00888  -0.0353   0.9619   1.0000
  -0.500   0.0629   0.01882   0.00888  -0.0377   0.9532   1.0000
  -0.250   0.1021   0.01910   0.00895  -0.0407   0.9454   1.0000
   0.000   0.1354   0.01933   0.00902  -0.0425   0.9355   1.0000
   0.250   0.1710   0.01958   0.00911  -0.0446   0.9264   1.0000
   0.500   0.2106   0.01983   0.00925  -0.0473   0.9182   1.0000
   0.750   0.2435   0.02007   0.00942  -0.0488   0.9076   1.0000
   1.000   0.2785   0.02030   0.00960  -0.0506   0.8975   1.0000
   1.250   0.3184   0.02051   0.00978  -0.0531   0.8889   1.0000
   1.500   0.3520   0.02071   0.00998  -0.0544   0.8780   1.0000
   1.750   0.3837   0.02092   0.01023  -0.0552   0.8662   1.0000
   2.000   0.4178   0.02103   0.01036  -0.0562   0.8531   1.0000
   2.250   0.4542   0.02096   0.01035  -0.0572   0.8374   1.0000
   2.500   0.4901   0.02080   0.01025  -0.0578   0.8209   1.0000
   2.750   0.5171   0.02081   0.01038  -0.0571   0.8017   1.0000
   3.000   0.5465   0.02081   0.01047  -0.0568   0.7852   1.0000
   3.250   0.5763   0.02078   0.01055  -0.0565   0.7687   1.0000
   3.500   0.6052   0.02074   0.01063  -0.0560   0.7511   1.0000
   3.750   0.6308   0.02077   0.01086  -0.0550   0.7304   1.0000
   4.000   0.6589   0.02070   0.01092  -0.0541   0.7095   1.0000
   4.250   0.6852   0.02068   0.01104  -0.0530   0.6850   1.0000
   4.500   0.7111   0.02067   0.01116  -0.0518   0.6574   1.0000
   4.750   0.7360   0.02071   0.01131  -0.0504   0.6248   1.0000
   5.000   0.7611   0.02076   0.01150  -0.0489   0.5872   1.0000
   5.250   0.7844   0.02096   0.01171  -0.0472   0.5423   1.0000
   5.500   0.8064   0.02130   0.01195  -0.0454   0.4911   1.0000
   5.750   0.8260   0.02190   0.01234  -0.0435   0.4357   1.0000
   6.000   0.8436   0.02274   0.01294  -0.0416   0.3805   1.0000
   6.250   0.8601   0.02371   0.01380  -0.0399   0.3282   1.0000
   6.500   0.8758   0.02483   0.01470  -0.0383   0.2807   1.0000
   6.750   0.8915   0.02604   0.01580  -0.0368   0.2357   1.0000
   7.000   0.9073   0.02734   0.01703  -0.0354   0.1929   1.0000
   7.250   0.9229   0.02883   0.01841  -0.0339   0.1580   1.0000
   7.500   0.9387   0.03040   0.01986  -0.0326   0.1301   1.0000
   7.750   0.9560   0.03202   0.02148  -0.0314   0.1112   1.0000
   8.000   0.9758   0.03379   0.02336  -0.0303   0.0997   1.0000
   8.250   0.9949   0.03560   0.02528  -0.0292   0.0883   1.0000
   8.500   1.0129   0.03752   0.02741  -0.0282   0.0780   1.0000
   8.750   1.0346   0.03995   0.03005  -0.0273   0.0703   1.0000
   9.000   1.0511   0.04262   0.03307  -0.0260   0.0611   1.0000
   9.250   1.0594   0.04493   0.03558  -0.0245   0.0516   1.0000
   9.500   1.0649   0.04696   0.03760  -0.0230   0.0441   1.0000
   9.750   1.0693   0.04990   0.04110  -0.0211   0.0378   1.0000
  10.000   1.0731   0.05297   0.04443  -0.0194   0.0348   1.0000
  10.250   1.0737   0.05615   0.04776  -0.0176   0.0330   1.0000
  10.500   1.0693   0.06002   0.05188  -0.0159   0.0320   1.0000
  10.750   1.0592   0.06395   0.05619  -0.0143   0.0315   1.0000
  11.000   1.0447   0.06831   0.06090  -0.0136   0.0312   1.0000
  11.250   1.0291   0.07299   0.06584  -0.0139   0.0311   1.0000
  11.500   1.0113   0.07830   0.07139  -0.0154   0.0311   1.0000
  11.750   0.9933   0.08414   0.07742  -0.0179   0.0312   1.0000
  12.000   0.9748   0.09074   0.08418  -0.0215   0.0314   1.0000
  12.250   0.9574   0.09794   0.09150  -0.0260   0.0317   1.0000
  12.500   0.9407   0.10588   0.09952  -0.0310   0.0320   1.0000
  12.750   0.9260   0.11416   0.10785  -0.0362   0.0323   1.0000
<< Back to HQ 2.0/9 AIRFOIL (hq209-il)

Polar data table (+)

Polar graphs


<< Back to HQ 2.0/9 AIRFOIL (hq209-il)