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

HQ 2.5/9 AIRFOIL (hq259-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: HQ 2.5/9 AIRFOIL (hq259-il)
Reynolds number: 50,000
Max Cl/Cd: 39.6 at α=5.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hq259-il-50000-n5.txt
Download as CSV file: xf-hq259-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 2.5/9 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.4050   0.09841   0.09157  -0.0321   1.0000   0.0470
  -8.500  -0.4078   0.09432   0.08758  -0.0345   1.0000   0.0439
  -8.250  -0.4133   0.08999   0.08335  -0.0378   1.0000   0.0401
  -8.000  -0.4131   0.08664   0.08008  -0.0379   1.0000   0.0393
  -7.750  -0.4167   0.08324   0.07679  -0.0385   1.0000   0.0386
  -7.500  -0.4196   0.07955   0.07319  -0.0402   1.0000   0.0380
  -7.250  -0.4230   0.07551   0.06922  -0.0423   1.0000   0.0376
  -7.000  -0.4259   0.07172   0.06547  -0.0437   1.0000   0.0371
  -6.750  -0.4281   0.06799   0.06174  -0.0446   1.0000   0.0366
  -6.500  -0.4297   0.06415   0.05785  -0.0455   1.0000   0.0366
  -6.250  -0.4291   0.06036   0.05395  -0.0460   1.0000   0.0369
  -6.000  -0.4256   0.05667   0.05010  -0.0463   1.0000   0.0371
  -5.750  -0.4190   0.05295   0.04613  -0.0467   1.0000   0.0376
  -5.500  -0.4090   0.04915   0.04197  -0.0470   1.0000   0.0386
  -5.250  -0.3956   0.04554   0.03793  -0.0471   1.0000   0.0393
  -5.000  -0.3794   0.04209   0.03403  -0.0471   1.0000   0.0396
  -4.750  -0.3608   0.03887   0.03033  -0.0469   1.0000   0.0398
  -4.500  -0.3397   0.03587   0.02679  -0.0465   1.0000   0.0403
  -4.250  -0.3171   0.03325   0.02360  -0.0459   1.0000   0.0410
  -4.000  -0.2952   0.03068   0.02079  -0.0455   1.0000   0.0429
  -3.750  -0.2729   0.02894   0.01872  -0.0448   1.0000   0.0463
  -3.500  -0.2495   0.02766   0.01697  -0.0438   1.0000   0.0539
  -3.250  -0.2276   0.02614   0.01545  -0.0429   1.0000   0.0610
  -3.000  -0.2052   0.02488   0.01404  -0.0418   1.0000   0.0719
  -2.750  -0.1830   0.02386   0.01290  -0.0409   1.0000   0.0904
  -2.500  -0.1485   0.02271   0.01170  -0.0427   0.9952   0.1244
  -2.250  -0.1126   0.02129   0.01060  -0.0451   0.9902   0.1941
  -2.000  -0.0893   0.01947   0.01081  -0.0441   0.9857   0.6378
  -1.750  -0.0698   0.01938   0.01082  -0.0411   0.9778   0.7709
  -1.500  -0.0302   0.01885   0.01048  -0.0410   0.9718   1.0000
  -1.250   0.0055   0.01903   0.01022  -0.0437   0.9637   1.0000
  -1.000   0.0427   0.01927   0.01007  -0.0464   0.9564   1.0000
  -0.750   0.0793   0.01952   0.00994  -0.0490   0.9486   1.0000
  -0.500   0.1138   0.01978   0.00992  -0.0511   0.9402   1.0000
  -0.250   0.1529   0.02006   0.00995  -0.0539   0.9330   1.0000
   0.000   0.1850   0.02031   0.01001  -0.0553   0.9235   1.0000
   0.250   0.2203   0.02057   0.01007  -0.0573   0.9149   1.0000
   0.500   0.2582   0.02082   0.01019  -0.0596   0.9067   1.0000
   0.750   0.2900   0.02107   0.01034  -0.0608   0.8965   1.0000
   1.000   0.3247   0.02131   0.01050  -0.0625   0.8871   1.0000
   1.250   0.3642   0.02150   0.01064  -0.0649   0.8789   1.0000
   1.500   0.3945   0.02173   0.01085  -0.0656   0.8678   1.0000
   1.750   0.4260   0.02194   0.01105  -0.0664   0.8568   1.0000
   2.000   0.4590   0.02210   0.01124  -0.0673   0.8457   1.0000
   2.250   0.4928   0.02215   0.01133  -0.0682   0.8331   1.0000
   2.500   0.5255   0.02214   0.01137  -0.0686   0.8186   1.0000
   2.750   0.5574   0.02208   0.01139  -0.0688   0.8034   1.0000
   3.000   0.5886   0.02203   0.01147  -0.0687   0.7881   1.0000
   3.250   0.6195   0.02196   0.01150  -0.0686   0.7727   1.0000
   3.500   0.6506   0.02183   0.01149  -0.0684   0.7568   1.0000
   3.750   0.6768   0.02183   0.01161  -0.0675   0.7374   1.0000
   4.000   0.7056   0.02170   0.01169  -0.0667   0.7176   1.0000
   4.250   0.7318   0.02164   0.01177  -0.0656   0.6945   1.0000
   4.500   0.7586   0.02154   0.01180  -0.0645   0.6691   1.0000
   4.750   0.7844   0.02149   0.01189  -0.0632   0.6400   1.0000
   5.000   0.8103   0.02146   0.01194  -0.0619   0.6070   1.0000
   5.250   0.8344   0.02156   0.01220  -0.0603   0.5677   1.0000
   5.500   0.8579   0.02177   0.01239  -0.0587   0.5237   1.0000
   5.750   0.8796   0.02221   0.01275  -0.0570   0.4758   1.0000
   6.000   0.9000   0.02286   0.01327  -0.0554   0.4294   1.0000
   6.250   0.9190   0.02371   0.01401  -0.0538   0.3855   1.0000
   6.500   0.9373   0.02468   0.01488  -0.0522   0.3459   1.0000
   6.750   0.9555   0.02573   0.01590  -0.0508   0.3104   1.0000
   7.000   0.9734   0.02686   0.01700  -0.0494   0.2792   1.0000
   7.250   0.9918   0.02805   0.01834  -0.0481   0.2512   1.0000
   7.500   1.0102   0.02928   0.01968  -0.0469   0.2246   1.0000
   7.750   1.0261   0.03055   0.02104  -0.0453   0.1970   1.0000
   8.000   1.0359   0.03181   0.02239  -0.0433   0.1608   1.0000
   8.250   1.0388   0.03332   0.02355  -0.0414   0.1049   1.0000
   8.500   1.0447   0.03582   0.02566  -0.0395   0.0572   1.0000
   8.750   1.0500   0.03849   0.02827  -0.0372   0.0433   1.0000
   9.000   1.0567   0.04090   0.03084  -0.0349   0.0367   1.0000
   9.250   1.0617   0.04328   0.03331  -0.0330   0.0320   1.0000
   9.500   1.0683   0.04577   0.03603  -0.0311   0.0288   1.0000
   9.750   1.0756   0.04831   0.03888  -0.0294   0.0259   1.0000
  10.000   1.0810   0.05101   0.04184  -0.0278   0.0242   1.0000
  10.250   1.0849   0.05400   0.04509  -0.0264   0.0232   1.0000
  10.500   1.0862   0.05723   0.04857  -0.0252   0.0224   1.0000
  10.750   1.0851   0.06076   0.05235  -0.0242   0.0219   1.0000
  11.000   1.0806   0.06470   0.05668  -0.0237   0.0215   1.0000
  11.250   1.0745   0.06886   0.06112  -0.0236   0.0214   1.0000
  11.500   1.0643   0.07362   0.06613  -0.0241   0.0211   1.0000
  11.750   1.0524   0.07877   0.07155  -0.0255   0.0211   1.0000
  12.000   1.0391   0.08441   0.07743  -0.0277   0.0212   1.0000
  12.250   1.0234   0.09092   0.08417  -0.0309   0.0212   1.0000
  12.500   1.0069   0.09827   0.09177  -0.0353   0.0216   1.0000
  12.750   0.9876   0.10725   0.10086  -0.0410   0.0220   1.0000
  13.000   0.9652   0.11837   0.11217  -0.0482   0.0228   1.0000
<< Back to HQ 2.5/9 AIRFOIL (hq259-il)

Polar data table (+)

Polar graphs


<< Back to HQ 2.5/9 AIRFOIL (hq259-il)