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HQ 2.5/9 B AIRFOIL (hq259b-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: HQ 2.5/9 B AIRFOIL (hq259b-il)
Reynolds number: 50,000
Max Cl/Cd: 39.05 at α=6°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq259b-il-50000.txt
Download as CSV file: xf-hq259b-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 2.5/9 B 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4123   0.10678   0.09970  -0.0194   1.0000   0.2418
  -8.750  -0.4351   0.10672   0.09982  -0.0211   1.0000   0.2506
  -8.500  -0.4269   0.10321   0.09635  -0.0201   1.0000   0.2651
  -8.250  -0.4219   0.10015   0.09333  -0.0191   1.0000   0.2798
  -8.000  -0.3975   0.09538   0.08853  -0.0168   1.0000   0.3008
  -7.750  -0.3981   0.09331   0.08655  -0.0154   1.0000   0.3218
  -7.500  -0.3912   0.09031   0.08361  -0.0136   1.0000   0.3429
  -7.250  -0.3916   0.08842   0.08181  -0.0115   1.0000   0.3668
  -7.000  -0.3757   0.08493   0.07834  -0.0094   1.0000   0.3914
  -6.750  -0.3733   0.08260   0.07610  -0.0069   1.0000   0.4163
  -6.500  -0.3853   0.08184   0.07548  -0.0033   1.0000   0.4414
  -6.250  -0.3769   0.07926   0.07294  -0.0003   1.0000   0.4709
  -6.000  -0.3461   0.07482   0.06847   0.0011   1.0000   0.5059
  -5.750  -0.3353   0.07234   0.06603   0.0040   1.0000   0.5415
  -4.750  -0.3962   0.04437   0.03643  -0.0465   1.0000   0.1626
  -4.500  -0.3699   0.04042   0.03191  -0.0474   1.0000   0.1456
  -4.250  -0.3466   0.03729   0.02846  -0.0476   1.0000   0.1416
  -4.000  -0.3209   0.03463   0.02530  -0.0479   1.0000   0.1400
  -3.750  -0.2939   0.03222   0.02234  -0.0479   1.0000   0.1373
  -3.500  -0.2669   0.03019   0.01983  -0.0476   1.0000   0.1361
  -3.250  -0.2413   0.02850   0.01786  -0.0471   1.0000   0.1385
  -3.000  -0.2156   0.02726   0.01626  -0.0464   1.0000   0.1462
  -2.750  -0.1915   0.02579   0.01480  -0.0454   1.0000   0.1537
  -2.500  -0.1680   0.02469   0.01357  -0.0441   1.0000   0.1634
  -2.250  -0.1449   0.02356   0.01252  -0.0430   1.0000   0.1850
  -2.000  -0.1187   0.02224   0.01140  -0.0427   1.0000   0.2248
  -1.750  -0.1121   0.01943   0.01149  -0.0361   1.0000   0.7449
  -1.500  -0.1107   0.01862   0.01095  -0.0269   1.0000   1.0000
  -1.250  -0.0847   0.01873   0.01051  -0.0278   1.0000   1.0000
  -1.000  -0.0586   0.01889   0.01025  -0.0288   1.0000   1.0000
  -0.750  -0.0328   0.01912   0.01012  -0.0296   1.0000   1.0000
  -0.500  -0.0075   0.01939   0.01009  -0.0303   1.0000   1.0000
  -0.250   0.0172   0.01969   0.01013  -0.0308   1.0000   1.0000
   0.000   0.0413   0.02004   0.01025  -0.0312   1.0000   1.0000
   0.250   0.0650   0.02042   0.01045  -0.0316   1.0000   1.0000
   0.500   0.0881   0.02084   0.01071  -0.0318   1.0000   1.0000
   0.750   0.1108   0.02130   0.01102  -0.0320   1.0000   1.0000
   1.000   0.1330   0.02180   0.01141  -0.0322   1.0000   1.0000
   1.250   0.1547   0.02235   0.01187  -0.0324   1.0000   1.0000
   1.500   0.1759   0.02295   0.01240  -0.0325   1.0000   1.0000
   1.750   0.1964   0.02361   0.01302  -0.0327   1.0000   1.0000
   2.000   0.2162   0.02435   0.01373  -0.0329   1.0000   1.0000
   2.250   0.2351   0.02518   0.01456  -0.0331   1.0000   1.0000
   2.500   0.2945   0.02684   0.01627  -0.0409   0.9779   1.0000
   2.750   0.3556   0.02830   0.01781  -0.0484   0.9530   1.0000
   3.000   0.4167   0.02946   0.01910  -0.0554   0.9270   1.0000
   3.250   0.4710   0.03024   0.02005  -0.0606   0.8988   1.0000
   3.500   0.5246   0.03076   0.02076  -0.0651   0.8699   1.0000
   3.750   0.5781   0.03096   0.02121  -0.0689   0.8407   1.0000
   4.000   0.6319   0.03082   0.02132  -0.0721   0.8117   1.0000
   4.250   0.6854   0.03029   0.02111  -0.0745   0.7832   1.0000
   4.500   0.7339   0.02948   0.02056  -0.0754   0.7545   1.0000
   4.750   0.7799   0.02836   0.01969  -0.0753   0.7250   1.0000
   5.000   0.8249   0.02681   0.01840  -0.0743   0.6932   1.0000
   5.250   0.8569   0.02573   0.01746  -0.0718   0.6545   1.0000
   5.500   0.8884   0.02461   0.01638  -0.0691   0.6118   1.0000
   5.750   0.9172   0.02389   0.01558  -0.0663   0.5637   1.0000
   6.000   0.9407   0.02409   0.01556  -0.0636   0.5089   1.0000
   6.250   0.9614   0.02500   0.01611  -0.0609   0.4471   1.0000
   6.500   0.9768   0.02648   0.01720  -0.0579   0.3761   1.0000
   6.750   0.9909   0.02828   0.01856  -0.0550   0.3065   1.0000
   7.000   1.0064   0.02994   0.01985  -0.0526   0.2535   1.0000
   7.250   1.0232   0.03161   0.02126  -0.0507   0.2128   1.0000
   7.500   1.0435   0.03369   0.02319  -0.0493   0.1817   1.0000
   7.750   1.0663   0.03610   0.02562  -0.0483   0.1591   1.0000
   8.000   1.0909   0.03889   0.02837  -0.0477   0.1431   1.0000
   8.250   1.1113   0.04213   0.03214  -0.0463   0.1340   1.0000
   8.500   1.1300   0.04539   0.03562  -0.0452   0.1256   1.0000
   8.750   1.1434   0.04883   0.03957  -0.0435   0.1200   1.0000
   9.000   1.1598   0.05272   0.04371  -0.0423   0.1168   1.0000
   9.250   1.1706   0.05735   0.04863  -0.0410   0.1145   1.0000
   9.500   1.1676   0.06168   0.05357  -0.0385   0.1135   1.0000
   9.750   1.1599   0.06626   0.05865  -0.0362   0.1129   1.0000
  10.000   1.1478   0.07104   0.06382  -0.0341   0.1128   1.0000
  10.250   1.1331   0.07594   0.06901  -0.0325   0.1132   1.0000
  10.500   1.1172   0.08088   0.07415  -0.0311   0.1139   1.0000
  10.750   1.1048   0.08603   0.07942  -0.0302   0.1146   1.0000
  11.000   1.0152   0.09638   0.09002  -0.0350   0.1267   1.0000
  11.250   1.0060   0.10317   0.09683  -0.0372   0.1281   1.0000
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