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HQ 2.5/8 AIRFOIL (hq258-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: HQ 2.5/8 AIRFOIL (hq258-il)
Reynolds number: 200,000
Max Cl/Cd: 80.95 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq258-il-200000.txt
Download as CSV file: xf-hq258-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 2.5/8 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
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4003   0.08715   0.08374  -0.0301   1.0000   0.0323
  -7.750  -0.4022   0.08407   0.08073  -0.0313   1.0000   0.0333
  -7.500  -0.4074   0.08122   0.07795  -0.0321   1.0000   0.0338
  -7.250  -0.4131   0.07821   0.07502  -0.0336   1.0000   0.0343
  -7.000  -0.4154   0.07436   0.07123  -0.0381   1.0000   0.0349
  -6.750  -0.4173   0.07079   0.06763  -0.0414   1.0000   0.0354
  -6.500  -0.4196   0.06789   0.06464  -0.0426   1.0000   0.0357
  -6.250  -0.4196   0.06525   0.06190  -0.0426   1.0000   0.0360
  -6.000  -0.4176   0.06252   0.05903  -0.0422   1.0000   0.0361
  -5.750  -0.4131   0.05965   0.05600  -0.0418   1.0000   0.0362
  -5.500  -0.4159   0.05249   0.04892  -0.0420   1.0000   0.0375
  -5.250  -0.4091   0.04952   0.04597  -0.0412   1.0000   0.0386
  -5.000  -0.3890   0.04620   0.04255  -0.0430   0.9983   0.0406
  -4.750  -0.3523   0.04198   0.03803  -0.0478   0.9946   0.0447
  -4.500  -0.3127   0.03720   0.03257  -0.0525   0.9903   0.0506
  -4.250  -0.2813   0.03362   0.02894  -0.0553   0.9864   0.0537
  -4.000  -0.2412   0.03058   0.02529  -0.0587   0.9830   0.0641
  -3.750  -0.1982   0.02307   0.01681  -0.0588   0.9818   0.0266
  -3.500  -0.1657   0.02199   0.01553  -0.0601   0.9765   0.0334
  -3.250  -0.1294   0.01937   0.01249  -0.0612   0.9733   0.0315
  -3.000  -0.0914   0.01736   0.01011  -0.0626   0.9706   0.0306
  -2.750  -0.0525   0.01584   0.00844  -0.0644   0.9683   0.0323
  -2.500  -0.0223   0.01484   0.00736  -0.0646   0.9620   0.0366
  -2.250   0.0148   0.01346   0.00608  -0.0667   0.9582   0.0779
  -2.000   0.0464   0.01108   0.00580  -0.0687   0.9555   0.5955
  -1.750   0.0721   0.01089   0.00587  -0.0677   0.9482   0.7170
  -1.500   0.1068   0.01077   0.00576  -0.0685   0.9437   0.7799
  -1.250   0.1380   0.01064   0.00565  -0.0687   0.9380   0.8210
  -1.000   0.1661   0.01037   0.00551  -0.0677   0.9320   0.8824
  -0.750   0.2234   0.01004   0.00523  -0.0727   0.9312   0.9774
  -0.500   0.2738   0.00985   0.00489  -0.0776   0.9283   1.0000
  -0.250   0.3034   0.00981   0.00475  -0.0783   0.9188   1.0000
   0.000   0.3388   0.00973   0.00457  -0.0800   0.9125   1.0000
   0.250   0.3707   0.00965   0.00442  -0.0807   0.9041   1.0000
   0.500   0.4005   0.00955   0.00425  -0.0808   0.8934   1.0000
   0.750   0.4296   0.00943   0.00406  -0.0807   0.8822   1.0000
   1.000   0.4583   0.00935   0.00392  -0.0805   0.8722   1.0000
   1.250   0.4861   0.00930   0.00383  -0.0801   0.8620   1.0000
   1.500   0.5123   0.00930   0.00382  -0.0795   0.8501   1.0000
   1.750   0.5386   0.00929   0.00379  -0.0788   0.8377   1.0000
   2.000   0.5649   0.00928   0.00375  -0.0782   0.8248   1.0000
   2.250   0.5911   0.00928   0.00374  -0.0775   0.8112   1.0000
   2.500   0.6174   0.00928   0.00373  -0.0769   0.7969   1.0000
   2.750   0.6437   0.00928   0.00376  -0.0762   0.7814   1.0000
   3.000   0.6693   0.00930   0.00380  -0.0754   0.7629   1.0000
   3.250   0.6951   0.00933   0.00382  -0.0746   0.7428   1.0000
   3.500   0.7207   0.00939   0.00388  -0.0739   0.7199   1.0000
   3.750   0.7460   0.00948   0.00396  -0.0730   0.6941   1.0000
   4.000   0.7709   0.00962   0.00405  -0.0721   0.6632   1.0000
   4.250   0.7949   0.00982   0.00424  -0.0710   0.6231   1.0000
   4.500   0.8179   0.01012   0.00440  -0.0698   0.5717   1.0000
   4.750   0.8393   0.01060   0.00464  -0.0684   0.5079   1.0000
   5.000   0.8594   0.01125   0.00501  -0.0670   0.4396   1.0000
   5.250   0.8781   0.01209   0.00548  -0.0655   0.3642   1.0000
   5.500   0.8961   0.01305   0.00600  -0.0641   0.2846   1.0000
   5.750   0.9163   0.01388   0.00663  -0.0631   0.2266   1.0000
   6.000   0.9359   0.01483   0.00722  -0.0621   0.1572   1.0000
   6.250   0.9479   0.01701   0.00850  -0.0601   0.0347   1.0000
   6.500   0.9664   0.01850   0.01007  -0.0582   0.0225   1.0000
   6.750   0.9859   0.01969   0.01140  -0.0568   0.0184   1.0000
   7.000   1.0009   0.02147   0.01332  -0.0549   0.0151   1.0000
   7.250   1.0186   0.02305   0.01504  -0.0532   0.0143   1.0000
   7.500   1.0369   0.02492   0.01706  -0.0515   0.0138   1.0000
   7.750   1.0566   0.02718   0.01951  -0.0500   0.0136   1.0000
   8.000   1.0771   0.02979   0.02238  -0.0487   0.0137   1.0000
   8.250   1.0960   0.03286   0.02578  -0.0472   0.0140   1.0000
   8.500   1.1107   0.03654   0.02988  -0.0454   0.0145   1.0000
   8.750   1.1198   0.04073   0.03458  -0.0432   0.0151   1.0000
   9.000   1.1292   0.04540   0.03955  -0.0414   0.0162   1.0000
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