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HQ 1.5/8.5 AIRFOIL (hq1585-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: HQ 1.5/8.5 AIRFOIL (hq1585-il)
Reynolds number: 100,000
Max Cl/Cd: 51.39 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq1585-il-100000.txt
Download as CSV file: xf-hq1585-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 1.5/8.5 AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5424   0.09030   0.08568  -0.0240   1.0000   0.1185
  -8.250  -0.5133   0.08647   0.08176  -0.0183   1.0000   0.1247
  -8.000  -0.5233   0.08291   0.07830  -0.0218   1.0000   0.1303
  -7.500  -0.5226   0.07501   0.07051  -0.0253   1.0000   0.1413
  -7.250  -0.5335   0.06964   0.06513  -0.0317   1.0000   0.1489
  -7.000  -0.5407   0.06575   0.06110  -0.0363   1.0000   0.1612
  -6.750  -0.5181   0.06312   0.05863  -0.0315   1.0000   0.1685
  -6.500  -0.5131   0.05983   0.05533  -0.0319   1.0000   0.1833
  -6.000  -0.4993   0.04010   0.03379  -0.0420   1.0000   0.0898
  -5.750  -0.4830   0.03532   0.02822  -0.0406   1.0000   0.0769
  -5.500  -0.4667   0.03160   0.02426  -0.0394   1.0000   0.0737
  -5.250  -0.4485   0.02845   0.02061  -0.0378   1.0000   0.0705
  -5.000  -0.4290   0.02600   0.01767  -0.0362   1.0000   0.0694
  -4.750  -0.4096   0.02443   0.01585  -0.0347   1.0000   0.0733
  -4.500  -0.3889   0.02290   0.01399  -0.0332   1.0000   0.0759
  -4.250  -0.3673   0.02152   0.01231  -0.0317   1.0000   0.0777
  -4.000  -0.3462   0.01975   0.01055  -0.0305   1.0000   0.0812
  -3.750  -0.3256   0.01883   0.00960  -0.0293   1.0000   0.0898
  -3.500  -0.3051   0.01765   0.00854  -0.0281   1.0000   0.0984
  -3.250  -0.2843   0.01666   0.00762  -0.0270   1.0000   0.1124
  -3.000  -0.2628   0.01553   0.00680  -0.0262   1.0000   0.1603
  -2.750  -0.2508   0.01305   0.00670  -0.0234   1.0000   0.6092
  -2.500  -0.2391   0.01313   0.00702  -0.0193   1.0000   0.7186
  -2.250  -0.2275   0.01321   0.00714  -0.0154   1.0000   0.7828
  -2.000  -0.2176   0.01317   0.00715  -0.0110   1.0000   0.8359
  -1.750  -0.2079   0.01306   0.00713  -0.0063   1.0000   0.8988
  -1.500  -0.1401   0.01311   0.00701  -0.0131   1.0000   0.9899
  -1.250  -0.1224   0.01303   0.00675  -0.0134   1.0000   1.0000
  -1.000  -0.1029   0.01308   0.00662  -0.0136   1.0000   1.0000
  -0.750  -0.0808   0.01324   0.00660  -0.0140   1.0000   1.0000
  -0.500  -0.0405   0.01363   0.00679  -0.0177   0.9944   1.0000
  -0.250   0.0052   0.01402   0.00704  -0.0222   0.9855   1.0000
   0.000   0.0492   0.01440   0.00730  -0.0262   0.9765   1.0000
   0.250   0.0973   0.01482   0.00762  -0.0309   0.9686   1.0000
   0.500   0.1376   0.01510   0.00784  -0.0340   0.9582   1.0000
   0.750   0.1833   0.01535   0.00805  -0.0379   0.9466   1.0000
   1.000   0.2314   0.01551   0.00820  -0.0419   0.9346   1.0000
   1.250   0.2818   0.01560   0.00831  -0.0463   0.9238   1.0000
   1.500   0.3393   0.01547   0.00826  -0.0516   0.9127   1.0000
   1.750   0.3830   0.01536   0.00821  -0.0542   0.8993   1.0000
   2.000   0.4208   0.01531   0.00822  -0.0556   0.8860   1.0000
   2.250   0.4555   0.01526   0.00827  -0.0563   0.8723   1.0000
   2.500   0.4883   0.01516   0.00825  -0.0564   0.8579   1.0000
   2.750   0.5192   0.01503   0.00821  -0.0560   0.8428   1.0000
   3.000   0.5465   0.01493   0.00819  -0.0549   0.8253   1.0000
   3.250   0.5720   0.01480   0.00817  -0.0533   0.8051   1.0000
   3.500   0.5977   0.01454   0.00797  -0.0513   0.7827   1.0000
   3.750   0.6220   0.01429   0.00776  -0.0491   0.7569   1.0000
   4.000   0.6445   0.01407   0.00762  -0.0468   0.7240   1.0000
   4.250   0.6667   0.01373   0.00732  -0.0442   0.6796   1.0000
   4.500   0.6883   0.01355   0.00697  -0.0416   0.6174   1.0000
   4.750   0.7086   0.01379   0.00692  -0.0392   0.5366   1.0000
   5.000   0.7285   0.01440   0.00713  -0.0373   0.4560   1.0000
   5.250   0.7484   0.01521   0.00764  -0.0358   0.3877   1.0000
   5.500   0.7676   0.01614   0.00826  -0.0344   0.3185   1.0000
   5.750   0.7830   0.01760   0.00915  -0.0327   0.2123   1.0000
   6.000   0.7976   0.01962   0.01045  -0.0308   0.1380   1.0000
   6.250   0.8182   0.02110   0.01183  -0.0295   0.1177   1.0000
   6.500   0.8406   0.02268   0.01335  -0.0285   0.1069   1.0000
   6.750   0.8639   0.02432   0.01494  -0.0277   0.0961   1.0000
   7.000   0.8875   0.02619   0.01673  -0.0272   0.0848   1.0000
   7.250   0.9095   0.02766   0.01831  -0.0264   0.0727   1.0000
   7.500   0.9327   0.02973   0.02070  -0.0253   0.0636   1.0000
   7.750   0.9541   0.03250   0.02369  -0.0243   0.0563   1.0000
   8.000   0.9751   0.03550   0.02711  -0.0230   0.0529   1.0000
   8.250   0.9938   0.03879   0.03077  -0.0217   0.0510   1.0000
   8.500   1.0093   0.04234   0.03464  -0.0206   0.0491   1.0000
   8.750   1.0169   0.04811   0.04079  -0.0194   0.0477   1.0000
   9.000   1.0211   0.05111   0.04438  -0.0172   0.0468   1.0000
   9.250   1.0228   0.05567   0.04935  -0.0155   0.0470   1.0000
   9.500   1.0274   0.06069   0.05465  -0.0143   0.0480   1.0000
   9.750   0.9900   0.06590   0.06081  -0.0109   0.0525   1.0000
  10.000   0.9644   0.07083   0.06598  -0.0094   0.0542   1.0000
  10.250   0.9403   0.07577   0.07106  -0.0098   0.0553   1.0000
  10.500   0.9177   0.08149   0.07689  -0.0121   0.0564   1.0000
  10.750   0.9057   0.08738   0.08281  -0.0145   0.0580   1.0000
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