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

HQ 1.5/9 B AIRFOIL (hq159b-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: HQ 1.5/9 B AIRFOIL (hq159b-il)
Reynolds number: 200,000
Max Cl/Cd: 60.59 at α=3.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hq159b-il-200000-n5.txt
Download as CSV file: xf-hq159b-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 1.5/9 B 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.5309   0.08569   0.08215  -0.0229   1.0000   0.0117
  -9.250  -0.5363   0.08026   0.07677  -0.0260   1.0000   0.0110
  -9.000  -0.5447   0.07404   0.07067  -0.0303   1.0000   0.0112
  -8.750  -0.5595   0.06616   0.06285  -0.0375   1.0000   0.0109
  -8.500  -0.5767   0.05974   0.05636  -0.0420   1.0000   0.0108
  -8.250  -0.5901   0.05294   0.04938  -0.0445   1.0000   0.0103
  -8.000  -0.5961   0.04708   0.04326  -0.0452   1.0000   0.0104
  -7.750  -0.5964   0.04192   0.03778  -0.0449   1.0000   0.0110
  -7.500  -0.5957   0.03636   0.03178  -0.0438   1.0000   0.0111
  -7.250  -0.5913   0.03124   0.02611  -0.0420   1.0000   0.0111
  -7.000  -0.5815   0.02727   0.02158  -0.0401   1.0000   0.0114
  -6.750  -0.5679   0.02428   0.01811  -0.0382   1.0000   0.0118
  -6.500  -0.5522   0.02200   0.01543  -0.0364   1.0000   0.0123
  -6.250  -0.5246   0.01995   0.01297  -0.0368   0.9965   0.0131
  -6.000  -0.4929   0.01839   0.01112  -0.0379   0.9915   0.0142
  -5.750  -0.4607   0.01742   0.01011  -0.0394   0.9862   0.0171
  -5.500  -0.4266   0.01643   0.00894  -0.0410   0.9820   0.0208
  -5.250  -0.3956   0.01552   0.00799  -0.0420   0.9758   0.0250
  -5.000  -0.3618   0.01486   0.00726  -0.0436   0.9711   0.0311
  -4.750  -0.3284   0.01464   0.00697  -0.0449   0.9656   0.0390
  -4.500  -0.2970   0.01389   0.00615  -0.0460   0.9592   0.0449
  -4.250  -0.2636   0.01340   0.00554  -0.0473   0.9539   0.0502
  -4.000  -0.2337   0.01290   0.00504  -0.0479   0.9461   0.0589
  -3.750  -0.2013   0.01246   0.00455  -0.0490   0.9402   0.0692
  -3.500  -0.1721   0.01205   0.00415  -0.0494   0.9318   0.0843
  -3.250  -0.1421   0.01157   0.00374  -0.0501   0.9249   0.1167
  -3.000  -0.1150   0.01083   0.00343  -0.0505   0.9165   0.2243
  -2.750  -0.0900   0.00994   0.00319  -0.0505   0.9080   0.3884
  -2.500  -0.0639   0.00941   0.00311  -0.0502   0.9002   0.5154
  -2.250  -0.0385   0.00919   0.00311  -0.0495   0.8906   0.5967
  -2.000  -0.0118   0.00910   0.00307  -0.0489   0.8816   0.6479
  -1.750   0.0148   0.00903   0.00303  -0.0483   0.8728   0.6830
  -1.500   0.0406   0.00899   0.00301  -0.0475   0.8628   0.7119
  -1.250   0.0670   0.00896   0.00295  -0.0469   0.8535   0.7335
  -1.000   0.0933   0.00893   0.00291  -0.0462   0.8450   0.7526
  -0.750   0.1195   0.00891   0.00289  -0.0456   0.8360   0.7692
  -0.500   0.1460   0.00889   0.00285  -0.0451   0.8272   0.7828
  -0.250   0.1722   0.00886   0.00279  -0.0444   0.8168   0.7967
   0.000   0.1980   0.00883   0.00275  -0.0437   0.8042   0.8101
   0.250   0.2241   0.00880   0.00271  -0.0431   0.7919   0.8222
   0.500   0.2503   0.00878   0.00268  -0.0426   0.7798   0.8341
   0.750   0.2764   0.00876   0.00265  -0.0419   0.7667   0.8470
   1.000   0.3025   0.00873   0.00262  -0.0413   0.7521   0.8611
   1.250   0.3292   0.00872   0.00262  -0.0408   0.7375   0.8768
   1.500   0.3568   0.00871   0.00263  -0.0405   0.7255   0.8956
   1.750   0.3869   0.00870   0.00266  -0.0407   0.7133   0.9193
   2.000   0.4217   0.00871   0.00269  -0.0420   0.6990   0.9480
   2.250   0.4596   0.00873   0.00273  -0.0441   0.6823   0.9843
   2.500   0.4887   0.00882   0.00279  -0.0445   0.6655   1.0000
   2.750   0.5147   0.00894   0.00289  -0.0441   0.6474   1.0000
   3.000   0.5403   0.00908   0.00299  -0.0436   0.6216   1.0000
   3.250   0.5641   0.00931   0.00304  -0.0427   0.5685   1.0000
   3.500   0.5850   0.00981   0.00311  -0.0414   0.4827   1.0000
   3.750   0.6061   0.01044   0.00335  -0.0403   0.4049   1.0000
   4.000   0.6280   0.01107   0.00364  -0.0395   0.3315   1.0000
   4.250   0.6508   0.01165   0.00401  -0.0389   0.2813   1.0000
   4.500   0.6748   0.01212   0.00436  -0.0384   0.2496   1.0000
   4.750   0.6990   0.01258   0.00473  -0.0380   0.2203   1.0000
   5.000   0.7231   0.01306   0.00511  -0.0375   0.1852   1.0000
   5.250   0.7461   0.01369   0.00552  -0.0370   0.1331   1.0000
   5.500   0.7678   0.01451   0.00610  -0.0363   0.0888   1.0000
   5.750   0.7900   0.01530   0.00680  -0.0356   0.0593   1.0000
   6.000   0.8130   0.01599   0.00745  -0.0349   0.0458   1.0000
   6.250   0.8364   0.01660   0.00810  -0.0343   0.0400   1.0000
   6.500   0.8587   0.01734   0.00891  -0.0336   0.0365   1.0000
   6.750   0.8808   0.01809   0.00977  -0.0327   0.0345   1.0000
   7.000   0.9030   0.01880   0.01061  -0.0320   0.0318   1.0000
   7.250   0.9237   0.01967   0.01151  -0.0312   0.0280   1.0000
   7.500   0.9438   0.02066   0.01260  -0.0302   0.0251   1.0000
   7.750   0.9649   0.02147   0.01351  -0.0294   0.0219   1.0000
   8.000   0.9820   0.02284   0.01499  -0.0281   0.0196   1.0000
   8.250   1.0011   0.02406   0.01637  -0.0270   0.0180   1.0000
   8.500   1.0210   0.02514   0.01764  -0.0260   0.0157   1.0000
   8.750   1.0401   0.02619   0.01883  -0.0250   0.0136   1.0000
   9.000   1.0573   0.02748   0.02024  -0.0238   0.0124   1.0000
   9.250   1.0712   0.02940   0.02231  -0.0224   0.0115   1.0000
   9.500   1.0834   0.03198   0.02513  -0.0207   0.0109   1.0000
   9.750   1.0964   0.03411   0.02756  -0.0191   0.0105   1.0000
  10.000   1.1061   0.03652   0.03030  -0.0173   0.0101   1.0000
  10.500   1.1155   0.04100   0.03543  -0.0129   0.0089   1.0000
  10.750   1.1130   0.04333   0.03803  -0.0103   0.0085   1.0000
  11.000   1.1062   0.04607   0.04104  -0.0079   0.0082   1.0000
  11.250   1.0981   0.04893   0.04414  -0.0062   0.0079   1.0000
  11.500   1.0859   0.05253   0.04799  -0.0052   0.0078   1.0000
  11.750   1.0706   0.05679   0.05249  -0.0053   0.0077   1.0000
  12.000   1.0517   0.06207   0.05802  -0.0066   0.0078   1.0000
  12.250   1.0313   0.06823   0.06440  -0.0094   0.0078   1.0000
  12.500   1.0092   0.07566   0.07204  -0.0139   0.0079   1.0000
  12.750   0.9849   0.08491   0.08147  -0.0203   0.0080   1.0000
  13.000   0.9454   0.10008   0.09681  -0.0308   0.0088   1.0000
  13.250   0.9079   0.11654   0.11333  -0.0406   0.0097   1.0000
<< Back to HQ 1.5/9 B AIRFOIL (hq159b-il)

Polar data table (+)

Polar graphs


<< Back to HQ 1.5/9 B AIRFOIL (hq159b-il)