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HQ 2.5/9 AIRFOIL (hq259-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: HQ 2.5/9 AIRFOIL (hq259-il)
Reynolds number: 500,000
Max Cl/Cd: 106.8 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq259-il-500000.txt
Download as CSV file: xf-hq259-il-500000.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 =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.3970   0.08519   0.08301  -0.0351   1.0000   0.0152
  -8.250  -0.3961   0.08187   0.07972  -0.0365   1.0000   0.0153
  -8.000  -0.3978   0.07844   0.07633  -0.0379   1.0000   0.0154
  -7.750  -0.4054   0.07526   0.07322  -0.0383   1.0000   0.0155
  -7.500  -0.4209   0.07276   0.07079  -0.0372   1.0000   0.0154
  -7.250  -0.4294   0.06844   0.06649  -0.0409   0.9990   0.0155
  -6.750  -0.3352   0.03308   0.03076  -0.0687   0.9828   0.0162
  -6.500  -0.3218   0.02634   0.02379  -0.0735   0.9773   0.0169
  -6.250  -0.2981   0.02278   0.02010  -0.0766   0.9732   0.0178
  -6.000  -0.2718   0.01916   0.01627  -0.0795   0.9698   0.0190
  -5.750  -0.2491   0.01589   0.01273  -0.0808   0.9631   0.0208
  -5.500  -0.2460   0.02205   0.01780  -0.0820   0.9679   0.0133
  -5.250  -0.2198   0.01803   0.01322  -0.0819   0.9610   0.0127
  -5.000  -0.1879   0.01636   0.01128  -0.0828   0.9571   0.0137
  -4.750  -0.1591   0.01506   0.00976  -0.0829   0.9516   0.0146
  -4.500  -0.1319   0.01321   0.00765  -0.0826   0.9453   0.0143
  -4.250  -0.1046   0.01190   0.00616  -0.0823   0.9396   0.0145
  -4.000  -0.0790   0.01101   0.00516  -0.0816   0.9321   0.0152
  -3.750  -0.0531   0.01000   0.00399  -0.0811   0.9257   0.0174
  -3.500  -0.0268   0.00952   0.00345  -0.0807   0.9183   0.0232
  -3.250   0.0000   0.00905   0.00302  -0.0803   0.9117   0.0472
  -3.000   0.0267   0.00882   0.00276  -0.0800   0.9043   0.0623
  -2.750   0.0534   0.00849   0.00250  -0.0798   0.8974   0.0916
  -2.500   0.0781   0.00766   0.00223  -0.0797   0.8895   0.2509
  -2.250   0.1018   0.00675   0.00207  -0.0794   0.8812   0.4894
  -2.000   0.1277   0.00650   0.00202  -0.0789   0.8738   0.5777
  -1.750   0.1541   0.00641   0.00197  -0.0784   0.8652   0.6236
  -1.500   0.1810   0.00636   0.00193  -0.0781   0.8582   0.6571
  -1.250   0.2078   0.00632   0.00192  -0.0776   0.8502   0.6861
  -1.000   0.2347   0.00630   0.00190  -0.0773   0.8427   0.7104
  -0.750   0.2612   0.00629   0.00188  -0.0767   0.8340   0.7359
  -0.500   0.2867   0.00625   0.00190  -0.0759   0.8237   0.7661
  -0.250   0.3128   0.00623   0.00187  -0.0753   0.8132   0.7850
   0.000   0.3395   0.00621   0.00183  -0.0749   0.8031   0.7987
   0.250   0.3664   0.00618   0.00179  -0.0746   0.7932   0.8102
   0.500   0.3932   0.00616   0.00178  -0.0742   0.7833   0.8226
   0.750   0.4199   0.00614   0.00176  -0.0739   0.7733   0.8360
   1.000   0.4462   0.00611   0.00173  -0.0734   0.7625   0.8511
   1.250   0.4719   0.00608   0.00172  -0.0728   0.7509   0.8693
   1.500   0.4970   0.00601   0.00172  -0.0720   0.7387   0.8943
   1.750   0.5258   0.00592   0.00171  -0.0719   0.7261   0.9412
   2.000   0.5638   0.00594   0.00170  -0.0742   0.7117   1.0000
   2.250   0.5910   0.00604   0.00172  -0.0741   0.6957   1.0000
   2.500   0.6180   0.00614   0.00179  -0.0740   0.6773   1.0000
   2.750   0.6448   0.00626   0.00185  -0.0738   0.6587   1.0000
   3.000   0.6713   0.00640   0.00192  -0.0735   0.6374   1.0000
   3.250   0.6974   0.00657   0.00201  -0.0732   0.6121   1.0000
   3.500   0.7230   0.00677   0.00214  -0.0727   0.5817   1.0000
   3.750   0.7477   0.00705   0.00228  -0.0722   0.5424   1.0000
   4.000   0.7715   0.00742   0.00246  -0.0715   0.4945   1.0000
   4.250   0.7948   0.00786   0.00269  -0.0708   0.4439   1.0000
   4.500   0.8182   0.00832   0.00295  -0.0701   0.3969   1.0000
   4.750   0.8418   0.00878   0.00326  -0.0695   0.3512   1.0000
   5.000   0.8639   0.00940   0.00360  -0.0687   0.2936   1.0000
   5.250   0.8859   0.01004   0.00396  -0.0680   0.2412   1.0000
   5.500   0.9092   0.01056   0.00430  -0.0674   0.2024   1.0000
   5.750   0.9328   0.01104   0.00464  -0.0669   0.1702   1.0000
   6.000   0.9555   0.01161   0.00504  -0.0662   0.1338   1.0000
   6.250   0.9755   0.01250   0.00559  -0.0652   0.0789   1.0000
   6.500   0.9903   0.01407   0.00670  -0.0633   0.0117   1.0000
   6.750   1.0135   0.01462   0.00733  -0.0625   0.0092   1.0000
   7.000   1.0359   0.01525   0.00806  -0.0617   0.0080   1.0000
   7.250   1.0567   0.01610   0.00903  -0.0605   0.0070   1.0000
   7.500   1.0721   0.01757   0.01071  -0.0584   0.0063   1.0000
   7.750   1.0899   0.01868   0.01193  -0.0568   0.0061   1.0000
   8.000   1.1064   0.01990   0.01328  -0.0551   0.0060   1.0000
   8.250   1.1217   0.02132   0.01483  -0.0532   0.0060   1.0000
   8.500   1.1357   0.02303   0.01669  -0.0511   0.0060   1.0000
   8.750   1.1494   0.02514   0.01901  -0.0491   0.0061   1.0000
   9.000   1.1634   0.02805   0.02211  -0.0473   0.0063   1.0000
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