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HQ 0/9 AIRFOIL (hq09-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: HQ 0/9 AIRFOIL (hq09-il)
Reynolds number: 50,000
Max Cl/Cd: 26 at α=4°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hq09-il-50000-n5.txt
Download as CSV file: xf-hq09-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 0/9 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.6749   0.09497   0.08827  -0.0004   1.0000   0.0413
  -9.500  -0.6810   0.08836   0.08171  -0.0050   1.0000   0.0410
  -9.250  -0.6895   0.08239   0.07576  -0.0092   1.0000   0.0403
  -9.000  -0.7021   0.07702   0.07039  -0.0126   1.0000   0.0397
  -8.750  -0.7163   0.07230   0.06561  -0.0142   1.0000   0.0394
  -8.500  -0.7263   0.06757   0.06088  -0.0155   1.0000   0.0392
  -8.250  -0.7327   0.06289   0.05597  -0.0161   1.0000   0.0393
  -8.000  -0.7355   0.05816   0.05090  -0.0163   1.0000   0.0399
  -7.750  -0.7335   0.05349   0.04569  -0.0160   1.0000   0.0403
  -7.500  -0.7272   0.04894   0.04061  -0.0153   1.0000   0.0410
  -7.250  -0.7167   0.04454   0.03558  -0.0143   1.0000   0.0422
  -7.000  -0.7022   0.04066   0.03108  -0.0133   1.0000   0.0445
  -6.750  -0.6848   0.03847   0.02879  -0.0128   1.0000   0.0512
  -6.500  -0.6637   0.03536   0.02504  -0.0118   1.0000   0.0546
  -6.250  -0.6397   0.03247   0.02170  -0.0108   1.0000   0.0570
  -6.000  -0.6157   0.03019   0.01930  -0.0101   1.0000   0.0608
  -5.750  -0.5906   0.02832   0.01717  -0.0091   1.0000   0.0661
  -5.500  -0.5677   0.02684   0.01561  -0.0083   1.0000   0.0773
  -5.250  -0.5447   0.02530   0.01397  -0.0073   1.0000   0.0873
  -5.000  -0.5228   0.02379   0.01237  -0.0061   1.0000   0.0998
  -4.750  -0.5024   0.02224   0.01082  -0.0051   1.0000   0.1259
  -4.500  -0.4836   0.02051   0.00963  -0.0043   1.0000   0.2004
  -4.250  -0.4698   0.01858   0.00870  -0.0027   1.0000   0.3515
  -4.000  -0.4568   0.01757   0.00860   0.0009   1.0000   0.5281
  -3.750  -0.4405   0.01738   0.00867   0.0046   1.0000   0.6410
  -3.500  -0.4240   0.01743   0.00878   0.0086   1.0000   0.7183
  -3.250  -0.4055   0.01752   0.00883   0.0123   1.0000   0.7732
  -3.000  -0.3834   0.01746   0.00859   0.0146   1.0000   0.8076
  -2.750  -0.3588   0.01731   0.00815   0.0158   1.0000   0.8331
  -2.500  -0.3297   0.01718   0.00784   0.0161   1.0000   0.8557
  -2.250  -0.2957   0.01713   0.00760   0.0157   1.0000   0.8824
  -2.000  -0.2518   0.01716   0.00742   0.0134   1.0000   0.9121
  -1.750  -0.2003   0.01709   0.00714   0.0090   1.0000   0.9345
  -1.500  -0.1564   0.01688   0.00670   0.0054   1.0000   0.9489
  -1.250  -0.1137   0.01665   0.00635   0.0018   1.0000   0.9626
  -1.000  -0.0714   0.01641   0.00602  -0.0019   1.0000   0.9765
  -0.750  -0.0290   0.01616   0.00571  -0.0057   1.0000   0.9910
  -0.500  -0.0013   0.01597   0.00552  -0.0070   1.0000   1.0000
  -0.250   0.0013   0.01587   0.00546  -0.0038   1.0000   1.0000
   0.000   0.0000   0.01583   0.00545   0.0000   1.0000   1.0000
   0.250  -0.0013   0.01587   0.00546   0.0038   1.0000   1.0000
   0.500   0.0013   0.01597   0.00552   0.0070   1.0000   1.0000
   0.750   0.0290   0.01616   0.00571   0.0057   0.9910   1.0000
   1.000   0.0714   0.01641   0.00602   0.0019   0.9765   1.0000
   1.250   0.1137   0.01665   0.00635  -0.0018   0.9626   1.0000
   1.500   0.1563   0.01688   0.00670  -0.0054   0.9490   1.0000
   1.750   0.2003   0.01709   0.00713  -0.0090   0.9345   1.0000
   2.000   0.2519   0.01715   0.00742  -0.0134   0.9121   1.0000
   2.250   0.2956   0.01713   0.00760  -0.0157   0.8825   1.0000
   2.500   0.3296   0.01718   0.00784  -0.0161   0.8558   1.0000
   2.750   0.3587   0.01730   0.00815  -0.0158   0.8332   1.0000
   3.000   0.3833   0.01746   0.00859  -0.0145   0.8077   1.0000
   3.250   0.4054   0.01752   0.00883  -0.0123   0.7733   1.0000
   3.500   0.4239   0.01743   0.00878  -0.0086   0.7184   1.0000
   3.750   0.4404   0.01738   0.00867  -0.0046   0.6409   1.0000
   4.000   0.4568   0.01757   0.00861  -0.0009   0.5287   1.0000
   4.250   0.4697   0.01858   0.00870   0.0027   0.3511   1.0000
   4.500   0.4835   0.02051   0.00963   0.0043   0.2003   1.0000
   4.750   0.5024   0.02225   0.01081   0.0051   0.1255   1.0000
   5.000   0.5228   0.02379   0.01237   0.0061   0.0998   1.0000
   5.250   0.5447   0.02530   0.01396   0.0073   0.0875   1.0000
   5.500   0.5677   0.02683   0.01560   0.0083   0.0771   1.0000
   5.750   0.5906   0.02832   0.01717   0.0091   0.0661   1.0000
   6.000   0.6157   0.03019   0.01930   0.0100   0.0608   1.0000
   6.250   0.6398   0.03249   0.02171   0.0108   0.0570   1.0000
   6.500   0.6638   0.03536   0.02505   0.0118   0.0546   1.0000
   6.750   0.6849   0.03848   0.02880   0.0128   0.0511   1.0000
   7.000   0.7024   0.04066   0.03109   0.0133   0.0445   1.0000
   7.250   0.7167   0.04458   0.03563   0.0143   0.0421   1.0000
   7.500   0.7274   0.04894   0.04062   0.0153   0.0410   1.0000
   7.750   0.7337   0.05349   0.04568   0.0159   0.0403   1.0000
   8.000   0.7357   0.05817   0.05092   0.0162   0.0399   1.0000
   8.250   0.7333   0.06287   0.05594   0.0161   0.0395   1.0000
   8.500   0.7261   0.06765   0.06096   0.0154   0.0390   1.0000
   8.750   0.7163   0.07237   0.06568   0.0141   0.0393   1.0000
   9.000   0.7022   0.07711   0.07047   0.0124   0.0397   1.0000
   9.250   0.6903   0.08238   0.07575   0.0092   0.0404   1.0000
   9.500   0.6810   0.08857   0.08192   0.0047   0.0410   1.0000
   9.750   0.6756   0.09497   0.08827   0.0003   0.0414   1.0000
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