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HT22 (ht22-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: HT22 (ht22-il)
Reynolds number: 500,000
Max Cl/Cd: 60.68 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ht22-il-500000.txt
Download as CSV file: xf-ht22-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HT22                                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.6645   0.09847   0.09630   0.0323   1.0001   0.0147
  -8.000  -0.6614   0.09400   0.09185   0.0290   1.0001   0.0147
  -7.750  -0.6644   0.08744   0.08533   0.0245   1.0001   0.0150
  -7.500  -0.6572   0.08331   0.08120   0.0215   1.0001   0.0153
  -7.250  -0.6469   0.07930   0.07718   0.0181   1.0001   0.0155
  -7.000  -0.6345   0.07498   0.07283   0.0141   1.0001   0.0159
  -6.750  -0.6198   0.07022   0.06803   0.0093   1.0001   0.0164
  -6.500  -0.6026   0.06503   0.06278   0.0043   1.0001   0.0171
  -6.250  -0.5824   0.05926   0.05690  -0.0009   1.0001   0.0181
  -6.000  -0.5488   0.05193   0.04926  -0.0071   1.0001   0.0200
  -5.750  -0.5262   0.04631   0.04335  -0.0096   1.0001   0.0201
  -5.500  -0.5134   0.03896   0.03581  -0.0122   1.0001   0.0211
  -5.250  -0.4918   0.03699   0.03377  -0.0128   1.0001   0.0220
  -5.000  -0.4677   0.03450   0.03113  -0.0134   1.0001   0.0239
  -4.750  -0.4359   0.03273   0.02894  -0.0130   1.0001   0.0271
  -4.500  -0.4137   0.02585   0.02145  -0.0140   1.0001   0.0282
  -4.250  -0.3866   0.01807   0.01293  -0.0134   1.0001   0.0182
  -4.000  -0.3590   0.01537   0.00980  -0.0129   1.0001   0.0183
  -3.750  -0.3314   0.01381   0.00799  -0.0125   1.0001   0.0189
  -3.500  -0.3034   0.01345   0.00749  -0.0122   1.0001   0.0200
  -3.250  -0.2768   0.01135   0.00525  -0.0119   1.0001   0.0217
  -3.000  -0.2494   0.01064   0.00451  -0.0116   1.0001   0.0233
  -2.750  -0.2218   0.01005   0.00388  -0.0113   1.0001   0.0254
  -2.500  -0.1941   0.00970   0.00350  -0.0111   1.0001   0.0274
  -2.250  -0.1667   0.00890   0.00270  -0.0109   1.0001   0.0327
  -2.000  -0.1389   0.00860   0.00238  -0.0107   1.0001   0.0368
  -1.750  -0.1112   0.00812   0.00193  -0.0105   1.0001   0.0471
  -1.500  -0.0834   0.00773   0.00160  -0.0103   1.0001   0.0650
  -1.250  -0.0559   0.00720   0.00140  -0.0103   1.0001   0.1424
  -1.000  -0.0294   0.00622   0.00127  -0.0105   1.0001   0.3875
  -0.750  -0.0062   0.00497   0.00123  -0.0099   1.0001   0.7234
  -0.500   0.0229   0.00429   0.00123  -0.0092   1.0001   0.9999
  -0.250   0.0501   0.00430   0.00118  -0.0089   1.0001   0.9999
   0.000   0.0887   0.00433   0.00116  -0.0111   0.9642   0.9999
   0.250   0.1183   0.00451   0.00109  -0.0107   0.8588   0.9999
   0.500   0.1400   0.00492   0.00104  -0.0088   0.7464   0.9999
   0.750   0.1651   0.00533   0.00103  -0.0080   0.6506   0.9999
   1.000   0.1917   0.00567   0.00104  -0.0076   0.5760   0.9999
   1.250   0.2189   0.00595   0.00106  -0.0074   0.5184   0.9999
   1.500   0.2463   0.00619   0.00111  -0.0072   0.4726   0.9999
   1.750   0.2738   0.00642   0.00116  -0.0071   0.4345   0.9999
   2.000   0.3014   0.00662   0.00123  -0.0070   0.4000   0.9999
   2.250   0.3290   0.00683   0.00130  -0.0068   0.3680   0.9999
   2.500   0.3566   0.00704   0.00138  -0.0067   0.3365   0.9999
   3.000   0.4118   0.00749   0.00160  -0.0065   0.2734   0.9999
   3.250   0.4394   0.00775   0.00173  -0.0065   0.2392   0.9999
   3.500   0.4668   0.00805   0.00188  -0.0064   0.2036   0.9999
   3.750   0.4941   0.00838   0.00209  -0.0063   0.1689   0.9999
   4.000   0.5214   0.00875   0.00231  -0.0063   0.1369   0.9999
   4.250   0.5486   0.00913   0.00257  -0.0062   0.1076   0.9999
   4.500   0.5757   0.00957   0.00289  -0.0061   0.0839   0.9999
   4.750   0.6029   0.00998   0.00326  -0.0060   0.0687   0.9999
   5.000   0.6300   0.01049   0.00373  -0.0059   0.0575   0.9999
   5.250   0.6572   0.01083   0.00409  -0.0057   0.0512   0.9999
   5.500   0.6839   0.01145   0.00474  -0.0055   0.0453   0.9999
   5.750   0.7109   0.01186   0.00521  -0.0053   0.0421   0.9999
   6.000   0.7376   0.01237   0.00575  -0.0051   0.0389   0.9999
   6.250   0.7627   0.01357   0.00703  -0.0048   0.0356   0.9999
   6.500   0.7895   0.01393   0.00748  -0.0046   0.0342   0.9999
   6.750   0.8158   0.01453   0.00817  -0.0043   0.0324   0.9999
   7.000   0.8418   0.01520   0.00891  -0.0040   0.0308   0.9999
   7.250   0.8674   0.01594   0.00971  -0.0038   0.0292   0.9999
   7.500   0.8897   0.01807   0.01200  -0.0033   0.0270   0.9999
   7.750   0.9151   0.01884   0.01293  -0.0030   0.0262   0.9999
   8.000   0.9402   0.01975   0.01401  -0.0026   0.0253   0.9999
   8.250   0.9645   0.02090   0.01533  -0.0023   0.0242   0.9999
   8.500   0.9886   0.02190   0.01647  -0.0020   0.0230   0.9999
   8.750   1.0129   0.02260   0.01727  -0.0017   0.0218   0.9999
   9.000   1.0347   0.02412   0.01891  -0.0015   0.0207   0.9999
   9.250   1.0475   0.02877   0.02408  -0.0009   0.0197   0.9999
   9.500   1.0683   0.03026   0.02584  -0.0006   0.0192   0.9999
   9.750   1.0850   0.03279   0.02872  -0.0002   0.0185   0.9999
  10.000   1.0990   0.03564   0.03192   0.0000   0.0178   0.9999
  10.250   1.1128   0.03807   0.03461   0.0000   0.0170   0.9999
  10.500   1.1280   0.03981   0.03651   0.0000   0.0163   0.9999
  10.750   1.1383   0.04225   0.03916  -0.0002   0.0158   0.9999
  11.000   1.1438   0.04524   0.04235  -0.0008   0.0155   0.9999
  11.250   1.1369   0.04990   0.04726  -0.0027   0.0153   0.9999
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