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GOE 101 AIRFOIL (goe101-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 101 AIRFOIL (goe101-il)
Reynolds number: 500,000
Max Cl/Cd: 104.4 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe101-il-500000.txt
Download as CSV file: xf-goe101-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 101 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.3790   0.09443   0.09222  -0.0190   1.0000   0.0153
  -8.000  -0.3765   0.09153   0.08934  -0.0202   1.0000   0.0153
  -7.750  -0.3763   0.08867   0.08652  -0.0212   1.0000   0.0153
  -7.500  -0.3723   0.08546   0.08334  -0.0231   1.0000   0.0154
  -7.250  -0.3656   0.08175   0.07965  -0.0262   1.0000   0.0154
  -7.000  -0.3574   0.07821   0.07613  -0.0288   1.0000   0.0154
  -6.750  -0.3592   0.07367   0.07162  -0.0275   1.0000   0.0158
  -6.500  -0.3525   0.07101   0.06898  -0.0273   1.0000   0.0162
  -6.250  -0.3448   0.06821   0.06619  -0.0283   1.0000   0.0166
  -6.000  -0.3370   0.06538   0.06337  -0.0294   1.0000   0.0171
  -5.750  -0.3295   0.06245   0.06044  -0.0305   1.0000   0.0178
  -5.500  -0.2875   0.05711   0.05498  -0.0401   0.9959   0.0198
  -5.250  -0.2337   0.05145   0.04909  -0.0507   0.9903   0.0208
  -5.000  -0.1954   0.04621   0.04367  -0.0567   0.9852   0.0209
  -4.750  -0.1591   0.04109   0.03832  -0.0615   0.9793   0.0209
  -4.500  -0.1326   0.03454   0.03164  -0.0664   0.9724   0.0221
  -4.250  -0.1011   0.03220   0.02921  -0.0690   0.9638   0.0234
  -4.000  -0.0650   0.02938   0.02617  -0.0717   0.9542   0.0262
  -3.750  -0.0253   0.02811   0.02447  -0.0724   0.9408   0.0287
  -3.500  -0.0006   0.02196   0.01780  -0.0735   0.9225   0.0299
  -3.250   0.0245   0.02031   0.01606  -0.0737   0.9014   0.0314
  -3.000   0.0496   0.01912   0.01467  -0.0733   0.8776   0.0337
  -2.750   0.0785   0.02028   0.01551  -0.0720   0.8510   0.0393
  -2.500   0.1001   0.01639   0.01127  -0.0719   0.8252   0.0431
  -2.250   0.1252   0.01569   0.01034  -0.0712   0.7969   0.0474
  -2.000   0.1504   0.01477   0.00910  -0.0705   0.7700   0.0556
  -1.250   0.2305   0.01133   0.00476  -0.0680   0.7016   0.0437
  -1.000   0.2569   0.01072   0.00398  -0.0674   0.6806   0.0424
  -0.750   0.2830   0.01028   0.00341  -0.0668   0.6594   0.0422
  -0.500   0.3092   0.00993   0.00295  -0.0663   0.6386   0.0416
  -0.250   0.3352   0.00960   0.00251  -0.0658   0.6182   0.0421
   0.000   0.3614   0.00939   0.00220  -0.0654   0.5999   0.0430
   0.250   0.3879   0.00924   0.00197  -0.0650   0.5828   0.0439
   0.500   0.4145   0.00915   0.00181  -0.0647   0.5667   0.0447
   0.750   0.4412   0.00911   0.00170  -0.0644   0.5521   0.0467
   1.000   0.4680   0.00909   0.00160  -0.0641   0.5390   0.0482
   1.250   0.4948   0.00911   0.00153  -0.0638   0.5268   0.0496
   1.500   0.5217   0.00913   0.00149  -0.0635   0.5150   0.0535
   1.750   0.5542   0.00701   0.00165  -0.0652   0.5040   1.0000
   2.000   0.5809   0.00713   0.00167  -0.0648   0.4956   1.0000
   2.250   0.6075   0.00727   0.00172  -0.0645   0.4873   1.0000
   2.500   0.6343   0.00738   0.00179  -0.0643   0.4780   1.0000
   2.750   0.6608   0.00751   0.00187  -0.0640   0.4690   1.0000
   3.000   0.6875   0.00763   0.00195  -0.0637   0.4602   1.0000
   3.250   0.7143   0.00776   0.00206  -0.0635   0.4533   1.0000
   3.500   0.7410   0.00789   0.00217  -0.0632   0.4461   1.0000
   3.750   0.7677   0.00801   0.00231  -0.0630   0.4397   1.0000
   4.000   0.7944   0.00814   0.00243  -0.0627   0.4322   1.0000
   4.250   0.8211   0.00827   0.00257  -0.0625   0.4244   1.0000
   4.500   0.8474   0.00842   0.00271  -0.0622   0.4140   1.0000
   5.000   0.9001   0.00869   0.00299  -0.0616   0.3836   1.0000
   5.250   0.9260   0.00887   0.00314  -0.0613   0.3611   1.0000
   5.500   0.9514   0.00913   0.00332  -0.0609   0.3300   1.0000
   5.750   0.9733   0.00984   0.00368  -0.0602   0.2510   1.0000
   6.000   0.9936   0.01083   0.00428  -0.0593   0.1811   1.0000
   6.250   1.0133   0.01191   0.00493  -0.0584   0.1029   1.0000
   6.500   1.0301   0.01340   0.00594  -0.0569   0.0280   1.0000
   6.750   1.0523   0.01416   0.00676  -0.0560   0.0212   1.0000
   7.000   1.0755   0.01472   0.00743  -0.0552   0.0192   1.0000
   7.250   1.0974   0.01543   0.00823  -0.0543   0.0173   1.0000
   7.500   1.1171   0.01640   0.00929  -0.0531   0.0159   1.0000
   7.750   1.1310   0.01802   0.01108  -0.0511   0.0147   1.0000
   8.000   1.1490   0.01907   0.01223  -0.0497   0.0143   1.0000
   8.250   1.1673   0.02006   0.01331  -0.0483   0.0138   1.0000
   8.500   1.1855   0.02103   0.01439  -0.0469   0.0130   1.0000
   8.750   1.2028   0.02207   0.01551  -0.0455   0.0123   1.0000
   9.000   1.2184   0.02335   0.01687  -0.0439   0.0118   1.0000
   9.250   1.2335   0.02471   0.01833  -0.0423   0.0114   1.0000
   9.500   1.2481   0.02619   0.01990  -0.0407   0.0111   1.0000
   9.750   1.2620   0.02789   0.02168  -0.0390   0.0107   1.0000
  10.000   1.2759   0.02989   0.02382  -0.0375   0.0105   1.0000
  10.250   1.2898   0.03222   0.02631  -0.0360   0.0104   1.0000
  10.500   1.3023   0.03476   0.02904  -0.0344   0.0104   1.0000
  10.750   1.2374   0.03050   0.02577  -0.0254   0.0122   1.0000
  11.000   1.2166   0.03976   0.03545  -0.0223   0.0142   1.0000
  11.250   1.2064   0.04082   0.03667  -0.0196   0.0138   1.0000
  11.500   1.1855   0.04710   0.04351  -0.0167   0.0184   1.0000
  11.750   1.1659   0.05150   0.04809  -0.0161   0.0183   1.0000
  12.000   1.1459   0.05644   0.05321  -0.0164   0.0181   1.0000
  12.250   1.1212   0.06288   0.05983  -0.0178   0.0184   1.0000
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