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GOE 397 AIRFOIL (goe397-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 397 AIRFOIL (goe397-il)
Reynolds number: 200,000
Max Cl/Cd: 81.7 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe397-il-200000.txt
Download as CSV file: xf-goe397-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 397 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.250  -0.3806   0.08734   0.08396  -0.0301   1.0000   0.0280
  -7.000  -0.3769   0.08280   0.07945  -0.0278   1.0000   0.0286
  -6.750  -0.3702   0.07938   0.07606  -0.0268   1.0000   0.0292
  -6.500  -0.3628   0.07633   0.07303  -0.0270   1.0000   0.0300
  -6.250  -0.3543   0.07331   0.07001  -0.0282   1.0000   0.0310
  -6.000  -0.3447   0.07033   0.06705  -0.0297   1.0000   0.0325
  -5.750  -0.3339   0.06734   0.06406  -0.0315   1.0000   0.0339
  -5.500  -0.3197   0.06444   0.06115  -0.0343   1.0000   0.0365
  -5.250  -0.2924   0.06244   0.05898  -0.0404   1.0000   0.0384
  -5.000  -0.2828   0.06024   0.05671  -0.0404   1.0000   0.0386
  -4.750  -0.2694   0.05756   0.05390  -0.0410   0.9993   0.0387
  -4.500  -0.2508   0.05009   0.04652  -0.0446   0.9946   0.0408
  -4.250  -0.2177   0.04659   0.04296  -0.0487   0.9888   0.0438
  -4.000  -0.1527   0.04427   0.03997  -0.0574   0.9835   0.0520
  -3.750  -0.1323   0.03823   0.03409  -0.0603   0.9768   0.0553
  -3.250  -0.0536   0.03187   0.02726  -0.0682   0.9651   0.0696
  -3.000  -0.0137   0.02929   0.02439  -0.0716   0.9590   0.0823
  -2.750   0.0272   0.02702   0.02184  -0.0751   0.9550   0.0958
  -2.500   0.0608   0.02518   0.01979  -0.0768   0.9459   0.1095
  -2.250   0.0968   0.02326   0.01773  -0.0791   0.9398   0.1242
  -2.000   0.1283   0.02176   0.01617  -0.0803   0.9303   0.1427
  -1.500   0.2042   0.01637   0.00945  -0.0805   0.9137   0.0735
  -1.250   0.2348   0.01516   0.00779  -0.0800   0.9023   0.0665
  -1.000   0.2630   0.01416   0.00664  -0.0797   0.8900   0.0672
  -0.750   0.2898   0.01338   0.00578  -0.0790   0.8762   0.0685
  -0.500   0.3160   0.01282   0.00510  -0.0780   0.8620   0.0673
  -0.250   0.3418   0.01233   0.00454  -0.0771   0.8494   0.0672
   0.000   0.3674   0.01188   0.00408  -0.0764   0.8387   0.0681
   0.250   0.3932   0.01155   0.00375  -0.0756   0.8289   0.0703
   0.500   0.4185   0.01135   0.00356  -0.0750   0.8180   0.0743
   0.750   0.4443   0.01120   0.00340  -0.0743   0.8083   0.0811
   1.000   0.4703   0.01103   0.00330  -0.0737   0.7996   0.1064
   1.250   0.5203   0.00896   0.00321  -0.0787   0.7908   1.0000
   1.500   0.5459   0.00909   0.00324  -0.0781   0.7814   1.0000
   1.750   0.5717   0.00922   0.00329  -0.0774   0.7727   1.0000
   2.000   0.5971   0.00934   0.00340  -0.0768   0.7626   1.0000
   2.250   0.6227   0.00947   0.00349  -0.0762   0.7530   1.0000
   2.500   0.6485   0.00959   0.00357  -0.0755   0.7442   1.0000
   2.750   0.6737   0.00970   0.00369  -0.0748   0.7325   1.0000
   3.000   0.6984   0.00975   0.00373  -0.0738   0.7173   1.0000
   3.250   0.7226   0.00980   0.00376  -0.0726   0.6993   1.0000
   3.500   0.7469   0.00987   0.00382  -0.0716   0.6804   1.0000
   3.750   0.7705   0.00994   0.00384  -0.0702   0.6560   1.0000
   4.000   0.7939   0.01006   0.00394  -0.0690   0.6291   1.0000
   4.250   0.8177   0.01023   0.00408  -0.0678   0.6034   1.0000
   4.500   0.8412   0.01037   0.00425  -0.0667   0.5710   1.0000
   4.750   0.8627   0.01056   0.00435  -0.0651   0.5088   1.0000
   5.000   0.8759   0.01159   0.00463  -0.0623   0.3705   1.0000
   5.250   0.8793   0.01441   0.00593  -0.0591   0.1116   1.0000
   5.500   0.8975   0.01559   0.00694  -0.0576   0.0806   1.0000
   5.750   0.9148   0.01687   0.00819  -0.0559   0.0643   1.0000
   6.000   0.9317   0.01823   0.00953  -0.0541   0.0517   1.0000
   6.250   0.9517   0.01929   0.01066  -0.0526   0.0447   1.0000
   6.500   0.9680   0.02155   0.01287  -0.0508   0.0400   1.0000
   6.750   0.9909   0.02235   0.01385  -0.0497   0.0361   1.0000
   7.000   1.0128   0.02371   0.01529  -0.0486   0.0330   1.0000
   7.250   1.0351   0.02561   0.01725  -0.0477   0.0311   1.0000
   7.500   1.0585   0.02896   0.02077  -0.0469   0.0299   1.0000
   7.750   1.0789   0.03336   0.02553  -0.0457   0.0292   1.0000
   8.000   1.0982   0.03403   0.02658  -0.0439   0.0277   1.0000
   8.250   1.1148   0.03725   0.03021  -0.0421   0.0274   1.0000
   8.500   1.1267   0.04154   0.03496  -0.0400   0.0279   1.0000
   8.750   1.1423   0.04610   0.03973  -0.0387   0.0301   1.0000
  10.750   0.9865   0.08617   0.08251  -0.0181   0.0524   1.0000
  11.000   0.9619   0.09112   0.08758  -0.0189   0.0524   1.0000
  11.250   0.9368   0.09675   0.09327  -0.0208   0.0524   1.0000
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