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GOE 504 AIRFOIL (goe504-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 504 AIRFOIL (goe504-il)
Reynolds number: 50,000
Max Cl/Cd: 13.69 at α=2.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe504-il-50000-n5.txt
Download as CSV file: xf-goe504-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 504 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
  ------ -------- --------- --------- -------- -------- --------
 -10.750  -0.2434   0.11149   0.10525  -0.0700   0.9663   0.0643
 -10.500  -0.2394   0.10566   0.09943  -0.0748   0.9607   0.0640
 -10.250  -0.2432   0.09976   0.09356  -0.0789   0.9523   0.0641
 -10.000  -0.2490   0.09234   0.08615  -0.0850   0.9461   0.0642
  -9.750  -0.2685   0.08440   0.07824  -0.0903   0.9358   0.0637
  -9.500  -0.3016   0.07471   0.06845  -0.0986   0.9252   0.0626
  -9.250  -0.3425   0.06793   0.06148  -0.1025   0.9126   0.0612
  -9.000  -0.3665   0.06382   0.05713  -0.1020   0.9000   0.0613
  -8.750  -0.3830   0.06049   0.05354  -0.1003   0.8879   0.0615
  -8.500  -0.3888   0.05729   0.05001  -0.0992   0.8785   0.0623
  -8.250  -0.3876   0.05424   0.04657  -0.0983   0.8701   0.0635
  -8.000  -0.3884   0.05184   0.04380  -0.0960   0.8603   0.0644
  -7.750  -0.3745   0.04899   0.04046  -0.0957   0.8541   0.0654
  -7.500  -0.3696   0.04722   0.03830  -0.0931   0.8446   0.0666
  -7.250  -0.3469   0.04535   0.03626  -0.0933   0.8391   0.0692
  -7.000  -0.3353   0.04423   0.03501  -0.0914   0.8307   0.0714
  -6.750  -0.3124   0.04267   0.03318  -0.0911   0.8246   0.0739
  -6.500  -0.2911   0.04130   0.03151  -0.0903   0.8182   0.0760
  -6.250  -0.2739   0.04029   0.03020  -0.0887   0.8105   0.0784
  -6.000  -0.2434   0.03889   0.02879  -0.0895   0.8065   0.0829
  -5.750  -0.2361   0.03841   0.02821  -0.0865   0.7965   0.0864
  -5.500  -0.2065   0.03730   0.02683  -0.0867   0.7919   0.0917
  -5.250  -0.1949   0.03668   0.02619  -0.0843   0.7836   0.0957
  -5.000  -0.1720   0.03578   0.02517  -0.0836   0.7775   0.1036
  -4.750  -0.1393   0.03445   0.02384  -0.0846   0.7740   0.1216
  -4.500  -0.1366   0.03417   0.02379  -0.0811   0.7636   0.1399
  -4.250  -0.1068   0.03338   0.02340  -0.0816   0.7592   0.2156
  -4.000  -0.0969   0.03370   0.02359  -0.0788   0.7500   0.2558
  -3.750  -0.0721   0.03369   0.02348  -0.0783   0.7445   0.2992
  -3.500  -0.0488   0.03370   0.02338  -0.0774   0.7389   0.3317
  -3.250  -0.0377   0.03401   0.02364  -0.0749   0.7300   0.3555
  -3.000  -0.0061   0.03374   0.02331  -0.0752   0.7262   0.3895
  -2.750  -0.0021   0.03429   0.02384  -0.0717   0.7163   0.4094
  -2.500   0.0271   0.03404   0.02355  -0.0717   0.7117   0.4359
  -2.250   0.0468   0.03415   0.02357  -0.0705   0.7048   0.4530
  -2.000   0.0674   0.03427   0.02358  -0.0696   0.6977   0.4667
  -1.750   0.1050   0.03382   0.02301  -0.0709   0.6942   0.4837
  -1.500   0.1117   0.03447   0.02362  -0.0682   0.6844   0.4967
  -1.250   0.1431   0.03418   0.02331  -0.0686   0.6798   0.5173
  -1.000   0.1815   0.03360   0.02273  -0.0699   0.6770   0.5470
  -0.750   0.1796   0.03456   0.02382  -0.0660   0.6659   0.5701
  -0.500   0.2135   0.03398   0.02344  -0.0663   0.6624   0.6248
  -0.250   0.2203   0.03468   0.02446  -0.0633   0.6531   0.6933
   0.000   0.3300   0.03421   0.02428  -0.0771   0.6503   0.9253
   0.250   0.4086   0.03404   0.02375  -0.0864   0.6474   1.0000
   0.750   0.4199   0.03528   0.02461  -0.0802   0.6322   1.0000
   1.000   0.4579   0.03493   0.02401  -0.0815   0.6292   1.0000
   1.250   0.4361   0.03667   0.02569  -0.0750   0.6172   1.0000
   1.500   0.4713   0.03636   0.02516  -0.0757   0.6140   1.0000
   1.750   0.4558   0.03819   0.02692  -0.0705   0.6023   1.0000
   2.000   0.4878   0.03799   0.02655  -0.0707   0.5988   1.0000
   2.500   0.5066   0.03976   0.02810  -0.0664   0.5837   1.0000
   2.750   0.5402   0.03947   0.02766  -0.0668   0.5806   1.0000
   3.000   0.5260   0.04176   0.02992  -0.0626   0.5686   1.0000
   3.250   0.5618   0.04127   0.02929  -0.0630   0.5659   1.0000
   3.750   0.5797   0.04350   0.03139  -0.0593   0.5507   1.0000
   4.250   0.5984   0.04586   0.03365  -0.0560   0.5353   1.0000
   4.750   0.6161   0.04852   0.03622  -0.0529   0.5200   1.0000
   5.000   0.6513   0.04786   0.03549  -0.0530   0.5178   1.0000
   5.500   0.6279   0.05415   0.04179  -0.0483   0.4942   1.0000
   5.750   0.6504   0.05445   0.04205  -0.0477   0.4894   1.0000
   6.000   0.6819   0.05394   0.04150  -0.0474   0.4868   1.0000
   6.500   0.6954   0.05743   0.04497  -0.0450   0.4714   1.0000
   7.000   0.7079   0.06117   0.04872  -0.0429   0.4560   1.0000
   7.500   0.7183   0.06529   0.05287  -0.0410   0.4408   1.0000
   7.750   0.7488   0.06478   0.05236  -0.0406   0.4387   1.0000
   8.000   0.7265   0.06982   0.05744  -0.0393   0.4259   1.0000
   8.250   0.7552   0.06948   0.05711  -0.0389   0.4235   1.0000
   8.750   0.7593   0.07465   0.06234  -0.0375   0.4085   1.0000
   9.250   0.7607   0.08034   0.06812  -0.0364   0.3938   1.0000
   9.750   0.7600   0.08646   0.07434  -0.0357   0.3795   1.0000
  10.000   0.7844   0.08670   0.07461  -0.0353   0.3772   1.0000
  10.250   0.7584   0.09301   0.08098  -0.0353   0.3663   1.0000
  10.500   0.7782   0.09376   0.08179  -0.0350   0.3631   1.0000
  10.750   0.8042   0.09375   0.08182  -0.0345   0.3610   1.0000
  11.000   0.7718   0.10115   0.08929  -0.0351   0.3500   1.0000
  11.250   0.7911   0.10205   0.09025  -0.0348   0.3472   1.0000
  11.750   0.7831   0.10980   0.09813  -0.0354   0.3345   1.0000
  12.000   0.8016   0.11083   0.09924  -0.0352   0.3318   1.0000
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