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GOE 391 AIRFOIL (goe391-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 391 AIRFOIL (goe391-il)
Reynolds number: 500,000
Max Cl/Cd: 105.98 at α=1.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe391-il-500000.txt
Download as CSV file: xf-goe391-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 391 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
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.4429   0.09065   0.08845  -0.0213   1.0000   0.0114
  -7.500  -0.4438   0.08768   0.08551  -0.0220   1.0000   0.0114
  -7.250  -0.4517   0.08293   0.08081  -0.0202   1.0000   0.0118
  -7.000  -0.4521   0.08029   0.07818  -0.0195   1.0000   0.0120
  -6.750  -0.4523   0.07781   0.07571  -0.0187   1.0000   0.0123
  -6.500  -0.4515   0.07519   0.07311  -0.0187   1.0000   0.0125
  -6.250  -0.4490   0.07246   0.07038  -0.0189   1.0000   0.0128
  -6.000  -0.4449   0.06972   0.06763  -0.0192   1.0000   0.0133
  -5.750  -0.4202   0.06551   0.06336  -0.0245   0.9982   0.0147
  -5.500  -0.3735   0.06092   0.05862  -0.0339   0.9956   0.0166
  -5.250  -0.3387   0.05669   0.05426  -0.0393   0.9934   0.0168
  -5.000  -0.3081   0.05244   0.04986  -0.0431   0.9903   0.0169
  -4.750  -0.2753   0.04804   0.04529  -0.0470   0.9874   0.0170
  -4.500  -0.2405   0.04361   0.04067  -0.0508   0.9851   0.0170
  -4.250  -0.2126   0.03619   0.03304  -0.0555   0.9830   0.0180
  -4.000  -0.1872   0.03352   0.03025  -0.0566   0.9781   0.0186
  -3.750  -0.1557   0.03085   0.02742  -0.0586   0.9749   0.0200
  -3.500  -0.1152   0.02971   0.02602  -0.0598   0.9729   0.0249
  -2.250   0.0393   0.01471   0.00929  -0.0632   0.9595   0.0278
  -2.000   0.0758   0.01415   0.00870  -0.0653   0.9580   0.0326
  -1.750   0.1122   0.01236   0.00665  -0.0666   0.9569   0.0326
  -1.500   0.1446   0.01146   0.00561  -0.0673   0.9543   0.0344
  -1.250   0.1736   0.01102   0.00509  -0.0673   0.9495   0.0376
  -1.000   0.2061   0.00990   0.00393  -0.0682   0.9467   0.0392
  -0.750   0.2412   0.00924   0.00327  -0.0697   0.9438   0.0413
  -0.500   0.2694   0.00884   0.00288  -0.0696   0.9378   0.0439
  -0.250   0.2993   0.00844   0.00248  -0.0699   0.9324   0.0459
   0.000   0.3294   0.00812   0.00214  -0.0701   0.9258   0.0482
   0.250   0.3579   0.00789   0.00189  -0.0700   0.9177   0.0498
   0.500   0.3849   0.00768   0.00168  -0.0696   0.9085   0.0531
   1.000   0.4894   0.00516   0.00150  -0.0806   0.8796   1.0000
   1.250   0.5149   0.00516   0.00142  -0.0798   0.8618   1.0000
   1.500   0.5389   0.00520   0.00136  -0.0787   0.8343   1.0000
   1.750   0.5617   0.00530   0.00131  -0.0773   0.7965   1.0000
   2.000   0.5813   0.00554   0.00127  -0.0752   0.7297   1.0000
   2.250   0.5974   0.00600   0.00132  -0.0725   0.6429   1.0000
   2.750   0.6320   0.00703   0.00165  -0.0681   0.4835   1.0000
   3.000   0.6469   0.00787   0.00190  -0.0657   0.3532   1.0000
   3.250   0.6664   0.00843   0.00216  -0.0642   0.2870   1.0000
   3.500   0.6835   0.00928   0.00248  -0.0623   0.1694   1.0000
   3.750   0.7005   0.01022   0.00292  -0.0604   0.0651   1.0000
   4.000   0.7227   0.01065   0.00327  -0.0593   0.0488   1.0000
   4.250   0.7453   0.01105   0.00370  -0.0582   0.0434   1.0000
   4.500   0.7675   0.01152   0.00423  -0.0570   0.0410   1.0000
   4.750   0.7903   0.01189   0.00466  -0.0560   0.0399   1.0000
   5.000   0.8128   0.01230   0.00513  -0.0550   0.0379   1.0000
   5.250   0.8346   0.01282   0.00571  -0.0538   0.0356   1.0000
   5.500   0.8550   0.01351   0.00646  -0.0525   0.0319   1.0000
   5.750   0.8723   0.01473   0.00778  -0.0505   0.0285   1.0000
   6.000   0.8954   0.01515   0.00827  -0.0496   0.0261   1.0000
   6.250   0.9173   0.01576   0.00894  -0.0485   0.0231   1.0000
   6.500   0.9362   0.01696   0.01015  -0.0471   0.0196   1.0000
   6.750   0.9568   0.01822   0.01157  -0.0458   0.0181   1.0000
   7.000   0.9792   0.01904   0.01250  -0.0447   0.0166   1.0000
   7.250   1.0010   0.02006   0.01364  -0.0436   0.0152   1.0000
   7.500   1.0224   0.02086   0.01453  -0.0425   0.0140   1.0000
   7.750   1.0418   0.02217   0.01592  -0.0414   0.0128   1.0000
   8.000   1.0578   0.02530   0.01942  -0.0395   0.0119   1.0000
   8.250   1.0753   0.02733   0.02173  -0.0377   0.0115   1.0000
   8.500   1.0894   0.03007   0.02487  -0.0354   0.0108   1.0000
   8.750   1.0971   0.03390   0.02916  -0.0323   0.0106   1.0000
   9.000   1.0974   0.03856   0.03430  -0.0285   0.0106   1.0000
   9.250   1.0910   0.04368   0.03987  -0.0243   0.0108   1.0000
   9.500   1.0799   0.04848   0.04503  -0.0203   0.0111   1.0000
   9.750   1.0638   0.05303   0.04985  -0.0165   0.0115   1.0000
  10.000   0.9853   0.04663   0.04377  -0.0077   0.0115   1.0000
  10.250   0.9609   0.05061   0.04791  -0.0052   0.0116   1.0000
  10.500   0.9355   0.05534   0.05279  -0.0044   0.0116   1.0000
  10.750   0.9116   0.06071   0.05829  -0.0053   0.0117   1.0000
  11.000   0.8873   0.06730   0.06498  -0.0082   0.0117   1.0000
  11.250   0.8637   0.07524   0.07303  -0.0128   0.0117   1.0000
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