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GOE 598 AIRFOIL (goe598-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: GOE 598 AIRFOIL (goe598-il)
Reynolds number: 1,000,000
Max Cl/Cd: 60.84 at α=4°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe598-il-1000000-n5.txt
Download as CSV file: xf-goe598-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 598 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.6584   0.08946   0.08798   0.0085   1.0000   0.0018
  -8.750  -0.6635   0.08356   0.08210   0.0045   1.0000   0.0018
  -7.750  -0.7246   0.02034   0.01678  -0.0272   1.0000   0.0019
  -7.500  -0.7035   0.01791   0.01392  -0.0266   1.0000   0.0020
  -7.250  -0.6804   0.01630   0.01202  -0.0262   1.0000   0.0021
  -7.000  -0.6561   0.01512   0.01062  -0.0258   1.0000   0.0022
  -6.750  -0.6318   0.01395   0.00924  -0.0254   1.0000   0.0022
  -6.500  -0.6079   0.01265   0.00773  -0.0250   1.0000   0.0025
  -6.250  -0.5823   0.01215   0.00717  -0.0248   1.0000   0.0027
  -6.000  -0.5568   0.01164   0.00659  -0.0245   1.0000   0.0030
  -5.750  -0.5316   0.01098   0.00582  -0.0241   1.0000   0.0033
  -5.500  -0.5064   0.01038   0.00512  -0.0236   1.0000   0.0036
  -5.250  -0.4812   0.00993   0.00456  -0.0232   1.0000   0.0038
  -5.000  -0.4509   0.00929   0.00382  -0.0239   0.9975   0.0045
  -4.750  -0.4195   0.00891   0.00340  -0.0248   0.9945   0.0053
  -4.500  -0.3887   0.00860   0.00304  -0.0256   0.9903   0.0062
  -4.250  -0.3573   0.00822   0.00265  -0.0265   0.9859   0.0088
  -4.000  -0.3264   0.00795   0.00236  -0.0273   0.9797   0.0129
  -3.750  -0.2958   0.00776   0.00219  -0.0280   0.9722   0.0191
  -3.500  -0.2660   0.00756   0.00197  -0.0285   0.9633   0.0236
  -3.250  -0.2376   0.00742   0.00180  -0.0286   0.9522   0.0277
  -3.000  -0.2103   0.00733   0.00165  -0.0284   0.9384   0.0296
  -2.750  -0.1838   0.00721   0.00148  -0.0281   0.9203   0.0318
  -2.500  -0.1578   0.00713   0.00131  -0.0276   0.8969   0.0335
  -2.250  -0.1318   0.00709   0.00115  -0.0271   0.8698   0.0352
  -2.000  -0.1054   0.00707   0.00103  -0.0267   0.8420   0.0371
  -1.750  -0.0785   0.00702   0.00092  -0.0265   0.8177   0.0470
  -1.500  -0.0513   0.00695   0.00086  -0.0264   0.7968   0.0728
  -1.250  -0.0237   0.00695   0.00079  -0.0264   0.7781   0.0763
  -1.000   0.0039   0.00693   0.00074  -0.0263   0.7607   0.0822
  -0.750   0.0316   0.00692   0.00069  -0.0263   0.7426   0.0866
  -0.500   0.0591   0.00696   0.00065  -0.0263   0.7158   0.0911
  -0.250   0.0866   0.00697   0.00061  -0.0263   0.6867   0.1032
   0.000   0.1138   0.00671   0.00057  -0.0264   0.6646   0.2045
   0.250   0.1413   0.00658   0.00056  -0.0265   0.6429   0.2646
   0.500   0.1683   0.00618   0.00054  -0.0267   0.6239   0.4095
   0.750   0.1953   0.00597   0.00057  -0.0267   0.5989   0.5123
   1.000   0.2223   0.00577   0.00062  -0.0268   0.5732   0.6092
   1.250   0.2480   0.00545   0.00068  -0.0265   0.5420   0.7462
   1.500   0.2709   0.00547   0.00078  -0.0255   0.4515   0.8565
   1.750   0.2914   0.00573   0.00094  -0.0238   0.3548   0.9466
   2.000   0.3247   0.00614   0.00109  -0.0253   0.2740   0.9867
   2.250   0.3550   0.00675   0.00127  -0.0263   0.1592   1.0000
   2.500   0.3817   0.00707   0.00143  -0.0263   0.1146   1.0000
   2.750   0.4078   0.00754   0.00164  -0.0262   0.0491   1.0000
   3.000   0.4344   0.00790   0.00187  -0.0261   0.0167   1.0000
   3.250   0.4616   0.00812   0.00209  -0.0260   0.0108   1.0000
   3.500   0.4887   0.00841   0.00244  -0.0258   0.0078   1.0000
   3.750   0.5158   0.00863   0.00267  -0.0257   0.0065   1.0000
   4.000   0.5427   0.00892   0.00298  -0.0256   0.0055   1.0000
   4.250   0.5690   0.00941   0.00354  -0.0253   0.0046   1.0000
   4.500   0.5956   0.00980   0.00399  -0.0251   0.0042   1.0000
   4.750   0.6217   0.01032   0.00459  -0.0248   0.0038   1.0000
   5.000   0.6478   0.01076   0.00508  -0.0246   0.0034   1.0000
   5.250   0.6740   0.01114   0.00548  -0.0244   0.0031   1.0000
   5.500   0.6982   0.01210   0.00655  -0.0239   0.0027   1.0000
   5.750   0.7236   0.01276   0.00730  -0.0235   0.0026   1.0000
   6.000   0.7484   0.01357   0.00823  -0.0230   0.0025   1.0000
   6.250   0.7726   0.01464   0.00946  -0.0224   0.0023   1.0000
   6.500   0.7964   0.01597   0.01101  -0.0216   0.0020   1.0000
   6.750   0.8191   0.01772   0.01302  -0.0207   0.0018   1.0000
   7.000   0.8402   0.02021   0.01586  -0.0196   0.0017   1.0000
   7.250   0.8521   0.02677   0.02316  -0.0167   0.0017   1.0000
   7.500   0.8517   0.03772   0.03494  -0.0126   0.0019   1.0000
   7.750   0.8558   0.04476   0.04238  -0.0108   0.0020   1.0000
   8.000   0.8581   0.05126   0.04917  -0.0098   0.0021   1.0000
   8.250   0.8578   0.05741   0.05555  -0.0096   0.0021   1.0000
   8.500   0.8538   0.06344   0.06175  -0.0102   0.0022   1.0000
   8.750   0.8460   0.06953   0.06798  -0.0120   0.0022   1.0000
   9.000   0.8311   0.07538   0.07392  -0.0147   0.0023   1.0000
   9.250   0.8172   0.08457   0.08316  -0.0252   0.0023   1.0000
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