Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

GOE 385 AIRFOIL (goe385-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 385 AIRFOIL (goe385-il)
Reynolds number: 500,000
Max Cl/Cd: 97.44 at α=6.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe385-il-500000.txt
Download as CSV file: xf-goe385-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 385 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  -0.3658   0.08732   0.08525  -0.0315   1.0000   0.0204
  -8.000  -0.3632   0.08321   0.08116  -0.0357   1.0000   0.0204
  -7.750  -0.3670   0.07836   0.07635  -0.0371   1.0000   0.0206
  -7.500  -0.3681   0.07595   0.07397  -0.0356   0.9999   0.0208
  -7.250  -0.3459   0.07218   0.07019  -0.0397   0.9957   0.0212
  -7.000  -0.3208   0.06782   0.06579  -0.0459   0.9901   0.0217
  -6.750  -0.2929   0.06290   0.06081  -0.0534   0.9835   0.0225
  -6.500  -0.2661   0.05773   0.05554  -0.0603   0.9725   0.0239
  -6.250  -0.2292   0.05050   0.04792  -0.0690   0.9581   0.0256
  -6.000  -0.2170   0.04339   0.04054  -0.0712   0.9425   0.0261
  -5.750  -0.2025   0.04116   0.03829  -0.0709   0.9257   0.0266
  -5.500  -0.1850   0.03919   0.03623  -0.0706   0.9070   0.0272
  -5.250  -0.1655   0.03697   0.03386  -0.0705   0.8880   0.0281
  -5.000  -0.1436   0.03443   0.03106  -0.0704   0.8702   0.0298
  -4.750  -0.1123   0.03300   0.02903  -0.0696   0.8539   0.0326
  -4.500  -0.0953   0.02729   0.02303  -0.0702   0.8394   0.0340
  -4.250  -0.0720   0.02590   0.02151  -0.0701   0.8245   0.0351
  -4.000  -0.0471   0.02450   0.01992  -0.0700   0.8104   0.0367
  -3.750  -0.0181   0.02424   0.01932  -0.0694   0.7954   0.0411
  -3.500   0.0065   0.02094   0.01554  -0.0693   0.7804   0.0436
  -3.250   0.0321   0.01973   0.01429  -0.0695   0.7622   0.0455
  -3.000   0.0590   0.01882   0.01321  -0.0694   0.7430   0.0484
  -2.750   0.0887   0.02011   0.01423  -0.0687   0.7242   0.0535
  -2.500   0.1139   0.01667   0.01056  -0.0691   0.7092   0.0575
  -2.250   0.1414   0.01599   0.00975  -0.0691   0.6947   0.0613
  -2.000   0.1691   0.01528   0.00882  -0.0692   0.6817   0.0711
  -1.750   0.1997   0.01360   0.00681  -0.0683   0.6711   0.0500
  -1.500   0.2287   0.01259   0.00548  -0.0677   0.6611   0.0431
  -1.250   0.2568   0.01212   0.00493  -0.0676   0.6515   0.0428
  -1.000   0.2846   0.01152   0.00427  -0.0676   0.6430   0.0432
  -0.750   0.3125   0.01112   0.00381  -0.0675   0.6348   0.0431
  -0.500   0.3404   0.01070   0.00337  -0.0674   0.6272   0.0435
  -0.250   0.3683   0.01039   0.00301  -0.0674   0.6196   0.0442
   0.000   0.3965   0.01016   0.00275  -0.0675   0.6128   0.0452
   0.250   0.4248   0.01000   0.00256  -0.0675   0.6054   0.0463
   0.500   0.4531   0.00989   0.00244  -0.0676   0.5981   0.0485
   0.750   0.4815   0.00980   0.00232  -0.0677   0.5895   0.0507
   1.000   0.5099   0.00974   0.00221  -0.0678   0.5809   0.0529
   1.250   0.5381   0.00969   0.00211  -0.0678   0.5721   0.0591
   1.500   0.5535   0.00745   0.00226  -0.0657   0.5646   0.8964
   1.750   0.5915   0.00741   0.00218  -0.0677   0.5556   1.0000
   2.000   0.6196   0.00747   0.00219  -0.0677   0.5458   1.0000
   2.250   0.6476   0.00755   0.00221  -0.0677   0.5357   1.0000
   2.500   0.6754   0.00765   0.00224  -0.0677   0.5241   1.0000
   2.750   0.7031   0.00775   0.00228  -0.0677   0.5107   1.0000
   3.000   0.7308   0.00787   0.00233  -0.0677   0.4949   1.0000
   3.250   0.7582   0.00801   0.00241  -0.0677   0.4763   1.0000
   3.500   0.7853   0.00819   0.00250  -0.0677   0.4562   1.0000
   3.750   0.8122   0.00841   0.00262  -0.0676   0.4353   1.0000
   4.000   0.8388   0.00868   0.00277  -0.0675   0.4119   1.0000
   4.250   0.8652   0.00896   0.00294  -0.0674   0.3897   1.0000
   4.500   0.8916   0.00924   0.00313  -0.0673   0.3712   1.0000
   4.750   0.9182   0.00950   0.00333  -0.0672   0.3575   1.0000
   5.000   0.9447   0.00976   0.00355  -0.0671   0.3456   1.0000
   5.250   0.9711   0.01004   0.00377  -0.0670   0.3343   1.0000
   5.500   0.9977   0.01028   0.00400  -0.0669   0.3247   1.0000
   5.750   1.0242   0.01052   0.00424  -0.0669   0.3160   1.0000
   6.000   1.0501   0.01082   0.00449  -0.0667   0.3058   1.0000
   6.250   1.0762   0.01107   0.00473  -0.0666   0.2938   1.0000
   6.500   1.1023   0.01132   0.00497  -0.0665   0.2816   1.0000
   6.750   1.1283   0.01158   0.00523  -0.0664   0.2713   1.0000
   7.000   1.1538   0.01188   0.00551  -0.0662   0.2617   1.0000
   7.250   1.1794   0.01216   0.00578  -0.0660   0.2499   1.0000
   7.500   1.2048   0.01245   0.00608  -0.0658   0.2370   1.0000
   7.750   1.2292   0.01283   0.00641  -0.0655   0.2208   1.0000
   8.000   1.2532   0.01324   0.00677  -0.0652   0.1993   1.0000
   8.250   1.2729   0.01409   0.00734  -0.0643   0.1515   1.0000
   8.500   1.2897   0.01520   0.00821  -0.0631   0.1146   1.0000
   8.750   1.3075   0.01617   0.00901  -0.0619   0.0779   1.0000
   9.000   1.3198   0.01762   0.01020  -0.0600   0.0399   1.0000
   9.250   1.3342   0.01879   0.01121  -0.0583   0.0220   1.0000
   9.500   1.3517   0.01960   0.01209  -0.0570   0.0193   1.0000
   9.750   1.3668   0.02056   0.01313  -0.0553   0.0174   1.0000
  10.000   1.3797   0.02160   0.01427  -0.0534   0.0163   1.0000
  10.250   1.3923   0.02250   0.01529  -0.0514   0.0157   1.0000
  10.500   1.4004   0.02353   0.01642  -0.0488   0.0152   1.0000
  10.750   1.4068   0.02474   0.01773  -0.0463   0.0147   1.0000
  11.000   1.4117   0.02615   0.01924  -0.0439   0.0142   1.0000
  11.250   1.4145   0.02783   0.02103  -0.0417   0.0138   1.0000
  11.500   1.4145   0.02988   0.02319  -0.0397   0.0134   1.0000
  11.750   1.4107   0.03250   0.02592  -0.0381   0.0130   1.0000
  12.000   1.4020   0.03581   0.02936  -0.0367   0.0127   1.0000
  12.250   1.3975   0.03890   0.03257  -0.0358   0.0124   1.0000
  12.500   1.4016   0.04123   0.03501  -0.0355   0.0122   1.0000
  12.750   1.4026   0.04398   0.03788  -0.0352   0.0120   1.0000
  13.000   1.4016   0.04703   0.04104  -0.0351   0.0118   1.0000
  13.250   1.4000   0.05020   0.04432  -0.0350   0.0116   1.0000
  13.500   1.3975   0.05348   0.04771  -0.0350   0.0114   1.0000
  13.750   1.3948   0.05682   0.05116  -0.0350   0.0112   1.0000
  14.000   1.3919   0.06018   0.05462  -0.0351   0.0111   1.0000
  14.250   1.3889   0.06359   0.05813  -0.0351   0.0109   1.0000
  14.500   1.3864   0.06701   0.06164  -0.0353   0.0107   1.0000
  14.750   1.3839   0.07049   0.06521  -0.0356   0.0106   1.0000
  15.000   1.3814   0.07400   0.06882  -0.0359   0.0105   1.0000
  15.250   1.3790   0.07756   0.07247  -0.0362   0.0103   1.0000
  15.500   1.3768   0.08113   0.07613  -0.0367   0.0102   1.0000
  15.750   1.3744   0.08473   0.07981  -0.0371   0.0100   1.0000
  16.000   1.3721   0.08834   0.08352  -0.0376   0.0100   1.0000
  16.250   1.3701   0.09185   0.08711  -0.0379   0.0098   1.0000
  16.500   1.3684   0.09525   0.09059  -0.0380   0.0096   1.0000
  16.750   1.3658   0.09862   0.09406  -0.0375   0.0095   1.0000
  17.250   1.3466   0.10884   0.10461  -0.0402   0.0093   1.0000
  17.500   1.3367   0.11467   0.11061  -0.0433   0.0092   1.0000
  17.750   1.3268   0.12064   0.11674  -0.0466   0.0092   1.0000
  18.000   1.3164   0.12686   0.12311  -0.0501   0.0091   1.0000
  18.250   1.3062   0.13304   0.12945  -0.0534   0.0091   1.0000
  18.500   1.2951   0.13963   0.13619  -0.0572   0.0091   1.0000
  18.750   1.2838   0.14643   0.14314  -0.0613   0.0091   1.0000
<< Back to GOE 385 AIRFOIL (goe385-il)

Polar data table (+)

Polar graphs


<< Back to GOE 385 AIRFOIL (goe385-il)