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

GOE 415 AIRFOIL (goe415-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: GOE 415 AIRFOIL (goe415-il)
Reynolds number: 1,000,000
Max Cl/Cd: 106.85 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe415-il-1000000-n5.txt
Download as CSV file: xf-goe415-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 415 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.3915   0.09544   0.09396  -0.0063   1.0000   0.0046
  -8.000  -0.3861   0.09188   0.09042  -0.0083   1.0000   0.0046
  -7.750  -0.3810   0.08830   0.08687  -0.0103   1.0000   0.0046
  -7.500  -0.3731   0.08449   0.08307  -0.0134   0.9816   0.0046
  -7.250  -0.3594   0.08004   0.07856  -0.0187   0.9379   0.0046
  -6.250  -0.2998   0.05754   0.05574  -0.0440   0.8710   0.0030
  -6.000  -0.2760   0.05169   0.04978  -0.0505   0.8557   0.0030
  -5.750  -0.2512   0.04645   0.04437  -0.0557   0.8357   0.0031
  -5.500  -0.2265   0.04183   0.03950  -0.0594   0.8025   0.0032
  -5.250  -0.2018   0.03729   0.03460  -0.0621   0.7540   0.0033
  -5.000  -0.1765   0.03290   0.02983  -0.0640   0.7156   0.0034
  -4.750  -0.1505   0.02847   0.02500  -0.0652   0.6876   0.0035
  -4.500  -0.1247   0.02520   0.02136  -0.0655   0.6584   0.0038
  -4.250  -0.0984   0.02161   0.01731  -0.0655   0.6295   0.0040
  -4.000  -0.0719   0.01718   0.01224  -0.0648   0.6018   0.0038
  -3.750  -0.0464   0.01172   0.00590  -0.0635   0.5790   0.0041
  -3.500  -0.0198   0.01086   0.00475  -0.0631   0.5537   0.0044
  -3.250   0.0073   0.01035   0.00403  -0.0629   0.5335   0.0047
  -3.000   0.0343   0.00975   0.00325  -0.0626   0.5181   0.0051
  -2.750   0.0619   0.00951   0.00293  -0.0625   0.5059   0.0055
  -2.500   0.0897   0.00936   0.00271  -0.0624   0.4949   0.0062
  -2.250   0.1175   0.00919   0.00246  -0.0623   0.4844   0.0071
  -2.000   0.1454   0.00892   0.00209  -0.0622   0.4758   0.0078
  -1.750   0.1733   0.00872   0.00182  -0.0621   0.4669   0.0094
  -1.500   0.2014   0.00863   0.00170  -0.0621   0.4578   0.0111
  -1.250   0.2294   0.00851   0.00150  -0.0620   0.4481   0.0138
  -1.000   0.2573   0.00845   0.00139  -0.0619   0.4371   0.0188
  -0.750   0.2853   0.00845   0.00134  -0.0619   0.4254   0.0221
  -0.500   0.3133   0.00842   0.00124  -0.0619   0.4146   0.0243
  -0.250   0.3412   0.00838   0.00115  -0.0618   0.4038   0.0296
   0.000   0.3691   0.00839   0.00109  -0.0618   0.3925   0.0345
   0.250   0.3964   0.00812   0.00104  -0.0618   0.3792   0.1503
   0.500   0.4233   0.00785   0.00107  -0.0618   0.3635   0.2929
   0.750   0.4507   0.00778   0.00109  -0.0618   0.3495   0.3552
   1.000   0.4780   0.00770   0.00115  -0.0618   0.3382   0.4319
   1.250   0.5054   0.00763   0.00119  -0.0617   0.3276   0.4989
   1.500   0.5282   0.00692   0.00130  -0.0609   0.3184   0.8032
   2.000   0.5912   0.00676   0.00140  -0.0624   0.2997   1.0000
   2.250   0.6186   0.00690   0.00147  -0.0623   0.2895   1.0000
   2.500   0.6459   0.00705   0.00155  -0.0622   0.2790   1.0000
   2.750   0.6734   0.00717   0.00162  -0.0621   0.2718   1.0000
   3.000   0.7008   0.00730   0.00172  -0.0620   0.2653   1.0000
   3.250   0.7283   0.00741   0.00181  -0.0620   0.2597   1.0000
   3.500   0.7556   0.00755   0.00191  -0.0619   0.2535   1.0000
   3.750   0.7830   0.00768   0.00203  -0.0618   0.2482   1.0000
   4.000   0.8102   0.00783   0.00214  -0.0617   0.2404   1.0000
   4.250   0.8375   0.00798   0.00227  -0.0617   0.2344   1.0000
   4.500   0.8647   0.00812   0.00241  -0.0616   0.2286   1.0000
   4.750   0.8913   0.00835   0.00258  -0.0615   0.2143   1.0000
   5.000   0.9178   0.00859   0.00276  -0.0614   0.1994   1.0000
   5.250   0.9437   0.00894   0.00299  -0.0612   0.1750   1.0000
   5.500   0.9668   0.00972   0.00346  -0.0607   0.1202   1.0000
   5.750   0.9915   0.01023   0.00386  -0.0604   0.0968   1.0000
   6.000   1.0112   0.01159   0.00480  -0.0594   0.0138   1.0000
   6.250   1.0370   0.01192   0.00516  -0.0592   0.0087   1.0000
   6.500   1.0627   0.01226   0.00554  -0.0589   0.0070   1.0000
   6.750   1.0880   0.01265   0.00597  -0.0586   0.0056   1.0000
   7.000   1.1132   0.01300   0.00636  -0.0584   0.0047   1.0000
   7.250   1.1377   0.01349   0.00689  -0.0579   0.0039   1.0000
   7.500   1.1623   0.01393   0.00738  -0.0576   0.0035   1.0000
   7.750   1.1866   0.01440   0.00791  -0.0571   0.0032   1.0000
   8.000   1.2105   0.01491   0.00850  -0.0567   0.0028   1.0000
   8.250   1.2337   0.01548   0.00913  -0.0562   0.0026   1.0000
   8.500   1.2551   0.01630   0.01004  -0.0554   0.0023   1.0000
   8.750   1.2780   0.01685   0.01066  -0.0549   0.0021   1.0000
   9.000   1.2998   0.01753   0.01142  -0.0542   0.0019   1.0000
   9.250   1.3204   0.01831   0.01229  -0.0534   0.0018   1.0000
   9.500   1.3403   0.01911   0.01318  -0.0525   0.0017   1.0000
   9.750   1.3595   0.01994   0.01409  -0.0515   0.0016   1.0000
  10.000   1.3780   0.02079   0.01502  -0.0506   0.0015   1.0000
  10.250   1.3947   0.02176   0.01608  -0.0494   0.0014   1.0000
  10.500   1.4058   0.02319   0.01764  -0.0475   0.0013   1.0000
  10.750   1.4134   0.02478   0.01937  -0.0453   0.0013   1.0000
  11.000   1.4212   0.02611   0.02086  -0.0431   0.0012   1.0000
  11.250   1.4235   0.02759   0.02247  -0.0403   0.0012   1.0000
  11.500   1.4226   0.02945   0.02446  -0.0378   0.0011   1.0000
  11.750   1.4200   0.03170   0.02686  -0.0359   0.0011   1.0000
  12.000   1.4156   0.03446   0.02977  -0.0347   0.0011   1.0000
  12.250   1.4099   0.03775   0.03321  -0.0342   0.0011   1.0000
  12.500   1.4031   0.04162   0.03723  -0.0347   0.0010   1.0000
  12.750   1.3956   0.04592   0.04168  -0.0356   0.0010   1.0000
  13.000   1.3867   0.05056   0.04648  -0.0368   0.0010   1.0000
  13.250   1.3766   0.05547   0.05153  -0.0381   0.0010   1.0000
  13.500   1.3654   0.06052   0.05674  -0.0394   0.0010   1.0000
  13.750   1.3530   0.06573   0.06209  -0.0407   0.0009   1.0000
  14.000   1.3389   0.07118   0.06769  -0.0419   0.0009   1.0000
  14.250   1.3243   0.07678   0.07344  -0.0433   0.0009   1.0000
  14.500   1.3086   0.08273   0.07954  -0.0448   0.0009   1.0000
  14.750   1.2934   0.08888   0.08584  -0.0466   0.0009   1.0000
  15.000   1.2789   0.09516   0.09227  -0.0488   0.0009   1.0000
<< Back to GOE 415 AIRFOIL (goe415-il)

Polar data table (+)

Polar graphs


<< Back to GOE 415 AIRFOIL (goe415-il)