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

GOE 5K AIRFOIL (goe05k-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: GOE 5K AIRFOIL (goe05k-il)
Reynolds number: 500,000
Max Cl/Cd: 82.41 at α=1.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe05k-il-500000-n5.txt
Download as CSV file: xf-goe05k-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 5K 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.5227   0.08978   0.08755  -0.0103   1.0000   0.0032
  -7.500  -0.5248   0.08705   0.08484  -0.0106   1.0000   0.0032
  -7.250  -0.5256   0.08394   0.08177  -0.0116   1.0000   0.0032
  -7.000  -0.5221   0.08052   0.07837  -0.0137   1.0000   0.0033
  -6.750  -0.5169   0.07690   0.07475  -0.0161   1.0000   0.0034
  -6.500  -0.5098   0.07321   0.07106  -0.0185   1.0000   0.0034
  -6.250  -0.5010   0.06938   0.06721  -0.0209   1.0000   0.0035
  -6.000  -0.4905   0.06546   0.06326  -0.0231   1.0000   0.0035
  -5.750  -0.4785   0.06157   0.05932  -0.0250   1.0000   0.0036
  -5.500  -0.4649   0.05766   0.05535  -0.0267   1.0000   0.0037
  -5.250  -0.4499   0.05378   0.05138  -0.0280   1.0000   0.0039
  -5.000  -0.4243   0.04924   0.04671  -0.0312   0.9990   0.0041
  -4.750  -0.3915   0.04452   0.04181  -0.0350   0.9971   0.0043
  -4.500  -0.3557   0.03988   0.03691  -0.0380   0.9954   0.0045
  -4.250  -0.3241   0.03583   0.03262  -0.0396   0.9932   0.0046
  -3.750  -0.2674   0.02834   0.02465  -0.0430   0.9890   0.0050
  -2.750  -0.1463   0.01676   0.01171  -0.0433   0.9806   0.0049
  -2.500  -0.1150   0.01425   0.00885  -0.0431   0.9794   0.0028
  -2.250  -0.0867   0.01231   0.00659  -0.0425   0.9774   0.0026
  -2.000  -0.0594   0.01079   0.00479  -0.0417   0.9747   0.0030
  -1.750  -0.0313   0.00977   0.00361  -0.0417   0.9724   0.0043
  -1.500  -0.0020   0.00887   0.00256  -0.0419   0.9705   0.0048
  -1.000   0.0556   0.00792   0.00151  -0.0423   0.9653   0.0338
  -0.750   0.1017   0.00523   0.00167  -0.0475   0.9717   0.9915
  -0.500   0.1362   0.00525   0.00158  -0.0492   0.9705   1.0000
   0.000   0.1967   0.00521   0.00143  -0.0506   0.9638   1.0000
   0.250   0.2283   0.00517   0.00138  -0.0516   0.9598   1.0000
   0.500   0.2671   0.00508   0.00130  -0.0541   0.9558   1.0000
   0.750   0.3036   0.00493   0.00117  -0.0560   0.9421   1.0000
   1.000   0.3439   0.00479   0.00112  -0.0587   0.9239   1.0000
   1.250   0.3906   0.00474   0.00098  -0.0625   0.8556   1.0000
   1.500   0.3842   0.00692   0.00114  -0.0549   0.3705   1.0000
   1.750   0.3939   0.00880   0.00168  -0.0518   0.0098   1.0000
   2.000   0.4169   0.00925   0.00224  -0.0506   0.0054   1.0000
   2.500   0.4601   0.01064   0.00383  -0.0477   0.0031   1.0000
   2.750   0.4808   0.01172   0.00504  -0.0459   0.0031   1.0000
   3.000   0.5028   0.01315   0.00663  -0.0442   0.0032   1.0000
   7.000   0.7804   0.06646   0.06400  -0.0178   0.0028   1.0000
   7.500   0.7141   0.06446   0.06240  -0.0158   0.0030   1.0000
   8.000   0.7814   0.08422   0.08204  -0.0215   0.0027   1.0000
   8.250   0.7728   0.08848   0.08631  -0.0237   0.0026   1.0000
   8.500   0.7668   0.09320   0.09101  -0.0277   0.0026   1.0000
<< Back to GOE 5K AIRFOIL (goe05k-il)

Polar data table (+)

Polar graphs


<< Back to GOE 5K AIRFOIL (goe05k-il)