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GOE 5K AIRFOIL (goe05k-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: GOE 5K AIRFOIL (goe05k-il)
Reynolds number: 1,000,000
Max Cl/Cd: 95.14 at α=0.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe05k-il-1000000.txt
Download as CSV file: xf-goe05k-il-1000000.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 =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4275   0.08340   0.08186  -0.0143   1.0000   0.0026
  -8.250  -0.4278   0.07976   0.07823  -0.0148   1.0000   0.0026
  -8.000  -0.4287   0.07622   0.07471  -0.0152   1.0000   0.0027
  -7.750  -0.4302   0.07270   0.07120  -0.0156   1.0000   0.0027
  -7.500  -0.4328   0.06933   0.06785  -0.0158   1.0000   0.0027
  -7.250  -0.4377   0.06623   0.06478  -0.0157   1.0000   0.0027
  -7.000  -0.4469   0.06361   0.06218  -0.0147   1.0000   0.0028
  -6.750  -0.4522   0.06047   0.05906  -0.0154   1.0000   0.0028
  -1.750  -0.0267   0.00881   0.00377  -0.0430   0.9803   0.0052
  -1.500   0.0056   0.00815   0.00300  -0.0438   0.9794   0.0067
  -1.250   0.0324   0.00728   0.00202  -0.0434   0.9766   0.0068
  -1.000   0.0611   0.00678   0.00144  -0.0435   0.9738   0.0158
  -0.750   0.1114   0.00418   0.00168  -0.0499   0.9816   1.0000
  -0.500   0.1471   0.00415   0.00157  -0.0519   0.9802   1.0000
  -0.250   0.1850   0.00410   0.00146  -0.0543   0.9790   1.0000
   0.000   0.2189   0.00395   0.00129  -0.0557   0.9731   1.0000
   0.250   0.2612   0.00379   0.00113  -0.0590   0.9697   1.0000
   0.500   0.2956   0.00364   0.00099  -0.0605   0.9609   1.0000
   0.750   0.3330   0.00350   0.00088  -0.0626   0.9482   1.0000
   1.250   0.3884   0.00421   0.00067  -0.0619   0.6811   1.0000
   1.500   0.3847   0.00664   0.00109  -0.0560   0.1342   1.0000
   1.750   0.4028   0.00761   0.00169  -0.0538   0.0073   1.0000
   2.000   0.4247   0.00818   0.00235  -0.0523   0.0047   1.0000
   2.250   0.4433   0.00926   0.00351  -0.0501   0.0051   1.0000
   4.500   0.6158   0.01130   0.00783  -0.0295   0.0032   1.0000
   4.750   0.6354   0.01398   0.01078  -0.0270   0.0029   1.0000
   5.000   0.6521   0.01716   0.01421  -0.0248   0.0026   1.0000
   5.250   0.6673   0.02066   0.01793  -0.0229   0.0024   1.0000
   5.500   0.6809   0.02439   0.02186  -0.0212   0.0023   1.0000
   5.750   0.6910   0.02848   0.02611  -0.0199   0.0021   1.0000
   6.000   0.6960   0.03376   0.03156  -0.0186   0.0021   1.0000
   6.250   0.7054   0.03849   0.03646  -0.0172   0.0021   1.0000
   6.500   0.7130   0.04333   0.04143  -0.0162   0.0021   1.0000
   6.750   0.7181   0.04838   0.04661  -0.0154   0.0020   1.0000
   7.000   0.7204   0.05337   0.05171  -0.0150   0.0020   1.0000
   7.250   0.7195   0.05829   0.05671  -0.0148   0.0020   1.0000
   7.500   0.7150   0.06322   0.06171  -0.0151   0.0020   1.0000
   7.750   0.7036   0.06724   0.06577  -0.0147   0.0020   1.0000
   8.000   0.6882   0.07145   0.07001  -0.0158   0.0020   1.0000
   8.250   0.6742   0.07667   0.07523  -0.0201   0.0020   1.0000
   8.500   0.6668   0.08189   0.08044  -0.0230   0.0021   1.0000
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