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AG12 (ag12-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: AG12 (ag12-il)
Reynolds number: 500,000
Max Cl/Cd: 79.62 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag12-il-500000.txt
Download as CSV file: xf-ag12-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG12                                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
   0.250   0.2509   0.00481   0.00107  -0.0467   0.8066   1.0000
   0.500   0.2771   0.00498   0.00105  -0.0460   0.7672   1.0000
   0.750   0.3036   0.00517   0.00104  -0.0455   0.7285   1.0000
   1.000   0.3303   0.00537   0.00105  -0.0451   0.6901   1.0000
   1.250   0.3572   0.00557   0.00108  -0.0447   0.6523   1.0000
   1.500   0.3843   0.00577   0.00112  -0.0444   0.6151   1.0000
   1.750   0.4115   0.00598   0.00118  -0.0441   0.5789   1.0000
   2.000   0.4387   0.00620   0.00125  -0.0439   0.5434   1.0000
   2.250   0.4660   0.00642   0.00133  -0.0437   0.5089   1.0000
   2.500   0.4932   0.00665   0.00143  -0.0436   0.4750   1.0000
   2.750   0.5205   0.00688   0.00153  -0.0434   0.4410   1.0000
   3.000   0.5477   0.00714   0.00168  -0.0432   0.4076   1.0000
   3.250   0.5749   0.00739   0.00182  -0.0431   0.3753   1.0000
   3.500   0.6020   0.00767   0.00198  -0.0430   0.3423   1.0000
   3.750   0.6291   0.00795   0.00216  -0.0428   0.3111   1.0000
   4.000   0.6561   0.00826   0.00237  -0.0427   0.2786   1.0000
   4.250   0.6831   0.00858   0.00259  -0.0426   0.2484   1.0000
   4.500   0.7099   0.00894   0.00284  -0.0424   0.2160   1.0000
   4.750   0.7366   0.00931   0.00311  -0.0423   0.1877   1.0000
   5.000   0.7632   0.00971   0.00340  -0.0421   0.1587   1.0000
   5.250   0.7896   0.01013   0.00376  -0.0420   0.1323   1.0000
   5.500   0.8159   0.01059   0.00412  -0.0418   0.1069   1.0000
   5.750   0.8422   0.01104   0.00452  -0.0416   0.0859   1.0000
   6.000   0.8682   0.01156   0.00497  -0.0413   0.0669   1.0000
   6.250   0.8941   0.01209   0.00546  -0.0411   0.0510   1.0000
   6.500   0.9198   0.01269   0.00601  -0.0408   0.0374   1.0000
   6.750   0.9452   0.01335   0.00672  -0.0405   0.0273   1.0000
   7.000   0.9703   0.01410   0.00751  -0.0400   0.0196   1.0000
   7.250   0.9934   0.01541   0.00893  -0.0393   0.0140   1.0000
   7.500   1.0182   0.01620   0.00983  -0.0388   0.0119   1.0000
   7.750   1.0418   0.01721   0.01095  -0.0382   0.0102   1.0000
   8.000   1.0604   0.01955   0.01352  -0.0370   0.0089   1.0000
   8.250   1.0768   0.02265   0.01695  -0.0355   0.0085   1.0000
   8.500   1.0991   0.02388   0.01837  -0.0348   0.0082   1.0000
   8.750   1.1210   0.02504   0.01976  -0.0341   0.0077   1.0000
   9.000   1.1412   0.02657   0.02149  -0.0333   0.0071   1.0000
   9.250   1.1575   0.02900   0.02422  -0.0322   0.0069   1.0000
   9.500   1.1716   0.03169   0.02723  -0.0311   0.0067   1.0000
   9.750   1.1826   0.03467   0.03054  -0.0299   0.0065   1.0000
  10.000   1.1899   0.03790   0.03410  -0.0287   0.0063   1.0000
  10.250   1.1927   0.04138   0.03790  -0.0275   0.0061   1.0000
  10.500   1.1882   0.04537   0.04219  -0.0264   0.0060   1.0000
  10.750   1.1739   0.04948   0.04658  -0.0252   0.0060   1.0000
  11.000   1.1506   0.05529   0.05267  -0.0273   0.0061   1.0000
  11.250   1.1258   0.06410   0.06174  -0.0348   0.0063   1.0000
  11.500   1.1008   0.07554   0.07339  -0.0450   0.0065   1.0000
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