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S2062 8% (s2062-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: S2062 8% (s2062-il)
Reynolds number: 500,000
Max Cl/Cd: 89.26 at α=3°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-s2062-il-500000.txt
Download as CSV file: xf-s2062-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: S2062 8%                                        
  
 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   0.2560   0.00524   0.00164  -0.0536   0.8604   0.9464
   0.500   0.2885   0.00523   0.00159  -0.0543   0.8471   0.9660
   0.750   0.3257   0.00523   0.00155  -0.0561   0.8329   0.9815
   1.000   0.3655   0.00524   0.00150  -0.0586   0.8171   0.9929
   1.250   0.3993   0.00527   0.00147  -0.0598   0.7984   1.0000
   1.500   0.4221   0.00534   0.00146  -0.0587   0.7771   1.0000
   1.750   0.4457   0.00543   0.00147  -0.0577   0.7537   1.0000
   2.000   0.4698   0.00555   0.00151  -0.0568   0.7269   1.0000
   2.250   0.4942   0.00570   0.00155  -0.0560   0.6969   1.0000
   2.500   0.5188   0.00589   0.00161  -0.0553   0.6646   1.0000
   2.750   0.5433   0.00610   0.00169  -0.0545   0.6295   1.0000
   3.000   0.5677   0.00636   0.00179  -0.0538   0.5894   1.0000
   3.250   0.5907   0.00674   0.00194  -0.0528   0.5282   1.0000
   3.500   0.6135   0.00719   0.00211  -0.0519   0.4617   1.0000
   3.750   0.6378   0.00757   0.00230  -0.0513   0.4116   1.0000
   4.000   0.6612   0.00807   0.00252  -0.0506   0.3503   1.0000
   4.250   0.6841   0.00866   0.00281  -0.0499   0.2787   1.0000
   4.500   0.7082   0.00915   0.00308  -0.0494   0.2283   1.0000
   4.750   0.7319   0.00970   0.00340  -0.0489   0.1756   1.0000
   5.000   0.7549   0.01037   0.00379  -0.0483   0.1180   1.0000
   5.250   0.7752   0.01147   0.00444  -0.0472   0.0394   1.0000
   5.500   0.7991   0.01210   0.00502  -0.0465   0.0254   1.0000
   5.750   0.8230   0.01273   0.00569  -0.0458   0.0214   1.0000
   6.000   0.8466   0.01339   0.00644  -0.0450   0.0201   1.0000
   6.250   0.8703   0.01400   0.00714  -0.0442   0.0194   1.0000
   6.500   0.8932   0.01471   0.00796  -0.0434   0.0186   1.0000
   6.750   0.9154   0.01554   0.00887  -0.0424   0.0179   1.0000
   7.000   0.9370   0.01647   0.00989  -0.0413   0.0172   1.0000
   7.250   0.9589   0.01735   0.01083  -0.0404   0.0163   1.0000
   7.500   0.9799   0.01839   0.01194  -0.0394   0.0154   1.0000
   7.750   0.9999   0.01983   0.01348  -0.0382   0.0150   1.0000
   8.000   1.0194   0.02177   0.01558  -0.0370   0.0146   1.0000
   8.250   1.0386   0.02433   0.01834  -0.0357   0.0143   1.0000
   8.500   1.0566   0.02734   0.02163  -0.0344   0.0142   1.0000
   8.750   1.0690   0.03142   0.02611  -0.0325   0.0140   1.0000
   9.000   1.0917   0.03043   0.02526  -0.0316   0.0131   1.0000
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