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

EPPLER 332 AIRFOIL (e332-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 332 AIRFOIL (e332-il)
Reynolds number: 50,000
Max Cl/Cd: 17.44 at α=1.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e332-il-50000.txt
Download as CSV file: xf-e332-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 332 AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.4476   0.12175   0.11533  -0.0047   1.0000   0.1756
  -9.500  -0.4173   0.11555   0.10911  -0.0020   1.0000   0.1868
  -9.250  -0.4412   0.11471   0.10843  -0.0060   1.0000   0.1910
  -9.000  -0.4146   0.10918   0.10291  -0.0037   1.0000   0.2035
  -8.750  -0.4093   0.10550   0.09928  -0.0040   1.0000   0.2115
  -8.500  -0.4277   0.10384   0.09776  -0.0065   1.0000   0.2214
  -8.250  -0.4096   0.09959   0.09355  -0.0050   1.0000   0.2348
  -8.000  -0.4038   0.09611   0.09016  -0.0045   1.0000   0.2489
  -7.750  -0.4025   0.09295   0.08710  -0.0042   1.0000   0.2644
  -7.500  -0.4116   0.09029   0.08457  -0.0042   1.0000   0.2818
  -7.250  -0.3938   0.08654   0.08088  -0.0020   1.0000   0.3043
  -7.000  -0.4061   0.08437   0.07886  -0.0009   1.0000   0.3279
  -6.750  -0.3991   0.08164   0.07624   0.0019   1.0000   0.3586
  -6.500  -0.3620   0.07803   0.07264   0.0069   1.0000   0.4134
  -6.250  -0.3396   0.07557   0.07023   0.0121   1.0000   0.4790
  -6.000  -0.2676   0.07179   0.06638   0.0186   1.0000   0.6115
  -5.750  -0.1860   0.06657   0.06107   0.0184   1.0000   0.7476
  -5.000  -0.2328   0.06029   0.05531   0.0208   1.0000   0.6775
  -4.750  -0.2861   0.05845   0.05384   0.0217   1.0000   0.6137
  -4.500  -0.3330   0.05782   0.05359   0.0271   1.0000   0.6003
  -4.000  -0.4152   0.05962   0.05584   0.0448   1.0000   0.6357
  -3.750  -0.3667   0.04551   0.03856  -0.0119   0.9559   0.1716
  -3.500  -0.3117   0.04190   0.03385  -0.0153   0.9398   0.1276
  -3.250  -0.2633   0.03841   0.02979  -0.0184   0.9244   0.1187
  -3.000  -0.2097   0.03592   0.02662  -0.0218   0.9092   0.1117
  -2.750  -0.1521   0.03312   0.02336  -0.0258   0.8952   0.1090
  -2.500   0.1633   0.02203   0.01469  -0.0653   0.9090   1.0000
  -2.250   0.2086   0.02187   0.01387  -0.0684   0.8842   1.0000
  -2.000   0.2376   0.02203   0.01361  -0.0686   0.8599   1.0000
  -1.750   0.2618   0.02228   0.01352  -0.0679   0.8388   1.0000
  -1.500   0.2824   0.02266   0.01362  -0.0667   0.8185   1.0000
  -1.250   0.3030   0.02305   0.01376  -0.0654   0.8007   1.0000
  -1.000   0.3233   0.02347   0.01396  -0.0641   0.7844   1.0000
  -0.750   0.3445   0.02393   0.01421  -0.0631   0.7691   1.0000
  -0.500   0.3652   0.02444   0.01454  -0.0621   0.7549   1.0000
  -0.250   0.3852   0.02497   0.01493  -0.0610   0.7413   1.0000
   0.000   0.4053   0.02555   0.01536  -0.0599   0.7289   1.0000
   0.250   0.4257   0.02607   0.01572  -0.0586   0.7176   1.0000
   0.500   0.4453   0.02673   0.01629  -0.0575   0.7059   1.0000
   0.750   0.4639   0.02756   0.01705  -0.0567   0.6943   1.0000
   1.000   0.4832   0.02830   0.01770  -0.0556   0.6843   1.0000
   1.250   0.5029   0.02897   0.01829  -0.0544   0.6746   1.0000
   1.500   0.5195   0.03006   0.01937  -0.0535   0.6639   1.0000
   1.750   0.5385   0.03087   0.02012  -0.0523   0.6554   1.0000
   2.000   0.5551   0.03192   0.02116  -0.0513   0.6458   1.0000
   2.250   0.5696   0.03320   0.02243  -0.0502   0.6365   1.0000
   2.500   0.5889   0.03398   0.02317  -0.0489   0.6287   1.0000
   2.750   0.5978   0.03576   0.02502  -0.0478   0.6194   1.0000
   3.000   0.6198   0.03633   0.02553  -0.0463   0.6124   1.0000
   3.250   0.6209   0.03872   0.02801  -0.0450   0.6031   1.0000
   3.500   0.6457   0.03913   0.02837  -0.0435   0.5967   1.0000
   3.750   0.6357   0.04225   0.03159  -0.0418   0.5876   1.0000
   4.000   0.6629   0.04259   0.03190  -0.0404   0.5814   1.0000
   4.250   0.6369   0.04664   0.03602  -0.0378   0.5734   1.0000
   4.500   0.6657   0.04701   0.03641  -0.0366   0.5667   1.0000
   4.750   0.6265   0.05161   0.04102  -0.0330   0.5609   1.0000
   5.000   0.6194   0.05413   0.04353  -0.0306   0.5551   1.0000
   5.250   0.6259   0.05603   0.04544  -0.0288   0.5490   1.0000
   5.500   0.5857   0.05981   0.04914  -0.0242   0.5473   1.0000
   5.750   0.5636   0.06291   0.05218  -0.0213   0.5458   1.0000
   6.000   0.5511   0.06578   0.05503  -0.0193   0.5445   1.0000
   6.250   0.5428   0.06880   0.05804  -0.0182   0.5461   1.0000
   6.500   0.5413   0.07179   0.06104  -0.0177   0.5483   1.0000
   6.750   0.5531   0.07482   0.06411  -0.0181   0.5509   1.0000
   7.000   0.4511   0.08208   0.07132  -0.0195   0.6600   1.0000
   7.250   0.4596   0.08402   0.07329  -0.0186   0.6459   1.0000
   7.500   0.4661   0.08602   0.07531  -0.0178   0.6334   1.0000
   7.750   0.4767   0.08835   0.07766  -0.0174   0.6211   1.0000
   8.000   0.4922   0.09097   0.08032  -0.0174   0.6088   1.0000
   8.250   0.5188   0.09444   0.08385  -0.0183   0.5964   1.0000
   8.500   0.5243   0.09628   0.08573  -0.0174   0.5825   1.0000
   8.750   0.5259   0.09806   0.08754  -0.0163   0.5692   1.0000
   9.000   0.5296   0.10027   0.08981  -0.0156   0.5568   1.0000
   9.250   0.5390   0.10300   0.09260  -0.0155   0.5463   1.0000
   9.500   0.5610   0.10653   0.09622  -0.0160   0.5348   1.0000
   9.750   0.5751   0.10931   0.09908  -0.0159   0.5213   1.0000
  10.000   0.5678   0.11083   0.10061  -0.0149   0.5096   1.0000
  10.250   0.5720   0.11339   0.10325  -0.0147   0.4979   1.0000
<< Back to EPPLER 332 AIRFOIL (e332-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 332 AIRFOIL (e332-il)