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EPPLER 908 AIRFOIL (e908-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 908 AIRFOIL (e908-il)
Reynolds number: 50,000
Max Cl/Cd: 24.42 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e908-il-50000-n5.txt
Download as CSV file: xf-e908-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 908 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4023   0.09843   0.09219  -0.0343   1.0000   0.0379
  -8.250  -0.4806   0.10201   0.09553  -0.0351   1.0000   0.0392
  -8.000  -0.4859   0.09903   0.09262  -0.0345   1.0000   0.0381
  -7.750  -0.4943   0.09585   0.08954  -0.0339   1.0000   0.0369
  -7.500  -0.5053   0.09282   0.08661  -0.0333   1.0000   0.0361
  -7.250  -0.5154   0.08950   0.08337  -0.0334   1.0000   0.0354
  -7.000  -0.5222   0.08538   0.07931  -0.0351   1.0000   0.0345
  -6.750  -0.5278   0.08076   0.07472  -0.0369   1.0000   0.0334
  -6.500  -0.5312   0.07572   0.06963  -0.0390   1.0000   0.0321
  -6.250  -0.5313   0.06973   0.06342  -0.0422   1.0000   0.0302
  -6.000  -0.5241   0.06581   0.05922  -0.0428   1.0000   0.0291
  -5.750  -0.5182   0.06134   0.05461  -0.0432   1.0000   0.0287
  -5.500  -0.5089   0.05694   0.04997  -0.0438   1.0000   0.0282
  -5.250  -0.4958   0.05240   0.04491  -0.0447   1.0000   0.0279
  -5.000  -0.4790   0.04815   0.04023  -0.0454   1.0000   0.0276
  -4.750  -0.4587   0.04413   0.03570  -0.0460   1.0000   0.0275
  -4.500  -0.4358   0.04048   0.03146  -0.0464   1.0000   0.0275
  -4.250  -0.4107   0.03719   0.02754  -0.0465   1.0000   0.0279
  -4.000  -0.3852   0.03449   0.02425  -0.0461   1.0000   0.0284
  -3.750  -0.3598   0.03235   0.02161  -0.0454   1.0000   0.0296
  -3.500  -0.3368   0.03025   0.01934  -0.0452   1.0000   0.0354
  -3.250  -0.3128   0.02892   0.01762  -0.0440   1.0000   0.0412
  -3.000  -0.2892   0.02731   0.01585  -0.0429   1.0000   0.0454
  -2.750  -0.2646   0.02611   0.01436  -0.0421   1.0000   0.0538
  -2.500  -0.2372   0.02463   0.01306  -0.0423   1.0000   0.0935
  -2.250  -0.2200   0.02172   0.01323  -0.0408   1.0000   0.6062
  -2.000  -0.2169   0.02148   0.01328  -0.0337   1.0000   0.7911
  -1.750  -0.2019   0.02130   0.01272  -0.0302   1.0000   0.8621
  -1.500  -0.1694   0.02081   0.01192  -0.0313   1.0000   1.0000
  -1.250  -0.1453   0.02093   0.01159  -0.0316   1.0000   1.0000
  -1.000  -0.1214   0.02111   0.01136  -0.0318   1.0000   1.0000
  -0.750  -0.0957   0.02137   0.01127  -0.0323   0.9991   1.0000
  -0.500  -0.0604   0.02192   0.01144  -0.0348   0.9944   1.0000
  -0.250  -0.0273   0.02235   0.01158  -0.0368   0.9890   1.0000
   0.000   0.0071   0.02290   0.01172  -0.0390   0.9841   1.0000
   0.250   0.0394   0.02336   0.01197  -0.0408   0.9787   1.0000
   0.500   0.0719   0.02385   0.01230  -0.0426   0.9731   1.0000
   0.750   0.1044   0.02436   0.01267  -0.0445   0.9677   1.0000
   1.000   0.1351   0.02479   0.01302  -0.0459   0.9615   1.0000
   1.250   0.1680   0.02533   0.01348  -0.0478   0.9562   1.0000
   1.500   0.1972   0.02574   0.01385  -0.0489   0.9494   1.0000
   1.750   0.2304   0.02628   0.01438  -0.0508   0.9440   1.0000
   2.000   0.2582   0.02669   0.01481  -0.0516   0.9367   1.0000
   2.250   0.2910   0.02723   0.01540  -0.0534   0.9309   1.0000
   2.500   0.3186   0.02766   0.01591  -0.0542   0.9233   1.0000
   2.750   0.3493   0.02817   0.01664  -0.0555   0.9168   1.0000
   3.000   0.3787   0.02864   0.01725  -0.0565   0.9094   1.0000
   3.250   0.4065   0.02911   0.01788  -0.0573   0.9015   1.0000
   3.500   0.4397   0.02962   0.01862  -0.0589   0.8948   1.0000
   3.750   0.4650   0.03006   0.01928  -0.0591   0.8853   1.0000
   4.000   0.4941   0.03052   0.02004  -0.0600   0.8765   1.0000
   4.250   0.5282   0.03098   0.02109  -0.0615   0.8685   1.0000
   4.500   0.5542   0.03138   0.02190  -0.0616   0.8575   1.0000
   4.750   0.5823   0.03175   0.02275  -0.0619   0.8459   1.0000
   5.000   0.6646   0.02722   0.01439  -0.0545   0.0473   1.0000
   5.250   0.6829   0.02883   0.01608  -0.0529   0.0328   1.0000
   5.500   0.7046   0.03079   0.01823  -0.0517   0.0289   1.0000
   5.750   0.7518   0.03301   0.02086  -0.0537   0.0264   1.0000
   6.000   0.8204   0.03668   0.02495  -0.0589   0.0248   1.0000
   6.250   0.8634   0.04022   0.02894  -0.0603   0.0246   1.0000
   6.500   0.8932   0.04364   0.03306  -0.0598   0.0245   1.0000
   6.750   0.9146   0.04688   0.03678  -0.0582   0.0241   1.0000
   7.000   0.9302   0.05001   0.04038  -0.0559   0.0232   1.0000
   7.250   0.9415   0.05311   0.04390  -0.0533   0.0222   1.0000
   7.500   0.9489   0.05616   0.04731  -0.0506   0.0209   1.0000
   7.750   0.9539   0.05933   0.05081  -0.0479   0.0203   1.0000
   8.000   0.9562   0.06262   0.05433  -0.0453   0.0193   1.0000
   8.250   0.9550   0.06612   0.05810  -0.0425   0.0188   1.0000
   8.500   0.9505   0.06972   0.06196  -0.0394   0.0186   1.0000
   8.750   0.9446   0.07303   0.06553  -0.0363   0.0186   1.0000
   9.000   0.9349   0.07611   0.06884  -0.0329   0.0187   1.0000
   9.250   0.9187   0.07926   0.07236  -0.0290   0.0201   1.0000
   9.500   0.9029   0.08301   0.07629  -0.0268   0.0205   1.0000
   9.750   0.8862   0.08711   0.08054  -0.0256   0.0208   1.0000
  10.000   0.8679   0.09208   0.08566  -0.0258   0.0224   1.0000
  10.250   0.8520   0.09729   0.09096  -0.0273   0.0231   1.0000
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