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EPPLER 562 AIRFOIL (e562-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 562 AIRFOIL (e562-il)
Reynolds number: 50,000
Max Cl/Cd: 11.63 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e562-il-50000.txt
Download as CSV file: xf-e562-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 562 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.3192   0.11771   0.11074  -0.0295   1.0000   0.2748
  -9.250  -0.3622   0.10457   0.09773  -0.0386   1.0000   0.1804
  -9.000  -0.3646   0.10012   0.09334  -0.0386   1.0000   0.1749
  -8.750  -0.3944   0.09374   0.08713  -0.0409   1.0000   0.1640
  -8.500  -0.5120   0.07709   0.07097  -0.0501   1.0000   0.1384
  -8.250  -0.5383   0.07094   0.06489  -0.0523   1.0000   0.1367
  -8.000  -0.5652   0.06459   0.05856  -0.0547   1.0000   0.1348
  -7.750  -0.5910   0.05733   0.05111  -0.0583   1.0000   0.1333
  -7.500  -0.6027   0.05030   0.04360  -0.0623   1.0000   0.1345
  -7.250  -0.5976   0.04443   0.03707  -0.0658   1.0000   0.1391
  -7.000  -0.5825   0.04296   0.03569  -0.0645   1.0000   0.1472
  -6.750  -0.5663   0.04037   0.03291  -0.0650   1.0000   0.1587
  -6.500  -0.5484   0.03851   0.03095  -0.0649   1.0000   0.1742
  -6.250  -0.5282   0.03665   0.02888  -0.0655   1.0000   0.1985
  -6.000  -0.5138   0.03733   0.02994  -0.0626   1.0000   0.2225
  -5.750  -0.4993   0.03807   0.03087  -0.0598   1.0000   0.2501
  -5.500  -0.4851   0.03890   0.03179  -0.0572   1.0000   0.2783
  -5.250  -0.4753   0.04064   0.03373  -0.0524   1.0000   0.2990
  -5.000  -0.4591   0.04059   0.03357  -0.0512   1.0000   0.3252
  -4.750  -0.4485   0.04155   0.03460  -0.0475   1.0000   0.3431
  -4.500  -0.4368   0.04211   0.03518  -0.0445   1.0000   0.3612
  -4.250  -0.4239   0.04231   0.03534  -0.0423   1.0000   0.3794
  -4.000  -0.4102   0.04231   0.03528  -0.0405   1.0000   0.3975
  -3.750  -0.3949   0.04204   0.03491  -0.0396   1.0000   0.4150
  -3.500  -0.3793   0.04176   0.03453  -0.0388   1.0000   0.4318
  -3.250  -0.3629   0.04143   0.03410  -0.0383   1.0000   0.4477
  -3.000  -0.3177   0.04119   0.03355  -0.0438   0.9908   0.4726
  -2.750  -0.2775   0.04161   0.03386  -0.0465   0.9778   0.4916
  -2.500  -0.2382   0.04176   0.03388  -0.0493   0.9648   0.5105
  -2.250  -0.1993   0.04171   0.03368  -0.0523   0.9519   0.5288
  -2.000  -0.1612   0.04155   0.03337  -0.0554   0.9393   0.5460
  -1.750  -0.1234   0.04138   0.03305  -0.0584   0.9265   0.5627
  -1.500  -0.0853   0.04122   0.03277  -0.0614   0.9141   0.5788
  -1.250  -0.0453   0.04113   0.03254  -0.0645   0.9022   0.5955
  -1.000  -0.0068   0.04102   0.03232  -0.0673   0.8900   0.6117
  -0.750   0.0225   0.04098   0.03220  -0.0687   0.8772   0.6269
  -0.500   0.0538   0.04099   0.03212  -0.0703   0.8647   0.6424
  -0.250   0.0885   0.04103   0.03209  -0.0722   0.8535   0.6590
   0.000   0.1259   0.04101   0.03201  -0.0742   0.8423   0.6763
   0.250   0.1488   0.04122   0.03218  -0.0744   0.8295   0.6921
   0.500   0.1752   0.04145   0.03239  -0.0750   0.8180   0.7088
   0.750   0.2178   0.04135   0.03223  -0.0772   0.8086   0.7294
   1.000   0.2344   0.04181   0.03268  -0.0765   0.7959   0.7460
   1.250   0.2546   0.04227   0.03314  -0.0762   0.7846   0.7648
   1.500   0.2932   0.04215   0.03300  -0.0774   0.7756   0.7890
   1.750   0.3043   0.04277   0.03367  -0.0758   0.7636   0.8092
   2.000   0.3196   0.04336   0.03430  -0.0748   0.7528   0.8344
   2.250   0.3508   0.04308   0.03408  -0.0744   0.7445   0.8686
   2.500   0.3525   0.04396   0.03510  -0.0720   0.7329   0.9070
   2.750   0.3948   0.04460   0.03579  -0.0770   0.7214   1.0000
   3.000   0.4685   0.04484   0.03585  -0.0863   0.7117   1.0000
   3.250   0.4887   0.04685   0.03774  -0.0893   0.6993   1.0000
   3.500   0.5190   0.04835   0.03910  -0.0922   0.6886   1.0000
   3.750   0.5673   0.04881   0.03944  -0.0957   0.6798   1.0000
   4.000   0.5691   0.05136   0.04191  -0.0952   0.6683   1.0000
   4.250   0.6012   0.05243   0.04290  -0.0966   0.6592   1.0000
   4.500   0.6153   0.05438   0.04482  -0.0967   0.6494   1.0000
   4.750   0.6236   0.05668   0.04708  -0.0964   0.6397   1.0000
   5.000   0.6614   0.05735   0.04772  -0.0976   0.6310   1.0000
   5.250   0.6500   0.06087   0.05122  -0.0963   0.6214   1.0000
   5.500   0.7045   0.06057   0.05094  -0.0980   0.6132   1.0000
   5.750   0.6740   0.06534   0.05570  -0.0961   0.6039   1.0000
   6.000   0.7101   0.06622   0.05659  -0.0968   0.5954   1.0000
   6.250   0.6948   0.07019   0.06058  -0.0958   0.5872   1.0000
   6.500   0.7253   0.07151   0.06191  -0.0963   0.5788   1.0000
   6.750   0.7103   0.07562   0.06605  -0.0957   0.5717   1.0000
   7.000   0.7425   0.07688   0.06736  -0.0961   0.5627   1.0000
   7.250   0.7240   0.08147   0.07198  -0.0957   0.5577   1.0000
   7.500   0.7645   0.08222   0.07279  -0.0961   0.5474   1.0000
   7.750   0.7416   0.08728   0.07788  -0.0960   0.5438   1.0000
   8.000   0.7356   0.09138   0.08203  -0.0964   0.5410   1.0000
   8.250   0.7360   0.09518   0.08589  -0.0970   0.5385   1.0000
   8.500   0.7574   0.09758   0.08836  -0.0973   0.5294   1.0000
   8.750   0.7579   0.10190   0.09274  -0.0983   0.5301   1.0000
   9.000   0.7698   0.10612   0.09705  -0.0997   0.5318   1.0000
   9.250   0.6828   0.11591   0.10690  -0.1034   0.6028   1.0000
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