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EPPLER E852 AIRFOIL (e852-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: EPPLER E852 AIRFOIL (e852-il)
Reynolds number: 500,000
Max Cl/Cd: 119.56 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e852-il-500000.txt
Download as CSV file: xf-e852-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER E852 AIRFOIL                             
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4430   0.08559   0.08350  -0.0427   1.0000   0.0110
  -8.750  -0.4508   0.08192   0.07989  -0.0430   1.0000   0.0111
  -8.500  -0.4611   0.07902   0.07703  -0.0423   1.0000   0.0114
  -8.250  -0.4668   0.07480   0.07286  -0.0446   0.9989   0.0117
  -8.000  -0.4612   0.06086   0.05886  -0.0654   0.9915   0.0111
  -7.750  -0.4455   0.05206   0.04982  -0.0787   0.9840   0.0112
  -7.500  -0.4287   0.04647   0.04400  -0.0852   0.9784   0.0116
  -7.250  -0.4094   0.04177   0.03906  -0.0898   0.9722   0.0124
  -7.000  -0.3836   0.03662   0.03357  -0.0945   0.9692   0.0131
  -6.750  -0.3639   0.03255   0.02912  -0.0959   0.9614   0.0140
  -6.500  -0.3330   0.02569   0.02146  -0.0972   0.9581   0.0098
  -6.250  -0.2992   0.02362   0.01913  -0.0993   0.9564   0.0096
  -6.000  -0.2792   0.01864   0.01361  -0.0997   0.9492   0.0111
  -5.750  -0.2460   0.01661   0.01132  -0.1011   0.9466   0.0115
  -5.500  -0.2112   0.01487   0.00934  -0.1025   0.9444   0.0105
  -5.250  -0.1809   0.01356   0.00790  -0.1031   0.9394   0.0102
  -5.000  -0.1499   0.01214   0.00631  -0.1040   0.9344   0.0103
  -4.750  -0.1156   0.01078   0.00476  -0.1057   0.9309   0.0114
  -4.500  -0.0861   0.00972   0.00359  -0.1060   0.9237   0.0294
  -4.250  -0.0536   0.00906   0.00312  -0.1075   0.9187   0.0804
  -4.000  -0.0257   0.00818   0.00267  -0.1084   0.9118   0.2034
  -3.750   0.0027   0.00726   0.00237  -0.1094   0.9056   0.3927
  -3.500   0.0312   0.00709   0.00229  -0.1097   0.8990   0.4516
  -3.250   0.0606   0.00702   0.00222  -0.1100   0.8930   0.4871
  -3.000   0.0893   0.00700   0.00215  -0.1103   0.8870   0.5087
  -2.750   0.1176   0.00698   0.00210  -0.1104   0.8808   0.5281
  -2.500   0.1466   0.00698   0.00204  -0.1107   0.8757   0.5431
  -2.250   0.1741   0.00697   0.00200  -0.1106   0.8697   0.5560
  -2.000   0.2030   0.00698   0.00199  -0.1109   0.8649   0.5722
  -1.750   0.2303   0.00700   0.00200  -0.1108   0.8594   0.5879
  -1.500   0.2583   0.00700   0.00197  -0.1109   0.8543   0.5994
  -1.250   0.2870   0.00701   0.00196  -0.1111   0.8500   0.6091
  -1.000   0.3144   0.00702   0.00196  -0.1111   0.8447   0.6183
  -0.750   0.3425   0.00701   0.00196  -0.1112   0.8402   0.6266
  -0.500   0.3708   0.00703   0.00198  -0.1114   0.8360   0.6355
  -0.250   0.3982   0.00704   0.00201  -0.1114   0.8310   0.6452
   0.000   0.4263   0.00704   0.00204  -0.1115   0.8267   0.6547
   0.250   0.4543   0.00706   0.00208  -0.1116   0.8223   0.6647
   0.500   0.4814   0.00706   0.00213  -0.1115   0.8168   0.6757
   0.750   0.5095   0.00708   0.00218  -0.1116   0.8120   0.6876
   1.000   0.5363   0.00708   0.00225  -0.1114   0.8059   0.7003
   1.250   0.5636   0.00708   0.00228  -0.1112   0.7993   0.7139
   1.500   0.5899   0.00706   0.00234  -0.1109   0.7915   0.7286
   1.750   0.6169   0.00705   0.00237  -0.1106   0.7837   0.7448
   2.000   0.6423   0.00701   0.00244  -0.1100   0.7735   0.7631
   2.250   0.6673   0.00697   0.00248  -0.1093   0.7617   0.7834
   2.500   0.6919   0.00692   0.00252  -0.1084   0.7491   0.8070
   2.750   0.7163   0.00689   0.00260  -0.1075   0.7382   0.8338
   3.000   0.7393   0.00683   0.00268  -0.1063   0.7270   0.8667
   3.250   0.7589   0.00674   0.00272  -0.1042   0.7116   0.9144
   3.500   0.7890   0.00669   0.00276  -0.1045   0.6919   1.0000
   3.750   0.8130   0.00680   0.00283  -0.1037   0.6634   1.0000
   4.000   0.8323   0.00710   0.00291  -0.1018   0.5993   1.0000
   4.250   0.8356   0.00841   0.00334  -0.0971   0.4254   1.0000
   4.500   0.8391   0.01016   0.00409  -0.0931   0.2341   1.0000
   4.750   0.8540   0.01120   0.00463  -0.0911   0.1376   1.0000
   5.000   0.8707   0.01211   0.00515  -0.0895   0.0699   1.0000
   5.250   0.8901   0.01281   0.00564  -0.0881   0.0334   1.0000
   5.500   0.9063   0.01390   0.00658  -0.0859   0.0058   1.0000
   5.750   0.9254   0.01473   0.00757  -0.0841   0.0048   1.0000
   6.000   0.9427   0.01572   0.00878  -0.0821   0.0046   1.0000
   6.250   0.9570   0.01708   0.01029  -0.0796   0.0044   1.0000
   6.500   0.9721   0.01860   0.01195  -0.0772   0.0044   1.0000
   6.750   0.9910   0.02000   0.01347  -0.0757   0.0044   1.0000
   7.000   1.0122   0.02141   0.01500  -0.0745   0.0045   1.0000
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