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EPPLER 908 AIRFOIL (e908-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 908 AIRFOIL (e908-il)
Reynolds number: 200,000
Max Cl/Cd: 62.83 at α=3°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e908-il-200000-n5.txt
Download as CSV file: xf-e908-il-200000-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 =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.4323   0.10602   0.10252  -0.0399   0.9937   0.0069
  -9.000  -0.4317   0.10176   0.09829  -0.0416   0.9921   0.0065
  -8.750  -0.4292   0.09780   0.09435  -0.0436   0.9903   0.0063
  -8.500  -0.4265   0.09350   0.09007  -0.0462   0.9881   0.0061
  -8.250  -0.4241   0.08925   0.08584  -0.0490   0.9859   0.0059
  -8.000  -0.4209   0.08503   0.08156  -0.0522   0.9839   0.0057
  -7.750  -0.4206   0.08042   0.07699  -0.0558   0.9812   0.0056
  -7.500  -0.4285   0.07665   0.07335  -0.0574   0.9764   0.0055
  -7.250  -0.4251   0.07062   0.06731  -0.0650   0.9722   0.0053
  -7.000  -0.4185   0.06444   0.06105  -0.0712   0.9684   0.0052
  -6.750  -0.4172   0.05917   0.05568  -0.0738   0.9634   0.0051
  -6.500  -0.4056   0.05311   0.04941  -0.0778   0.9601   0.0049
  -6.250  -0.3882   0.04694   0.04286  -0.0815   0.9579   0.0046
  -6.000  -0.3817   0.04250   0.03814  -0.0806   0.9531   0.0045
  -5.750  -0.3646   0.03754   0.03275  -0.0813   0.9499   0.0044
  -5.500  -0.3417   0.03303   0.02771  -0.0822   0.9479   0.0042
  -5.250  -0.3150   0.02919   0.02330  -0.0830   0.9464   0.0043
  -5.000  -0.2858   0.02616   0.01972  -0.0837   0.9454   0.0044
  -4.750  -0.2667   0.02417   0.01733  -0.0821   0.9421   0.0047
  -4.500  -0.2422   0.02239   0.01518  -0.0814   0.9396   0.0051
  -4.250  -0.2138   0.02232   0.01494  -0.0816   0.9373   0.0078
  -4.000  -0.1865   0.01985   0.01221  -0.0819   0.9361   0.0106
  -3.750  -0.1566   0.01858   0.01079  -0.0824   0.9348   0.0112
  -3.500  -0.1253   0.01756   0.00963  -0.0832   0.9336   0.0122
  -3.250  -0.0926   0.01677   0.00861  -0.0842   0.9325   0.0149
  -3.000  -0.0736   0.01617   0.00780  -0.0825   0.9278   0.0287
  -2.750  -0.0466   0.01512   0.00735  -0.0831   0.9253   0.1682
  -2.500  -0.0182   0.01417   0.00713  -0.0842   0.9234   0.3519
  -2.250   0.0079   0.01334   0.00725  -0.0842   0.9217   0.6088
  -2.000   0.0399   0.01326   0.00713  -0.0850   0.9202   0.6597
  -1.750   0.0609   0.01328   0.00711  -0.0837   0.9155   0.6863
  -1.500   0.0877   0.01325   0.00704  -0.0835   0.9121   0.7106
  -1.250   0.1182   0.01320   0.00686  -0.0841   0.9098   0.7298
  -1.000   0.1500   0.01312   0.00677  -0.0849   0.9081   0.7472
  -0.750   0.1827   0.01304   0.00668  -0.0860   0.9067   0.7652
  -0.500   0.1986   0.01312   0.00678  -0.0836   0.8999   0.7833
  -0.250   0.2258   0.01301   0.00676  -0.0833   0.8972   0.8123
   0.000   0.2544   0.01287   0.00671  -0.0832   0.8952   0.8430
   0.250   0.2845   0.01274   0.00664  -0.0835   0.8934   0.8639
   0.500   0.3007   0.01279   0.00675  -0.0812   0.8861   0.8821
   0.750   0.3323   0.01268   0.00668  -0.0820   0.8837   0.8995
   1.000   0.3678   0.01256   0.00662  -0.0837   0.8821   0.9196
   1.250   0.4084   0.01244   0.00658  -0.0866   0.8811   0.9456
   1.750   0.4666   0.01246   0.00677  -0.0878   0.8707   1.0000
   2.000   0.5016   0.01234   0.00672  -0.0894   0.8688   1.0000
   2.250   0.5213   0.01247   0.00691  -0.0880   0.8607   1.0000
   2.500   0.5538   0.01236   0.00691  -0.0891   0.8576   1.0000
   2.750   0.5894   0.01221   0.00690  -0.0907   0.8554   1.0000
   3.000   0.6622   0.01054   0.00392  -0.0955   0.5171   1.0000
   3.250   0.6361   0.01342   0.00482  -0.0858   0.1386   1.0000
   3.500   0.6440   0.01498   0.00571  -0.0822   0.0073   1.0000
   3.750   0.6649   0.01552   0.00640  -0.0808   0.0055   1.0000
   4.000   0.6841   0.01625   0.00730  -0.0791   0.0048   1.0000
   4.250   0.7023   0.01710   0.00828  -0.0772   0.0045   1.0000
   4.500   0.7198   0.01817   0.00946  -0.0751   0.0043   1.0000
   4.750   0.7398   0.01945   0.01083  -0.0736   0.0042   1.0000
   5.000   0.7672   0.02110   0.01259  -0.0734   0.0042   1.0000
   5.250   0.8018   0.02324   0.01490  -0.0743   0.0043   1.0000
   5.500   0.8392   0.02587   0.01781  -0.0754   0.0045   1.0000
   5.750   0.8721   0.02878   0.02110  -0.0755   0.0048   1.0000
   6.000   0.8987   0.03193   0.02476  -0.0743   0.0052   1.0000
   6.250   0.9194   0.03511   0.02833  -0.0724   0.0056   1.0000
   6.500   0.9353   0.03839   0.03199  -0.0698   0.0061   1.0000
   6.750   0.9468   0.04185   0.03580  -0.0667   0.0065   1.0000
   7.000   0.9533   0.04558   0.03988  -0.0633   0.0070   1.0000
  12.500   0.6877   0.14443   0.14160  -0.0557   0.0176   1.0000
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