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EPPLER 559 AIRFOIL (e559-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 559 AIRFOIL (e559-il)
Reynolds number: 50,000
Max Cl/Cd: 8.77 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e559-il-50000.txt
Download as CSV file: xf-e559-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 559 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
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.2735   0.12349   0.11634  -0.0318   1.0000   0.2695
 -10.000  -0.2702   0.12098   0.11387  -0.0311   1.0000   0.2797
  -9.750  -0.2949   0.12081   0.11385  -0.0311   1.0000   0.2825
  -9.500  -0.2722   0.11592   0.10892  -0.0297   1.0000   0.2969
  -9.250  -0.2631   0.11225   0.10528  -0.0285   1.0000   0.3074
  -9.000  -0.2834   0.11176   0.10493  -0.0275   1.0000   0.3149
  -8.750  -0.2730   0.10831   0.10149  -0.0259   1.0000   0.3301
  -8.500  -0.2662   0.10507   0.09828  -0.0245   1.0000   0.3409
  -8.250  -0.2928   0.10514   0.09852  -0.0222   1.0000   0.3487
  -8.000  -0.3794   0.09590   0.08963  -0.0280   1.0000   0.2505
  -7.750  -0.3950   0.09314   0.08696  -0.0258   1.0000   0.2492
  -7.500  -0.4156   0.09084   0.08477  -0.0231   1.0000   0.2486
  -7.250  -0.4405   0.08868   0.08275  -0.0203   1.0000   0.2479
  -7.000  -0.4655   0.08590   0.08008  -0.0187   1.0000   0.2472
  -6.750  -0.4903   0.08235   0.07664  -0.0186   1.0000   0.2468
  -6.500  -0.5154   0.07731   0.07167  -0.0208   1.0000   0.2470
  -6.250  -0.5128   0.07715   0.07153  -0.0171   1.0000   0.2586
  -6.000  -0.5313   0.07048   0.06482  -0.0227   1.0000   0.2621
  -5.750  -0.5326   0.06706   0.06136  -0.0241   1.0000   0.2712
  -5.500  -0.5299   0.06007   0.05410  -0.0320   1.0000   0.2842
  -5.250  -0.5220   0.06113   0.05520  -0.0274   1.0000   0.2942
  -5.000  -0.5130   0.05936   0.05335  -0.0276   1.0000   0.3059
  -4.750  -0.4716   0.05778   0.05152  -0.0332   0.9891   0.3259
  -4.500  -0.4283   0.05530   0.04875  -0.0404   0.9771   0.3459
  -4.250  -0.3825   0.05227   0.04532  -0.0492   0.9651   0.3669
  -4.000  -0.3437   0.05123   0.04408  -0.0531   0.9526   0.3840
  -3.750  -0.3127   0.05137   0.04421  -0.0531   0.9401   0.3968
  -3.500  -0.2748   0.05041   0.04307  -0.0565   0.9281   0.4126
  -3.250  -0.2322   0.04934   0.04180  -0.0608   0.9168   0.4287
  -3.000  -0.2006   0.04833   0.04062  -0.0635   0.9040   0.4427
  -2.750  -0.1657   0.04735   0.03945  -0.0668   0.8917   0.4575
  -2.500  -0.1259   0.04651   0.03839  -0.0708   0.8806   0.4730
  -2.250  -0.0830   0.04575   0.03744  -0.0748   0.8694   0.4888
  -2.000  -0.0613   0.04564   0.03729  -0.0743   0.8569   0.4996
  -1.750  -0.0289   0.04525   0.03678  -0.0763   0.8458   0.5126
  -1.500   0.0157   0.04468   0.03606  -0.0799   0.8357   0.5282
  -1.250   0.0403   0.04454   0.03577  -0.0813   0.8234   0.5409
  -1.000   0.0698   0.04449   0.03565  -0.0823   0.8128   0.5542
  -0.750   0.1075   0.04422   0.03531  -0.0842   0.8029   0.5685
  -0.500   0.1253   0.04453   0.03557  -0.0840   0.7913   0.5799
  -0.250   0.1644   0.04435   0.03530  -0.0862   0.7824   0.5957
   0.000   0.1873   0.04467   0.03554  -0.0869   0.7713   0.6095
   0.250   0.2076   0.04515   0.03600  -0.0866   0.7612   0.6229
   0.500   0.2438   0.04510   0.03591  -0.0881   0.7525   0.6395
   0.750   0.2527   0.04605   0.03687  -0.0867   0.7419   0.6515
   1.000   0.2971   0.04577   0.03654  -0.0888   0.7344   0.6720
   1.250   0.2968   0.04727   0.03805  -0.0869   0.7237   0.6836
   1.500   0.3446   0.04688   0.03762  -0.0890   0.7169   0.7077
   1.750   0.3361   0.04881   0.03959  -0.0864   0.7063   0.7196
   2.000   0.3814   0.04847   0.03925  -0.0880   0.7000   0.7470
   2.250   0.3679   0.05081   0.04164  -0.0854   0.6901   0.7618
   2.500   0.4041   0.05062   0.04152  -0.0855   0.6837   0.7942
   2.750   0.3896   0.05291   0.04391  -0.0827   0.6755   0.8146
   3.000   0.4049   0.05338   0.04454  -0.0809   0.6690   0.8586
   3.250   0.4140   0.05456   0.04596  -0.0807   0.6618   1.0000
   3.500   0.4567   0.05662   0.04779  -0.0880   0.6534   1.0000
   3.750   0.4915   0.05858   0.04951  -0.0923   0.6463   1.0000
   4.000   0.4955   0.06137   0.05218  -0.0932   0.6400   1.0000
   4.250   0.5453   0.06216   0.05277  -0.0961   0.6328   1.0000
   4.500   0.5328   0.06571   0.05627  -0.0954   0.6291   1.0000
   4.750   0.5308   0.06875   0.05926  -0.0953   0.6259   1.0000
   5.000   0.5355   0.07158   0.06204  -0.0956   0.6235   1.0000
   5.250   0.5371   0.07478   0.06522  -0.0960   0.6248   1.0000
   5.500   0.5418   0.07803   0.06845  -0.0966   0.6278   1.0000
   5.750   0.5578   0.08122   0.07161  -0.0980   0.6312   1.0000
   6.000   0.4690   0.08958   0.08018  -0.0988   0.7320   1.0000
   6.250   0.4858   0.09153   0.08208  -0.0992   0.7191   1.0000
   6.500   0.4915   0.09346   0.08401  -0.0987   0.7093   1.0000
   6.750   0.5253   0.09690   0.08741  -0.1009   0.7008   1.0000
   7.000   0.5245   0.09828   0.08878  -0.0998   0.6889   1.0000
   7.250   0.5508   0.10172   0.09221  -0.1014   0.6815   1.0000
   7.500   0.5612   0.10356   0.09405  -0.1012   0.6690   1.0000
   7.750   0.5672   0.10581   0.09632  -0.1010   0.6592   1.0000
   8.000   0.6025   0.10960   0.10011  -0.1030   0.6498   1.0000
   8.250   0.5977   0.11092   0.10145  -0.1018   0.6378   1.0000
   8.500   0.6235   0.11476   0.10531  -0.1033   0.6312   1.0000
   8.750   0.6317   0.11660   0.10719  -0.1031   0.6181   1.0000
   9.000   0.6345   0.11886   0.10948  -0.1028   0.6077   1.0000
   9.250   0.6682   0.12298   0.11364  -0.1046   0.5994   1.0000
   9.500   0.6616   0.12436   0.11506  -0.1037   0.5874   1.0000
   9.750   0.6812   0.12803   0.11876  -0.1047   0.5798   1.0000
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