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EPPLER E851 AIRFOIL (e851-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: EPPLER E851 AIRFOIL (e851-il)
Reynolds number: 200,000
Max Cl/Cd: 73.47 at α=3.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e851-il-200000-n5.txt
Download as CSV file: xf-e851-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER E851 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.4702   0.08592   0.08244  -0.0391   1.0000   0.0061
  -8.500  -0.4777   0.08229   0.07888  -0.0395   1.0000   0.0060
  -8.250  -0.4886   0.07867   0.07533  -0.0395   1.0000   0.0059
  -8.000  -0.5033   0.07541   0.07215  -0.0390   1.0000   0.0059
  -7.750  -0.5129   0.07003   0.06682  -0.0433   0.9974   0.0058
  -7.500  -0.5029   0.06063   0.05725  -0.0572   0.9898   0.0057
  -7.250  -0.4915   0.05369   0.05006  -0.0647   0.9835   0.0055
  -7.000  -0.4792   0.04753   0.04359  -0.0693   0.9775   0.0053
  -6.750  -0.4633   0.04195   0.03763  -0.0723   0.9720   0.0051
  -6.500  -0.4405   0.03657   0.03179  -0.0752   0.9688   0.0048
  -6.250  -0.4235   0.03222   0.02696  -0.0753   0.9627   0.0046
  -6.000  -0.3985   0.02799   0.02215  -0.0762   0.9593   0.0044
  -5.750  -0.3701   0.02461   0.01822  -0.0770   0.9572   0.0043
  -5.500  -0.3412   0.02198   0.01514  -0.0776   0.9553   0.0043
  -5.250  -0.3185   0.02010   0.01297  -0.0767   0.9507   0.0043
  -5.000  -0.2907   0.01845   0.01107  -0.0768   0.9480   0.0045
  -4.750  -0.2613   0.01706   0.00949  -0.0774   0.9460   0.0048
  -4.500  -0.2303   0.01588   0.00802  -0.0783   0.9443   0.0055
  -4.250  -0.1995   0.01500   0.00691  -0.0792   0.9423   0.0069
  -4.000  -0.1752   0.01435   0.00621  -0.0788   0.9378   0.0166
  -3.750  -0.1465   0.01317   0.00549  -0.0798   0.9353   0.1021
  -3.500  -0.1163   0.01227   0.00498  -0.0812   0.9334   0.2219
  -3.250  -0.0863   0.01134   0.00481  -0.0825   0.9318   0.4303
  -3.000  -0.0537   0.01116   0.00473  -0.0836   0.9305   0.4994
  -2.750  -0.0295   0.01114   0.00468  -0.0830   0.9261   0.5280
  -2.500  -0.0007   0.01109   0.00455  -0.0832   0.9231   0.5558
  -2.250   0.0303   0.01104   0.00451  -0.0839   0.9210   0.5852
  -2.000   0.0626   0.01097   0.00442  -0.0849   0.9193   0.6029
  -1.750   0.0960   0.01091   0.00431  -0.0861   0.9179   0.6138
  -1.500   0.1228   0.01090   0.00427  -0.0861   0.9142   0.6237
  -1.250   0.1503   0.01088   0.00422  -0.0861   0.9106   0.6328
  -1.000   0.1812   0.01083   0.00417  -0.0868   0.9081   0.6422
  -0.750   0.2136   0.01077   0.00411  -0.0879   0.9062   0.6524
  -0.500   0.2467   0.01071   0.00406  -0.0891   0.9045   0.6631
  -0.250   0.2708   0.01074   0.00414  -0.0884   0.8996   0.6732
   0.000   0.2998   0.01071   0.00417  -0.0888   0.8961   0.6843
   0.250   0.3316   0.01065   0.00416  -0.0896   0.8935   0.6961
   0.500   0.3649   0.01056   0.00415  -0.0908   0.8914   0.7088
   0.750   0.3883   0.01059   0.00427  -0.0899   0.8853   0.7221
   1.000   0.4188   0.01052   0.00430  -0.0904   0.8813   0.7367
   1.250   0.4526   0.01039   0.00428  -0.0916   0.8781   0.7526
   1.500   0.4760   0.01038   0.00446  -0.0906   0.8704   0.7700
   1.750   0.5094   0.01019   0.00440  -0.0914   0.8652   0.7888
   2.000   0.5340   0.01008   0.00443  -0.0904   0.8549   0.8104
   2.250   0.5609   0.00989   0.00438  -0.0897   0.8428   0.8353
   2.500   0.5895   0.00960   0.00421  -0.0891   0.8249   0.8644
   2.750   0.6150   0.00936   0.00415  -0.0878   0.7992   0.9039
   3.000   0.6488   0.00919   0.00411  -0.0886   0.7689   0.9902
   3.250   0.6752   0.00919   0.00406  -0.0879   0.7217   1.0000
   3.500   0.6932   0.00976   0.00390  -0.0852   0.5670   1.0000
   3.750   0.6944   0.01130   0.00440  -0.0801   0.3729   1.0000
   4.000   0.7028   0.01275   0.00503  -0.0770   0.2117   1.0000
   4.250   0.7193   0.01377   0.00563  -0.0753   0.1208   1.0000
   4.500   0.7391   0.01458   0.00624  -0.0741   0.0727   1.0000
   4.750   0.7582   0.01551   0.00692  -0.0727   0.0318   1.0000
   5.000   0.7773   0.01656   0.00788  -0.0711   0.0103   1.0000
   5.250   0.7975   0.01751   0.00884  -0.0698   0.0051   1.0000
   5.500   0.8177   0.01856   0.01019  -0.0683   0.0042   1.0000
   5.750   0.8373   0.01985   0.01166  -0.0667   0.0037   1.0000
   6.000   0.8579   0.02134   0.01332  -0.0653   0.0034   1.0000
   6.250   0.8803   0.02314   0.01532  -0.0642   0.0032   1.0000
   6.500   0.9043   0.02531   0.01775  -0.0632   0.0031   1.0000
   6.750   0.9279   0.02788   0.02068  -0.0622   0.0030   1.0000
   7.000   0.9487   0.03093   0.02416  -0.0606   0.0030   1.0000
   7.250   0.9651   0.03447   0.02819  -0.0584   0.0031   1.0000
   7.500   0.9767   0.03843   0.03266  -0.0555   0.0031   1.0000
   7.750   0.9833   0.04277   0.03748  -0.0522   0.0032   1.0000
   8.000   0.9859   0.04723   0.04235  -0.0487   0.0033   1.0000
   8.250   0.9846   0.05174   0.04722  -0.0452   0.0034   1.0000
   8.500   0.9798   0.05611   0.05190  -0.0418   0.0035   1.0000
   8.750   0.9713   0.06026   0.05630  -0.0384   0.0036   1.0000
   9.000   0.9574   0.06395   0.06017  -0.0347   0.0036   1.0000
   9.250   0.9414   0.06769   0.06408  -0.0317   0.0037   1.0000
   9.500   0.9237   0.07181   0.06836  -0.0299   0.0037   1.0000
   9.750   0.9055   0.07644   0.07312  -0.0295   0.0038   1.0000
  10.000   0.8867   0.08192   0.07872  -0.0309   0.0038   1.0000
  10.250   0.8694   0.08845   0.08534  -0.0343   0.0038   1.0000
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