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EPPLER E851 AIRFOIL (e851-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: EPPLER E851 AIRFOIL (e851-il)
Reynolds number: 500,000
Max Cl/Cd: 90.54 at α=2.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e851-il-500000-n5.txt
Download as CSV file: xf-e851-il-500000-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 =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.4462   0.11004   0.10766  -0.0331   1.0000   0.0077
 -10.000  -0.4458   0.10634   0.10399  -0.0338   1.0000   0.0077
  -8.500  -0.4552   0.07922   0.07703  -0.0455   0.9967   0.0065
  -8.250  -0.4516   0.07207   0.06990  -0.0535   0.9936   0.0055
  -8.000  -0.4490   0.06349   0.06132  -0.0648   0.9870   0.0055
  -7.750  -0.4415   0.05416   0.05180  -0.0788   0.9799   0.0052
  -7.500  -0.4405   0.04173   0.03893  -0.0853   0.9702   0.0032
  -7.250  -0.4213   0.03877   0.03578  -0.0866   0.9651   0.0030
  -7.000  -0.4044   0.03526   0.03203  -0.0885   0.9596   0.0027
  -6.750  -0.3851   0.02922   0.02551  -0.0910   0.9563   0.0025
  -6.500  -0.3670   0.02511   0.02096  -0.0911   0.9500   0.0024
  -6.250  -0.3414   0.02125   0.01660  -0.0920   0.9466   0.0023
  -6.000  -0.3121   0.01820   0.01311  -0.0930   0.9444   0.0022
  -5.750  -0.2836   0.01601   0.01057  -0.0935   0.9418   0.0021
  -5.500  -0.2577   0.01430   0.00861  -0.0933   0.9373   0.0021
  -5.250  -0.2285   0.01287   0.00697  -0.0939   0.9342   0.0021
  -5.000  -0.1979   0.01171   0.00557  -0.0947   0.9316   0.0021
  -4.750  -0.1684   0.01088   0.00458  -0.0953   0.9283   0.0022
  -4.500  -0.1401   0.01030   0.00384  -0.0956   0.9241   0.0023
  -4.250  -0.1096   0.00986   0.00324  -0.0963   0.9208   0.0026
  -4.000  -0.0782   0.00950   0.00277  -0.0971   0.9180   0.0047
  -3.750  -0.0504   0.00904   0.00246  -0.0975   0.9139   0.0342
  -3.500  -0.0236   0.00827   0.00214  -0.0980   0.9097   0.1450
  -3.250   0.0046   0.00757   0.00185  -0.0988   0.9062   0.2698
  -3.000   0.0323   0.00698   0.00170  -0.0994   0.9028   0.4095
  -2.750   0.0595   0.00682   0.00162  -0.0994   0.8987   0.4595
  -2.500   0.0882   0.00675   0.00155  -0.0997   0.8951   0.4822
  -2.250   0.1177   0.00669   0.00150  -0.1002   0.8920   0.5049
  -2.000   0.1456   0.00664   0.00148  -0.1003   0.8884   0.5286
  -1.750   0.1732   0.00660   0.00147  -0.1003   0.8845   0.5458
  -1.250   0.2311   0.00656   0.00138  -0.1009   0.8782   0.5654
  -1.000   0.2585   0.00655   0.00138  -0.1009   0.8744   0.5731
  -0.750   0.2862   0.00654   0.00138  -0.1010   0.8705   0.5809
  -0.500   0.3148   0.00653   0.00138  -0.1012   0.8670   0.5885
  -0.250   0.3431   0.00653   0.00140  -0.1014   0.8631   0.5968
   0.000   0.3701   0.00652   0.00143  -0.1013   0.8578   0.6047
   0.250   0.3985   0.00651   0.00143  -0.1014   0.8526   0.6133
   0.500   0.4252   0.00651   0.00147  -0.1012   0.8455   0.6223
   0.750   0.4531   0.00650   0.00149  -0.1012   0.8380   0.6314
   1.000   0.4795   0.00649   0.00152  -0.1008   0.8278   0.6411
   1.250   0.5059   0.00649   0.00160  -0.1005   0.8162   0.6514
   1.500   0.5320   0.00650   0.00162  -0.1000   0.8010   0.6625
   1.750   0.5568   0.00651   0.00165  -0.0993   0.7773   0.6741
   2.000   0.5812   0.00657   0.00168  -0.0984   0.7470   0.6862
   2.250   0.6048   0.00668   0.00174  -0.0974   0.7084   0.6994
   2.500   0.6233   0.00704   0.00185  -0.0953   0.6305   0.7136
   2.750   0.6316   0.00809   0.00227  -0.0914   0.4728   0.7287
   3.000   0.6421   0.00928   0.00275  -0.0884   0.3013   0.7455
   3.250   0.6598   0.01003   0.00313  -0.0867   0.1991   0.7646
   3.500   0.6801   0.01057   0.00347  -0.0855   0.1332   0.7861
   3.750   0.6984   0.01127   0.00388  -0.0839   0.0584   0.8113
   4.000   0.7177   0.01177   0.00426  -0.0823   0.0199   0.8424
   4.250   0.7360   0.01212   0.00464  -0.0803   0.0042   0.8866
   4.500   0.7607   0.01245   0.00513  -0.0795   0.0019   1.0000
   4.750   0.7845   0.01291   0.00569  -0.0788   0.0017   1.0000
   5.000   0.8073   0.01350   0.00639  -0.0778   0.0015   1.0000
   5.250   0.8289   0.01425   0.00735  -0.0765   0.0014   1.0000
   5.500   0.8497   0.01515   0.00837  -0.0752   0.0013   1.0000
   5.750   0.8697   0.01629   0.00966  -0.0736   0.0013   1.0000
   6.000   0.8903   0.01761   0.01115  -0.0721   0.0013   1.0000
   6.250   0.9119   0.01926   0.01299  -0.0709   0.0013   1.0000
   6.500   0.9345   0.02130   0.01529  -0.0697   0.0013   1.0000
   6.750   0.9565   0.02374   0.01806  -0.0684   0.0014   1.0000
   7.000   0.9755   0.02677   0.02148  -0.0665   0.0014   1.0000
   7.250   0.9903   0.03028   0.02544  -0.0640   0.0015   1.0000
   7.500   1.0003   0.03427   0.02986  -0.0608   0.0015   1.0000
   7.750   1.0047   0.03888   0.03491  -0.0571   0.0016   1.0000
   8.000   1.0051   0.04373   0.04014  -0.0532   0.0016   1.0000
   8.250   1.0009   0.04872   0.04546  -0.0491   0.0017   1.0000
   8.500   0.9923   0.05356   0.05058  -0.0452   0.0017   1.0000
   8.750   0.9790   0.05800   0.05523  -0.0412   0.0018   1.0000
   9.000   0.9583   0.06049   0.05787  -0.0366   0.0019   1.0000
   9.250   0.9464   0.06374   0.06125  -0.0337   0.0019   1.0000
   9.500   0.9339   0.06730   0.06493  -0.0318   0.0020   1.0000
   9.750   0.9192   0.07145   0.06921  -0.0310   0.0020   1.0000
  10.000   0.9031   0.07641   0.07428  -0.0315   0.0020   1.0000
  10.250   0.8829   0.08308   0.08106  -0.0343   0.0020   1.0000
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