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EPPLER E851 AIRFOIL (e851-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: EPPLER E851 AIRFOIL (e851-il)
Reynolds number: 100,000
Max Cl/Cd: 50.74 at α=4.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e851-il-100000-n5.txt
Download as CSV file: xf-e851-il-100000-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 =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4615   0.09223   0.08724  -0.0395   1.0000   0.0110
  -8.750  -0.4671   0.08819   0.08329  -0.0404   1.0000   0.0107
  -8.500  -0.4744   0.08415   0.07933  -0.0412   1.0000   0.0104
  -8.250  -0.4842   0.08012   0.07539  -0.0419   1.0000   0.0102
  -8.000  -0.4976   0.07604   0.07141  -0.0424   1.0000   0.0100
  -7.750  -0.5165   0.07206   0.06752  -0.0428   1.0000   0.0098
  -7.500  -0.5337   0.06756   0.06304  -0.0445   1.0000   0.0096
  -7.250  -0.5470   0.06331   0.05871  -0.0449   1.0000   0.0094
  -7.000  -0.5562   0.05900   0.05427  -0.0449   1.0000   0.0092
  -6.750  -0.5597   0.05473   0.04982  -0.0447   1.0000   0.0089
  -6.500  -0.5570   0.05064   0.04545  -0.0445   1.0000   0.0086
  -6.250  -0.5361   0.04555   0.03990  -0.0478   0.9964   0.0084
  -6.000  -0.5119   0.04104   0.03487  -0.0505   0.9931   0.0082
  -5.750  -0.4877   0.03726   0.03057  -0.0522   0.9897   0.0081
  -5.500  -0.4619   0.03344   0.02618  -0.0537   0.9868   0.0082
  -5.250  -0.4343   0.03003   0.02220  -0.0548   0.9844   0.0083
  -5.000  -0.4053   0.02715   0.01885  -0.0557   0.9825   0.0087
  -4.750  -0.3763   0.02541   0.01685  -0.0568   0.9804   0.0107
  -4.500  -0.3495   0.02426   0.01545  -0.0569   0.9775   0.0141
  -4.250  -0.3209   0.02306   0.01385  -0.0570   0.9749   0.0173
  -4.000  -0.2932   0.02098   0.01166  -0.0573   0.9728   0.0207
  -3.750  -0.2625   0.01982   0.01042  -0.0582   0.9706   0.0384
  -3.500  -0.2308   0.01838   0.00915  -0.0596   0.9689   0.0912
  -3.250  -0.2055   0.01647   0.00845  -0.0608   0.9666   0.3445
  -3.000  -0.1806   0.01617   0.00860  -0.0603   0.9630   0.5081
  -2.750  -0.1525   0.01621   0.00862  -0.0603   0.9598   0.5669
  -2.500  -0.1229   0.01631   0.00865  -0.0605   0.9571   0.6105
  -2.250  -0.0975   0.01639   0.00871  -0.0599   0.9537   0.6448
  -2.000  -0.0717   0.01640   0.00854  -0.0595   0.9497   0.6654
  -1.750  -0.0422   0.01641   0.00844  -0.0600   0.9466   0.6787
  -1.500  -0.0105   0.01644   0.00836  -0.0610   0.9440   0.6914
  -1.250   0.0157   0.01645   0.00831  -0.0609   0.9401   0.7039
  -1.000   0.0426   0.01648   0.00828  -0.0610   0.9360   0.7167
  -0.750   0.0732   0.01651   0.00828  -0.0618   0.9329   0.7301
  -0.500   0.1057   0.01654   0.00824  -0.0629   0.9305   0.7443
  -0.250   0.1287   0.01658   0.00830  -0.0621   0.9252   0.7591
   0.000   0.1574   0.01661   0.00835  -0.0624   0.9213   0.7759
   0.250   0.1892   0.01662   0.00841  -0.0634   0.9183   0.7948
   0.500   0.2131   0.01665   0.00851  -0.0627   0.9132   0.8164
   0.750   0.2400   0.01665   0.00861  -0.0625   0.9085   0.8433
   1.000   0.2725   0.01659   0.00867  -0.0634   0.9053   0.8793
   1.250   0.3086   0.01654   0.00876  -0.0654   0.9006   0.9516
   1.500   0.3410   0.01661   0.00884  -0.0668   0.8951   1.0000
   1.750   0.3790   0.01666   0.00893  -0.0691   0.8917   1.0000
   2.000   0.4048   0.01681   0.00922  -0.0690   0.8840   1.0000
   2.250   0.4424   0.01680   0.00930  -0.0711   0.8795   1.0000
   2.500   0.4709   0.01687   0.00946  -0.0713   0.8713   1.0000
   2.750   0.5126   0.01665   0.00940  -0.0738   0.8658   1.0000
   3.000   0.5446   0.01650   0.00940  -0.0742   0.8551   1.0000
   3.250   0.5855   0.01601   0.00921  -0.0759   0.8418   1.0000
   3.500   0.6243   0.01519   0.00859  -0.0764   0.8177   1.0000
   3.750   0.6589   0.01449   0.00807  -0.0761   0.7853   1.0000
   4.000   0.6820   0.01418   0.00790  -0.0741   0.7384   1.0000
   4.250   0.7230   0.01425   0.00685  -0.0739   0.4756   1.0000
   4.500   0.7270   0.01629   0.00760  -0.0699   0.2491   1.0000
   4.750   0.7389   0.01817   0.00855  -0.0677   0.1033   1.0000
   5.000   0.7562   0.01957   0.00972  -0.0659   0.0622   1.0000
   5.250   0.7735   0.02100   0.01107  -0.0640   0.0370   1.0000
   5.500   0.7914   0.02245   0.01256  -0.0622   0.0227   1.0000
   5.750   0.8104   0.02410   0.01451  -0.0604   0.0163   1.0000
   6.000   0.8311   0.02697   0.01752  -0.0590   0.0138   1.0000
   6.250   0.8581   0.02923   0.02003  -0.0584   0.0123   1.0000
   6.500   0.8831   0.03092   0.02203  -0.0577   0.0095   1.0000
   6.750   0.9048   0.03258   0.02396  -0.0568   0.0075   1.0000
   7.000   0.9243   0.03495   0.02666  -0.0555   0.0067   1.0000
   7.250   0.9414   0.03812   0.03028  -0.0537   0.0065   1.0000
   7.500   0.9545   0.04151   0.03413  -0.0515   0.0064   1.0000
   7.750   0.9631   0.04512   0.03820  -0.0488   0.0063   1.0000
   8.000   0.9681   0.04882   0.04234  -0.0459   0.0063   1.0000
   8.250   0.9690   0.05266   0.04658  -0.0428   0.0063   1.0000
   8.500   0.9661   0.05657   0.05085  -0.0397   0.0064   1.0000
   8.750   0.9593   0.06052   0.05512  -0.0365   0.0065   1.0000
   9.000   0.9469   0.06424   0.05908  -0.0330   0.0066   1.0000
   9.250   0.9321   0.06818   0.06323  -0.0302   0.0067   1.0000
   9.500   0.9154   0.07244   0.06767  -0.0284   0.0068   1.0000
   9.750   0.8977   0.07718   0.07258  -0.0280   0.0069   1.0000
  10.000   0.8789   0.08281   0.07835  -0.0293   0.0071   1.0000
  10.250   0.8606   0.08973   0.08537  -0.0326   0.0073   1.0000
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