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EPPLER E851 AIRFOIL (e851-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER E851 AIRFOIL (e851-il)
Reynolds number: 50,000
Max Cl/Cd: 27.26 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e851-il-50000.txt
Download as CSV file: xf-e851-il-50000.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 =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4440   0.09871   0.09221  -0.0204   1.0000   0.2969
  -7.750  -0.4437   0.09590   0.08948  -0.0188   1.0000   0.3107
  -7.500  -0.4462   0.09355   0.08707  -0.0167   1.0000   0.3286
  -7.250  -0.4566   0.09189   0.08555  -0.0144   1.0000   0.3441
  -7.000  -0.4575   0.08974   0.08349  -0.0110   1.0000   0.3710
  -5.750  -0.5365   0.05569   0.04882  -0.0464   1.0000   0.1184
  -5.500  -0.5199   0.05016   0.04285  -0.0474   1.0000   0.1018
  -5.250  -0.4994   0.04529   0.03726  -0.0484   1.0000   0.0905
  -5.000  -0.4769   0.04094   0.03226  -0.0488   1.0000   0.0841
  -4.750  -0.4506   0.03779   0.02820  -0.0485   1.0000   0.0802
  -4.500  -0.4255   0.03462   0.02454  -0.0481   1.0000   0.0813
  -4.250  -0.4014   0.03233   0.02189  -0.0477   1.0000   0.0937
  -4.000  -0.3759   0.02946   0.01881  -0.0467   1.0000   0.1031
  -3.750  -0.3499   0.02707   0.01629  -0.0450   1.0000   0.1167
  -3.500  -0.3265   0.02499   0.01437  -0.0437   1.0000   0.1573
  -3.250  -0.2980   0.02090   0.01199  -0.0441   1.0000   0.3376
  -3.000  -0.3072   0.02136   0.01378  -0.0318   1.0000   0.6781
  -2.750  -0.3113   0.02171   0.01417  -0.0221   1.0000   0.7474
  -2.500  -0.3148   0.02169   0.01411  -0.0131   1.0000   0.8024
  -2.250  -0.3190   0.02140   0.01380  -0.0043   1.0000   0.8565
  -2.000  -0.2964   0.02127   0.01347   0.0007   1.0000   0.9371
  -1.750  -0.1643   0.02169   0.01285  -0.0195   1.0000   0.9900
  -1.500  -0.1490   0.02105   0.01191  -0.0204   1.0000   1.0000
  -1.250  -0.1557   0.02021   0.01100  -0.0172   1.0000   1.0000
  -1.000  -0.1442   0.01971   0.01030  -0.0171   1.0000   1.0000
  -0.750  -0.1189   0.01956   0.00988  -0.0191   1.0000   1.0000
  -0.500  -0.0901   0.01961   0.00964  -0.0214   1.0000   1.0000
  -0.250  -0.0613   0.01976   0.00954  -0.0233   1.0000   1.0000
   0.000  -0.0336   0.01999   0.00954  -0.0248   1.0000   1.0000
   0.250  -0.0070   0.02026   0.00954  -0.0259   1.0000   1.0000
   0.500   0.0185   0.02057   0.00968  -0.0266   1.0000   1.0000
   0.750   0.0432   0.02091   0.00988  -0.0272   1.0000   1.0000
   1.000   0.0672   0.02128   0.01014  -0.0276   1.0000   1.0000
   1.250   0.0906   0.02168   0.01046  -0.0278   1.0000   1.0000
   1.500   0.1135   0.02211   0.01084  -0.0280   1.0000   1.0000
   1.750   0.1359   0.02257   0.01127  -0.0281   1.0000   1.0000
   2.000   0.1580   0.02306   0.01175  -0.0282   1.0000   1.0000
   2.250   0.1796   0.02359   0.01228  -0.0282   1.0000   1.0000
   2.500   0.2008   0.02416   0.01288  -0.0282   1.0000   1.0000
   2.750   0.2216   0.02477   0.01353  -0.0282   1.0000   1.0000
   3.000   0.2419   0.02543   0.01426  -0.0282   1.0000   1.0000
   3.250   0.2618   0.02614   0.01506  -0.0282   1.0000   1.0000
   3.500   0.2812   0.02690   0.01604  -0.0281   1.0000   1.0000
   3.750   0.3000   0.02773   0.01699  -0.0281   1.0000   1.0000
   4.000   0.3182   0.02863   0.01803  -0.0282   1.0000   1.0000
   4.250   0.3357   0.02961   0.01917  -0.0282   1.0000   1.0000
   4.500   0.3525   0.03068   0.02043  -0.0284   1.0000   1.0000
   4.750   0.3880   0.03239   0.02245  -0.0323   0.9887   1.0000
   5.000   0.4684   0.03463   0.02546  -0.0436   0.9475   1.0000
   5.250   0.6115   0.03342   0.02558  -0.0587   0.8602   1.0000
   5.500   0.7450   0.02733   0.01649  -0.0507   0.1628   1.0000
   5.750   0.8093   0.03188   0.02077  -0.0545   0.1088   1.0000
   6.000   0.8507   0.03592   0.02516  -0.0552   0.0900   1.0000
   6.250   0.8827   0.04023   0.02988  -0.0546   0.0853   1.0000
   6.500   0.9027   0.04314   0.03327  -0.0526   0.0773   1.0000
   6.750   0.9207   0.04710   0.03753  -0.0511   0.0726   1.0000
   7.000   0.9384   0.05162   0.04244  -0.0492   0.0736   1.0000
   7.250   0.9459   0.05508   0.04688  -0.0453   0.0789   1.0000
   7.500   0.9537   0.05988   0.05213  -0.0428   0.0839   1.0000
   7.750   0.9558   0.06458   0.05746  -0.0398   0.0928   1.0000
   8.000   0.9553   0.06983   0.06318  -0.0375   0.1049   1.0000
   8.250   0.9486   0.07574   0.06950  -0.0359   0.1220   1.0000
   8.500   0.8783   0.07503   0.06941  -0.0289   0.1272   1.0000
   8.750   0.8458   0.07909   0.07364  -0.0266   0.1286   1.0000
   9.000   0.8149   0.08386   0.07850  -0.0261   0.1297   1.0000
   9.250   0.7847   0.08968   0.08436  -0.0279   0.1313   1.0000
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