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EPPLER E850 AIRFOIL (e850-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: EPPLER E850 AIRFOIL (e850-il)
Reynolds number: 500,000
Max Cl/Cd: 97.47 at α=2.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e850-il-500000.txt
Download as CSV file: xf-e850-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER E850 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4439   0.08243   0.08037  -0.0314   1.0000   0.0124
  -8.750  -0.4509   0.07837   0.07634  -0.0315   1.0000   0.0125
  -8.500  -0.4579   0.07485   0.07285  -0.0312   1.0000   0.0128
  -8.250  -0.4684   0.07064   0.06867  -0.0314   1.0000   0.0127
  -8.000  -0.4816   0.06628   0.06434  -0.0318   1.0000   0.0127
  -7.750  -0.5046   0.05978   0.05786  -0.0350   1.0000   0.0122
  -7.500  -0.5307   0.05718   0.05523  -0.0345   1.0000   0.0118
  -7.250  -0.5313   0.05034   0.04824  -0.0419   0.9967   0.0119
  -7.000  -0.5258   0.04384   0.04156  -0.0472   0.9940   0.0122
  -6.750  -0.5178   0.03902   0.03660  -0.0499   0.9909   0.0127
  -6.500  -0.5056   0.03468   0.03209  -0.0521   0.9879   0.0134
  -6.250  -0.4890   0.03041   0.02761  -0.0543   0.9854   0.0143
  -6.000  -0.4679   0.02629   0.02324  -0.0566   0.9836   0.0155
  -5.750  -0.4432   0.02226   0.01890  -0.0586   0.9822   0.0172
  -5.500  -0.4142   0.01903   0.01532  -0.0600   0.9811   0.0199
  -5.250  -0.3903   0.01998   0.01595  -0.0583   0.9773   0.0220
  -5.000  -0.3733   0.01315   0.00853  -0.0594   0.9751   0.0238
  -4.750  -0.3474   0.01059   0.00582  -0.0602   0.9736   0.0260
  -4.500  -0.3186   0.00871   0.00371  -0.0607   0.9723   0.0284
  -4.250  -0.3122   0.01860   0.01305  -0.0588   0.9737   0.0195
  -4.000  -0.2815   0.01581   0.00994  -0.0578   0.9731   0.0096
  -3.750  -0.2524   0.01387   0.00781  -0.0577   0.9722   0.0080
  -3.500  -0.2214   0.01259   0.00635  -0.0583   0.9712   0.0070
  -3.250  -0.1886   0.01182   0.00539  -0.0593   0.9702   0.0065
  -3.000  -0.1550   0.01144   0.00484  -0.0605   0.9693   0.0065
  -2.750  -0.1229   0.01125   0.00448  -0.0614   0.9679   0.0076
  -2.500  -0.1004   0.01109   0.00421  -0.0603   0.9640   0.0093
  -2.250  -0.0724   0.00975   0.00388  -0.0615   0.9622   0.2791
  -2.000  -0.0430   0.00906   0.00390  -0.0626   0.9605   0.4811
  -1.750  -0.0121   0.00886   0.00412  -0.0634   0.9589   0.5970
  -1.500   0.0213   0.00885   0.00416  -0.0647   0.9576   0.6261
  -1.250   0.0559   0.00882   0.00416  -0.0663   0.9565   0.6391
  -1.000   0.0829   0.00878   0.00413  -0.0662   0.9531   0.6489
  -0.750   0.1117   0.00871   0.00409  -0.0665   0.9495   0.6588
  -0.500   0.1459   0.00861   0.00402  -0.0680   0.9473   0.6694
  -0.250   0.1820   0.00850   0.00395  -0.0698   0.9455   0.6807
   0.000   0.2202   0.00835   0.00386  -0.0721   0.9441   0.6927
   0.250   0.2610   0.00815   0.00373  -0.0749   0.9430   0.7054
   0.500   0.2851   0.00802   0.00368  -0.0740   0.9363   0.7187
   0.750   0.3268   0.00771   0.00348  -0.0769   0.9339   0.7332
   1.000   0.3751   0.00726   0.00315  -0.0810   0.9317   0.7486
   1.250   0.4116   0.00685   0.00286  -0.0824   0.9238   0.7656
   1.500   0.4507   0.00637   0.00258  -0.0842   0.9124   0.7840
   1.750   0.4820   0.00611   0.00246  -0.0846   0.9002   0.8044
   2.000   0.5097   0.00589   0.00237  -0.0840   0.8812   0.8277
   2.250   0.5350   0.00571   0.00226  -0.0829   0.8435   0.8551
   2.500   0.5575   0.00572   0.00217  -0.0810   0.7692   0.8891
   2.750   0.5631   0.00660   0.00232  -0.0757   0.5790   0.9482
   3.000   0.5738   0.00838   0.00283  -0.0730   0.2924   1.0000
   3.250   0.5880   0.00979   0.00335  -0.0709   0.0947   1.0000
   3.500   0.6058   0.01117   0.00444  -0.0688   0.0132   1.0000
   3.750   0.6279   0.01208   0.00549  -0.0673   0.0113   1.0000
   4.000   0.6495   0.01314   0.00665  -0.0658   0.0104   1.0000
   4.250   0.6731   0.01382   0.00734  -0.0651   0.0069   1.0000
   4.500   0.6940   0.01680   0.01056  -0.0633   0.0057   1.0000
   4.750   0.7201   0.01750   0.01137  -0.0628   0.0051   1.0000
   5.000   0.7460   0.01954   0.01366  -0.0618   0.0047   1.0000
   5.250   0.7698   0.02258   0.01707  -0.0602   0.0047   1.0000
   5.500   0.7878   0.02692   0.02191  -0.0575   0.0052   1.0000
   5.750   0.8125   0.03169   0.02718  -0.0539   0.0097   1.0000
   6.000   0.8276   0.03545   0.03129  -0.0511   0.0097   1.0000
   6.250   0.8410   0.03903   0.03518  -0.0484   0.0094   1.0000
   6.500   0.8523   0.04276   0.03919  -0.0458   0.0090   1.0000
   6.750   0.8617   0.04644   0.04313  -0.0433   0.0085   1.0000
   7.000   0.8691   0.05016   0.04707  -0.0409   0.0082   1.0000
   7.250   0.8741   0.05394   0.05106  -0.0387   0.0079   1.0000
   7.500   0.8766   0.05774   0.05503  -0.0366   0.0076   1.0000
   7.750   0.8765   0.06160   0.05906  -0.0345   0.0074   1.0000
   8.000   0.8731   0.06556   0.06316  -0.0326   0.0072   1.0000
   8.250   0.8670   0.06941   0.06713  -0.0308   0.0071   1.0000
   8.500   0.8546   0.07307   0.07089  -0.0285   0.0071   1.0000
   8.750   0.8388   0.07703   0.07493  -0.0268   0.0071   1.0000
   9.000   0.8225   0.08179   0.07977  -0.0272   0.0072   1.0000
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