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EH 1.0/9.0 (eh1090-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: EH 1.0/9.0 (eh1090-il)
Reynolds number: 500,000
Max Cl/Cd: 74.39 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-eh1090-il-500000.txt
Download as CSV file: xf-eh1090-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EH 1.0/9.0                                      
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.6529   0.05092   0.04814  -0.0191   1.0000   0.0213
  -7.500  -0.6512   0.04674   0.04368  -0.0180   1.0000   0.0213
  -7.250  -0.6814   0.03090   0.02679  -0.0137   1.0000   0.0141
  -7.000  -0.6703   0.02656   0.02212  -0.0119   1.0000   0.0131
  -6.750  -0.6538   0.02385   0.01904  -0.0103   1.0000   0.0132
  -6.500  -0.6342   0.02192   0.01683  -0.0089   1.0000   0.0134
  -6.250  -0.6139   0.02025   0.01484  -0.0076   1.0000   0.0141
  -6.000  -0.5932   0.01827   0.01262  -0.0063   1.0000   0.0141
  -5.750  -0.5722   0.01647   0.01063  -0.0049   1.0000   0.0141
  -5.500  -0.5524   0.01467   0.00866  -0.0033   1.0000   0.0144
  -5.250  -0.5337   0.01317   0.00703  -0.0015   1.0000   0.0153
  -5.000  -0.5132   0.01236   0.00616  -0.0001   1.0000   0.0165
  -4.750  -0.4922   0.01177   0.00554   0.0013   1.0000   0.0179
  -4.500  -0.4716   0.01124   0.00495   0.0027   1.0000   0.0195
  -4.250  -0.4498   0.01089   0.00453   0.0039   0.9997   0.0213
  -4.000  -0.4120   0.00993   0.00359   0.0017   0.9943   0.0372
  -3.750  -0.3752   0.00934   0.00317  -0.0004   0.9874   0.0731
  -3.500  -0.3367   0.00892   0.00289  -0.0029   0.9810   0.1100
  -3.250  -0.3017   0.00840   0.00262  -0.0047   0.9703   0.1702
  -3.000  -0.2690   0.00781   0.00237  -0.0061   0.9564   0.2590
  -2.750  -0.2405   0.00713   0.00213  -0.0065   0.9376   0.3777
  -2.500  -0.2172   0.00644   0.00194  -0.0056   0.9151   0.5194
  -2.250  -0.1960   0.00597   0.00183  -0.0040   0.8907   0.6352
  -2.000  -0.1736   0.00574   0.00175  -0.0025   0.8667   0.7093
  -1.750  -0.1500   0.00563   0.00170  -0.0012   0.8431   0.7599
  -1.500  -0.1263   0.00557   0.00168   0.0002   0.8194   0.8054
  -1.250  -0.1016   0.00558   0.00166   0.0013   0.7965   0.8351
  -1.000  -0.0765   0.00562   0.00162   0.0022   0.7744   0.8586
  -0.750  -0.0509   0.00566   0.00160   0.0030   0.7522   0.8761
  -0.500  -0.0254   0.00573   0.00160   0.0039   0.7311   0.8943
  -0.250  -0.0003   0.00581   0.00161   0.0049   0.7101   0.9137
   0.000   0.0259   0.00592   0.00164   0.0056   0.6901   0.9297
   0.250   0.0536   0.00603   0.00168   0.0060   0.6699   0.9432
   0.500   0.0831   0.00618   0.00173   0.0060   0.6505   0.9556
   0.750   0.1204   0.00634   0.00181   0.0043   0.6306   0.9635
   1.000   0.1566   0.00649   0.00186   0.0027   0.6109   0.9707
   1.250   0.1979   0.00662   0.00189  -0.0002   0.5913   0.9741
   1.500   0.2353   0.00673   0.00192  -0.0022   0.5709   0.9786
   1.750   0.2696   0.00684   0.00196  -0.0037   0.5522   0.9832
   2.000   0.3077   0.00692   0.00197  -0.0060   0.5325   0.9858
   2.250   0.3447   0.00701   0.00199  -0.0080   0.5134   0.9893
   2.500   0.3803   0.00712   0.00205  -0.0098   0.4946   0.9934
   2.750   0.4174   0.00720   0.00207  -0.0119   0.4729   0.9964
   3.000   0.4541   0.00730   0.00211  -0.0140   0.4503   0.9995
   3.250   0.4795   0.00739   0.00215  -0.0137   0.4324   1.0000
   3.500   0.5025   0.00748   0.00223  -0.0129   0.4164   1.0000
   3.750   0.5254   0.00760   0.00232  -0.0120   0.4009   1.0000
   4.000   0.5483   0.00773   0.00242  -0.0111   0.3836   1.0000
   4.250   0.5713   0.00788   0.00254  -0.0103   0.3636   1.0000
   4.500   0.5940   0.00807   0.00268  -0.0094   0.3441   1.0000
   4.750   0.6167   0.00829   0.00285  -0.0085   0.3168   1.0000
   5.000   0.6386   0.00863   0.00303  -0.0075   0.2730   1.0000
   5.250   0.6605   0.00901   0.00325  -0.0066   0.2344   1.0000
   5.500   0.6825   0.00940   0.00352  -0.0057   0.1979   1.0000
   5.750   0.7038   0.00991   0.00385  -0.0047   0.1551   1.0000
   6.000   0.7236   0.01068   0.00435  -0.0036   0.0979   1.0000
   6.250   0.7395   0.01206   0.00531  -0.0018   0.0260   1.0000
   6.500   0.7592   0.01296   0.00627  -0.0004   0.0185   1.0000
   6.750   0.7813   0.01349   0.00688   0.0006   0.0170   1.0000
   7.000   0.8026   0.01415   0.00764   0.0017   0.0156   1.0000
   7.250   0.8232   0.01492   0.00850   0.0029   0.0143   1.0000
   7.500   0.8419   0.01596   0.00961   0.0042   0.0129   1.0000
   7.750   0.8550   0.01788   0.01169   0.0063   0.0116   1.0000
   8.000   0.8765   0.01858   0.01251   0.0073   0.0110   1.0000
   8.250   0.8960   0.01964   0.01368   0.0084   0.0104   1.0000
   8.500   0.9144   0.02095   0.01513   0.0097   0.0100   1.0000
   8.750   0.9322   0.02245   0.01677   0.0111   0.0095   1.0000
   9.000   0.9493   0.02407   0.01854   0.0124   0.0091   1.0000
   9.250   0.9648   0.02611   0.02079   0.0138   0.0089   1.0000
   9.500   0.9776   0.02869   0.02364   0.0155   0.0090   1.0000
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