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EH 0.0/9.0 (eh0009-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: EH 0.0/9.0 (eh0009-il)
Reynolds number: 200,000
Max Cl/Cd: 47.7 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-eh0009-il-200000.txt
Download as CSV file: xf-eh0009-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EH 0.0/9.0                                      
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.6738   0.08718   0.08386  -0.0061   1.0000   0.0574
  -9.500  -0.6850   0.08025   0.07698  -0.0115   1.0000   0.0582
  -9.250  -0.7791   0.05756   0.05377  -0.0239   1.0000   0.0266
  -9.000  -0.7904   0.05390   0.04980  -0.0227   1.0000   0.0277
  -8.750  -0.7986   0.04896   0.04464  -0.0216   1.0000   0.0265
  -8.500  -0.8016   0.04433   0.03966  -0.0201   1.0000   0.0257
  -8.250  -0.7976   0.04060   0.03552  -0.0184   1.0000   0.0264
  -8.000  -0.7915   0.03635   0.03081  -0.0165   1.0000   0.0259
  -7.750  -0.7815   0.03212   0.02607  -0.0144   1.0000   0.0250
  -7.500  -0.7665   0.02869   0.02217  -0.0126   1.0000   0.0249
  -7.250  -0.7478   0.02594   0.01902  -0.0110   1.0000   0.0255
  -7.000  -0.7269   0.02373   0.01649  -0.0096   1.0000   0.0267
  -6.750  -0.7044   0.02267   0.01512  -0.0085   1.0000   0.0291
  -6.500  -0.6853   0.01989   0.01226  -0.0072   1.0000   0.0323
  -6.250  -0.6642   0.01853   0.01084  -0.0059   1.0000   0.0357
  -6.000  -0.6426   0.01745   0.00962  -0.0045   1.0000   0.0398
  -5.750  -0.6243   0.01608   0.00829  -0.0029   1.0000   0.0486
  -5.500  -0.6050   0.01497   0.00718  -0.0014   1.0000   0.0619
  -5.250  -0.5872   0.01369   0.00613   0.0003   1.0000   0.0985
  -5.000  -0.5692   0.01257   0.00540   0.0017   1.0000   0.1701
  -4.750  -0.5497   0.01181   0.00492   0.0029   1.0000   0.2399
  -4.500  -0.5296   0.01121   0.00458   0.0041   1.0000   0.3032
  -4.250  -0.5092   0.01071   0.00429   0.0054   1.0000   0.3637
  -4.000  -0.4892   0.01025   0.00407   0.0067   1.0000   0.4258
  -3.750  -0.4689   0.00989   0.00389   0.0082   1.0000   0.4795
  -3.500  -0.4483   0.00960   0.00373   0.0096   1.0000   0.5280
  -3.250  -0.4279   0.00934   0.00359   0.0111   1.0000   0.5731
  -3.000  -0.4076   0.00912   0.00352   0.0127   1.0000   0.6170
  -2.750  -0.3875   0.00895   0.00347   0.0143   1.0000   0.6580
  -2.500  -0.3676   0.00883   0.00346   0.0160   1.0000   0.6967
  -2.250  -0.3480   0.00876   0.00349   0.0177   1.0000   0.7344
  -2.000  -0.3291   0.00871   0.00355   0.0197   1.0000   0.7692
  -1.750  -0.3108   0.00871   0.00361   0.0217   1.0000   0.8034
  -1.500  -0.2915   0.00875   0.00373   0.0235   0.9995   0.8362
  -1.250  -0.2494   0.00888   0.00391   0.0208   0.9911   0.8690
  -1.000  -0.2055   0.00902   0.00405   0.0179   0.9833   0.8971
  -0.750  -0.1599   0.00914   0.00416   0.0146   0.9755   0.9188
  -0.500  -0.1076   0.00927   0.00427   0.0100   0.9697   0.9355
  -0.250  -0.0555   0.00935   0.00434   0.0053   0.9628   0.9479
   0.000   0.0000   0.00938   0.00436   0.0000   0.9568   0.9568
   0.250   0.0555   0.00935   0.00434  -0.0053   0.9479   0.9629
   0.500   0.1077   0.00927   0.00427  -0.0100   0.9356   0.9698
   0.750   0.1599   0.00914   0.00416  -0.0146   0.9190   0.9756
   1.000   0.2055   0.00902   0.00404  -0.0179   0.8970   0.9833
   1.250   0.2494   0.00888   0.00390  -0.0208   0.8692   0.9912
   1.500   0.2915   0.00874   0.00373  -0.0235   0.8365   0.9996
   1.750   0.3106   0.00871   0.00361  -0.0216   0.8035   1.0000
   2.000   0.3288   0.00871   0.00355  -0.0196   0.7692   1.0000
   2.250   0.3477   0.00875   0.00348  -0.0177   0.7344   1.0000
   2.500   0.3673   0.00883   0.00346  -0.0159   0.6969   1.0000
   2.750   0.3872   0.00895   0.00347  -0.0142   0.6582   1.0000
   3.000   0.4073   0.00912   0.00351  -0.0126   0.6172   1.0000
   3.250   0.4276   0.00933   0.00359  -0.0110   0.5734   1.0000
   3.500   0.4480   0.00960   0.00373  -0.0095   0.5282   1.0000
   3.750   0.4686   0.00989   0.00389  -0.0081   0.4796   1.0000
   4.000   0.4889   0.01025   0.00407  -0.0067   0.4262   1.0000
   4.250   0.5089   0.01070   0.00429  -0.0053   0.3643   1.0000
   4.500   0.5293   0.01120   0.00457  -0.0040   0.3035   1.0000
   4.750   0.5494   0.01181   0.00492  -0.0028   0.2399   1.0000
   5.000   0.5689   0.01257   0.00540  -0.0016   0.1704   1.0000
   5.250   0.5870   0.01369   0.00613  -0.0002   0.0985   1.0000
   5.500   0.6049   0.01497   0.00718   0.0014   0.0621   1.0000
   5.750   0.6242   0.01607   0.00828   0.0030   0.0486   1.0000
   6.000   0.6426   0.01744   0.00962   0.0045   0.0398   1.0000
   6.250   0.6642   0.01852   0.01083   0.0059   0.0356   1.0000
   6.500   0.6854   0.01988   0.01225   0.0072   0.0323   1.0000
   6.750   0.7043   0.02275   0.01520   0.0085   0.0291   1.0000
   7.000   0.7270   0.02374   0.01650   0.0096   0.0267   1.0000
   7.250   0.7480   0.02593   0.01900   0.0110   0.0255   1.0000
   7.500   0.7667   0.02870   0.02217   0.0126   0.0249   1.0000
   7.750   0.7816   0.03216   0.02612   0.0144   0.0250   1.0000
   8.000   0.7918   0.03630   0.03076   0.0164   0.0259   1.0000
   8.250   0.7972   0.04083   0.03574   0.0183   0.0268   1.0000
   8.500   0.8006   0.04474   0.04006   0.0200   0.0264   1.0000
   8.750   0.7982   0.04923   0.04489   0.0215   0.0269   1.0000
   9.000   0.7920   0.05336   0.04933   0.0228   0.0267   1.0000
   9.250   0.7796   0.05801   0.05415   0.0236   0.0276   1.0000
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