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Eh 1.0/7.0 (eh1070-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: Eh 1.0/7.0 (eh1070-il)
Reynolds number: 1,000,000
Max Cl/Cd: 66.61 at α=3.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-eh1070-il-1000000-n5.txt
Download as CSV file: xf-eh1070-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: Eh 1.0/7.0                                      
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.6240   0.08423   0.08272   0.0068   1.0000   0.0017
  -8.500  -0.6311   0.07913   0.07765   0.0035   1.0000   0.0017
  -8.250  -0.6421   0.07460   0.07315  -0.0005   1.0000   0.0017
  -8.000  -0.6495   0.06853   0.06704  -0.0049   1.0000   0.0017
  -7.750  -0.6515   0.06241   0.06085  -0.0083   1.0000   0.0016
  -7.500  -0.6529   0.05523   0.05353  -0.0109   1.0000   0.0016
  -7.000  -0.6902   0.01975   0.01604  -0.0072   1.0000   0.0020
  -6.750  -0.6683   0.01801   0.01404  -0.0064   1.0000   0.0021
  -6.500  -0.6444   0.01698   0.01283  -0.0059   1.0000   0.0022
  -6.250  -0.6201   0.01610   0.01181  -0.0054   1.0000   0.0023
  -6.000  -0.5990   0.01381   0.00916  -0.0043   1.0000   0.0026
  -5.750  -0.5746   0.01295   0.00817  -0.0038   1.0000   0.0028
  -5.500  -0.5495   0.01238   0.00752  -0.0034   1.0000   0.0030
  -5.250  -0.5245   0.01175   0.00682  -0.0029   1.0000   0.0032
  -5.000  -0.4994   0.01123   0.00623  -0.0024   1.0000   0.0034
  -4.750  -0.4690   0.01071   0.00561  -0.0031   0.9908   0.0039
  -4.500  -0.4352   0.01020   0.00504  -0.0045   0.9759   0.0046
  -4.250  -0.4019   0.00971   0.00446  -0.0057   0.9532   0.0052
  -4.000  -0.3740   0.00954   0.00419  -0.0056   0.9184   0.0055
  -3.750  -0.3515   0.00880   0.00320  -0.0043   0.8811   0.0067
  -3.500  -0.3270   0.00854   0.00276  -0.0034   0.8462   0.0074
  -3.250  -0.3015   0.00835   0.00243  -0.0028   0.8139   0.0081
  -3.000  -0.2754   0.00818   0.00211  -0.0024   0.7855   0.0086
  -2.750  -0.2489   0.00804   0.00179  -0.0020   0.7588   0.0089
  -2.500  -0.2221   0.00794   0.00156  -0.0017   0.7338   0.0095
  -2.250  -0.1952   0.00786   0.00137  -0.0014   0.7094   0.0112
  -2.000  -0.1685   0.00770   0.00120  -0.0012   0.6854   0.0296
  -1.750  -0.1418   0.00755   0.00107  -0.0010   0.6626   0.0583
  -1.500  -0.1148   0.00744   0.00094  -0.0009   0.6415   0.0817
  -1.250  -0.0881   0.00727   0.00084  -0.0008   0.6213   0.1241
  -1.000  -0.0621   0.00693   0.00075  -0.0006   0.6021   0.2180
  -0.750  -0.0382   0.00617   0.00063  -0.0003   0.5838   0.4285
  -0.250   0.0115   0.00538   0.00053   0.0005   0.5484   0.6789
   0.000   0.0375   0.00524   0.00051   0.0009   0.5314   0.7390
   0.250   0.0632   0.00513   0.00053   0.0014   0.5150   0.7932
   0.500   0.0890   0.00509   0.00055   0.0020   0.4971   0.8372
   0.750   0.1142   0.00505   0.00059   0.0027   0.4776   0.8752
   1.000   0.1384   0.00505   0.00065   0.0037   0.4596   0.9185
   1.250   0.1647   0.00510   0.00070   0.0043   0.4448   0.9440
   1.500   0.1931   0.00517   0.00074   0.0042   0.4316   0.9559
   1.750   0.2229   0.00525   0.00078   0.0038   0.4166   0.9643
   2.000   0.2532   0.00536   0.00083   0.0033   0.3980   0.9707
   2.250   0.2844   0.00549   0.00091   0.0025   0.3767   0.9758
   2.500   0.3139   0.00562   0.00097   0.0020   0.3568   0.9815
   2.750   0.3469   0.00576   0.00105   0.0008   0.3353   0.9844
   3.000   0.3790   0.00593   0.00115  -0.0003   0.3111   0.9874
   3.250   0.4103   0.00616   0.00126  -0.0012   0.2763   0.9906
   3.500   0.4396   0.00663   0.00147  -0.0019   0.2067   0.9937
   3.750   0.4706   0.00711   0.00169  -0.0030   0.1423   0.9959
   4.000   0.5006   0.00783   0.00204  -0.0040   0.0579   0.9985
   4.250   0.5293   0.00833   0.00236  -0.0045   0.0188   1.0000
   4.500   0.5534   0.00861   0.00262  -0.0038   0.0109   1.0000
   4.750   0.5778   0.00888   0.00291  -0.0031   0.0071   1.0000
   5.000   0.6024   0.00913   0.00319  -0.0026   0.0057   1.0000
   5.250   0.6269   0.00944   0.00357  -0.0020   0.0048   1.0000
   5.500   0.6511   0.00986   0.00404  -0.0013   0.0039   1.0000
   5.750   0.6750   0.01040   0.00468  -0.0006   0.0032   1.0000
   6.000   0.6995   0.01081   0.00515   0.0000   0.0029   1.0000
   6.250   0.7232   0.01145   0.00590   0.0007   0.0026   1.0000
   6.500   0.7462   0.01227   0.00685   0.0015   0.0025   1.0000
   6.750   0.7683   0.01328   0.00800   0.0024   0.0024   1.0000
   7.000   0.7899   0.01446   0.00934   0.0034   0.0023   1.0000
   7.250   0.8106   0.01591   0.01096   0.0045   0.0023   1.0000
   7.500   0.8309   0.01757   0.01285   0.0056   0.0023   1.0000
   7.750   0.8505   0.01952   0.01504   0.0068   0.0023   1.0000
   8.000   0.8702   0.02128   0.01701   0.0077   0.0022   1.0000
   8.250   0.8899   0.02281   0.01873   0.0086   0.0020   1.0000
   8.500   0.9080   0.02459   0.02073   0.0095   0.0019   1.0000
   8.750   0.9256   0.02623   0.02257   0.0104   0.0018   1.0000
   9.000   0.9372   0.02911   0.02576   0.0117   0.0018   1.0000
   9.250   0.9451   0.03231   0.02927   0.0132   0.0017   1.0000
   9.500   0.9490   0.03568   0.03293   0.0146   0.0017   1.0000
   9.750   0.9412   0.04033   0.03791   0.0165   0.0017   1.0000
  10.000   0.9342   0.04367   0.04145   0.0180   0.0016   1.0000
  10.250   0.9109   0.04743   0.04539   0.0201   0.0016   1.0000
  10.500   0.8998   0.05063   0.04869   0.0189   0.0016   1.0000
  10.750   0.8716   0.05828   0.05652   0.0133   0.0017   1.0000
  11.000   0.8459   0.06814   0.06653   0.0050   0.0017   1.0000
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