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Eh 1.0/7.0 (eh1070-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: Eh 1.0/7.0 (eh1070-il)
Reynolds number: 500,000
Max Cl/Cd: 61.46 at α=3.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-eh1070-il-500000-n5.txt
Download as CSV file: xf-eh1070-il-500000-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 =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.6128   0.08428   0.08218   0.0065   1.0000   0.0033
  -8.250  -0.6189   0.07935   0.07728   0.0031   1.0000   0.0032
  -8.000  -0.6299   0.07424   0.07220  -0.0012   1.0000   0.0032
  -7.750  -0.6343   0.06801   0.06592  -0.0056   1.0000   0.0030
  -7.500  -0.6352   0.06172   0.05954  -0.0089   1.0000   0.0030
  -7.250  -0.6332   0.05514   0.05282  -0.0111   1.0000   0.0030
  -7.000  -0.6282   0.04844   0.04590  -0.0123   1.0000   0.0030
  -6.750  -0.6206   0.04153   0.03869  -0.0124   1.0000   0.0030
  -6.500  -0.6092   0.03545   0.03226  -0.0117   1.0000   0.0031
  -6.250  -0.5955   0.02975   0.02613  -0.0104   1.0000   0.0033
  -6.000  -0.5788   0.02512   0.02103  -0.0089   1.0000   0.0037
  -5.750  -0.5592   0.02170   0.01717  -0.0076   1.0000   0.0042
  -5.500  -0.5371   0.01938   0.01448  -0.0065   1.0000   0.0046
  -5.250  -0.5126   0.01842   0.01333  -0.0060   1.0000   0.0052
  -5.000  -0.4865   0.01835   0.01317  -0.0056   1.0000   0.0056
  -4.750  -0.4654   0.01515   0.00947  -0.0042   1.0000   0.0060
  -4.500  -0.4426   0.01330   0.00736  -0.0031   1.0000   0.0063
  -4.250  -0.4194   0.01210   0.00604  -0.0021   1.0000   0.0069
  -4.000  -0.3955   0.01140   0.00527  -0.0013   1.0000   0.0079
  -3.750  -0.3673   0.01081   0.00461  -0.0014   0.9952   0.0091
  -3.500  -0.3332   0.01034   0.00407  -0.0029   0.9821   0.0104
  -3.250  -0.2992   0.00986   0.00350  -0.0042   0.9649   0.0112
  -3.000  -0.2651   0.00938   0.00286  -0.0055   0.9411   0.0115
  -2.750  -0.2353   0.00902   0.00233  -0.0056   0.9095   0.0130
  -2.500  -0.2098   0.00879   0.00201  -0.0048   0.8752   0.0214
  -2.250  -0.1853   0.00853   0.00177  -0.0040   0.8438   0.0531
  -2.000  -0.1605   0.00830   0.00159  -0.0034   0.8140   0.0971
  -1.750  -0.1361   0.00793   0.00142  -0.0028   0.7859   0.1788
  -1.500  -0.1130   0.00725   0.00121  -0.0023   0.7593   0.3443
  -1.250  -0.0918   0.00641   0.00105  -0.0013   0.7337   0.5646
  -1.000  -0.0695   0.00597   0.00099  -0.0001   0.7092   0.6983
  -0.750  -0.0458   0.00579   0.00097   0.0010   0.6859   0.7742
  -0.500  -0.0209   0.00574   0.00096   0.0019   0.6638   0.8212
  -0.250   0.0043   0.00573   0.00095   0.0027   0.6427   0.8571
   0.000   0.0295   0.00574   0.00096   0.0036   0.6221   0.8927
   0.250   0.0567   0.00580   0.00101   0.0041   0.6018   0.9315
   0.500   0.0899   0.00591   0.00105   0.0033   0.5809   0.9575
   0.750   0.1217   0.00601   0.00107   0.0025   0.5612   0.9684
   1.000   0.1557   0.00611   0.00108   0.0011   0.5410   0.9731
   1.250   0.1877   0.00622   0.00109   0.0002   0.5195   0.9787
   1.500   0.2203   0.00633   0.00111  -0.0010   0.4978   0.9828
   1.750   0.2545   0.00643   0.00115  -0.0024   0.4785   0.9860
   2.000   0.2877   0.00653   0.00119  -0.0037   0.4612   0.9898
   2.250   0.3198   0.00664   0.00128  -0.0047   0.4427   0.9937
   2.500   0.3531   0.00677   0.00134  -0.0061   0.4197   0.9968
   2.750   0.3861   0.00691   0.00142  -0.0073   0.3961   0.9999
   3.000   0.4107   0.00706   0.00152  -0.0068   0.3750   1.0000
   3.250   0.4349   0.00724   0.00164  -0.0061   0.3513   1.0000
   3.500   0.4591   0.00747   0.00182  -0.0055   0.3182   1.0000
   3.750   0.4821   0.00794   0.00200  -0.0048   0.2476   1.0000
   4.000   0.5049   0.00850   0.00226  -0.0042   0.1771   1.0000
   4.250   0.5270   0.00923   0.00263  -0.0035   0.0957   1.0000
   4.500   0.5486   0.01010   0.00314  -0.0027   0.0230   1.0000
   4.750   0.5728   0.01050   0.00356  -0.0021   0.0131   1.0000
   5.000   0.5971   0.01096   0.00407  -0.0014   0.0098   1.0000
   5.250   0.6206   0.01161   0.00490  -0.0006   0.0072   1.0000
   5.500   0.6451   0.01202   0.00538   0.0000   0.0063   1.0000
   5.750   0.6690   0.01261   0.00607   0.0006   0.0055   1.0000
   6.000   0.6925   0.01327   0.00681   0.0013   0.0046   1.0000
   6.250   0.7154   0.01408   0.00772   0.0020   0.0041   1.0000
   6.500   0.7366   0.01530   0.00907   0.0030   0.0037   1.0000
   6.750   0.7560   0.01701   0.01097   0.0043   0.0035   1.0000
   7.000   0.7755   0.01894   0.01311   0.0055   0.0035   1.0000
   7.250   0.7946   0.02118   0.01561   0.0067   0.0034   1.0000
   7.500   0.8135   0.02343   0.01813   0.0078   0.0034   1.0000
   7.750   0.8312   0.02581   0.02087   0.0090   0.0034   1.0000
   8.000   0.8478   0.02820   0.02355   0.0101   0.0034   1.0000
   8.250   0.8627   0.03065   0.02629   0.0112   0.0035   1.0000
   8.500   0.8751   0.03332   0.02927   0.0123   0.0035   1.0000
   8.750   0.8847   0.03614   0.03239   0.0135   0.0036   1.0000
   9.000   0.8866   0.03979   0.03637   0.0148   0.0035   1.0000
   9.250   0.8767   0.04445   0.04136   0.0161   0.0035   1.0000
  10.250   0.8232   0.06088   0.05844   0.0112   0.0035   1.0000
  10.750   0.8033   0.07490   0.07268  -0.0005   0.0036   1.0000
  11.250   0.6618   0.10057   0.09854  -0.0129   0.0044   1.0000
  11.500   0.6489   0.10919   0.10714  -0.0163   0.0046   1.0000
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