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Eh 1.0/7.0 (eh1070-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: Eh 1.0/7.0 (eh1070-il)
Reynolds number: 50,000
Max Cl/Cd: 30.27 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-eh1070-il-50000-n5.txt
Download as CSV file: xf-eh1070-il-50000-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 =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.5100   0.10412   0.09787   0.0082   1.0000   0.0532
  -9.000  -0.5986   0.10468   0.09805   0.0098   1.0000   0.0436
  -8.750  -0.5958   0.10039   0.09380   0.0083   1.0000   0.0421
  -8.500  -0.5953   0.09599   0.08947   0.0058   1.0000   0.0415
  -8.250  -0.5954   0.09156   0.08510   0.0032   1.0000   0.0403
  -8.000  -0.5978   0.08691   0.08053  -0.0002   1.0000   0.0397
  -7.750  -0.5997   0.08222   0.07586  -0.0033   1.0000   0.0387
  -7.500  -0.5994   0.07742   0.07105  -0.0062   1.0000   0.0382
  -7.250  -0.5976   0.07256   0.06613  -0.0087   1.0000   0.0377
  -7.000  -0.5937   0.06791   0.06137  -0.0106   1.0000   0.0378
  -6.750  -0.5874   0.06343   0.05674  -0.0120   1.0000   0.0381
  -6.500  -0.5789   0.05908   0.05218  -0.0130   1.0000   0.0390
  -6.250  -0.5682   0.05472   0.04754  -0.0137   1.0000   0.0394
  -6.000  -0.5553   0.05041   0.04288  -0.0139   1.0000   0.0393
  -5.750  -0.5401   0.04614   0.03805  -0.0137   1.0000   0.0387
  -5.500  -0.5227   0.04211   0.03355  -0.0133   1.0000   0.0383
  -5.250  -0.5030   0.03845   0.02937  -0.0125   1.0000   0.0382
  -5.000  -0.4814   0.03507   0.02545  -0.0116   1.0000   0.0386
  -4.750  -0.4578   0.03207   0.02194  -0.0107   1.0000   0.0395
  -4.500  -0.4327   0.02941   0.01880  -0.0097   1.0000   0.0409
  -4.250  -0.4064   0.02708   0.01604  -0.0087   1.0000   0.0430
  -4.000  -0.3806   0.02516   0.01389  -0.0078   1.0000   0.0483
  -3.750  -0.3548   0.02380   0.01225  -0.0069   1.0000   0.0586
  -3.500  -0.3297   0.02214   0.01055  -0.0060   1.0000   0.0680
  -3.250  -0.3049   0.02068   0.00910  -0.0053   1.0000   0.0949
  -3.000  -0.2840   0.01876   0.00761  -0.0043   1.0000   0.1754
  -2.750  -0.2763   0.01590   0.00691  -0.0009   1.0000   0.5345
  -2.500  -0.1741   0.01559   0.00678  -0.0091   1.0000   0.9703
  -2.250  -0.1153   0.01527   0.00588  -0.0156   1.0000   1.0000
  -2.000  -0.0958   0.01507   0.00545  -0.0146   1.0000   1.0000
  -1.750  -0.0763   0.01491   0.00501  -0.0135   1.0000   1.0000
  -1.500  -0.0572   0.01479   0.00472  -0.0124   1.0000   1.0000
  -1.250  -0.0384   0.01469   0.00448  -0.0112   1.0000   1.0000
  -1.000  -0.0201   0.01463   0.00431  -0.0098   1.0000   1.0000
  -0.750  -0.0026   0.01459   0.00419  -0.0084   1.0000   1.0000
  -0.500   0.0138   0.01459   0.00413  -0.0068   1.0000   1.0000
  -0.250   0.0288   0.01464   0.00411  -0.0050   1.0000   1.0000
   0.000   0.0422   0.01474   0.00418  -0.0030   1.0000   1.0000
   0.250   0.0538   0.01492   0.00433  -0.0009   1.0000   1.0000
   0.500   0.1034   0.01510   0.00449  -0.0060   0.9726   1.0000
   0.750   0.1538   0.01525   0.00466  -0.0110   0.9445   1.0000
   1.000   0.2010   0.01537   0.00483  -0.0150   0.9148   1.0000
   1.250   0.2416   0.01550   0.00497  -0.0174   0.8824   1.0000
   1.500   0.2766   0.01563   0.00512  -0.0184   0.8502   1.0000
   1.750   0.3055   0.01581   0.00529  -0.0182   0.8184   1.0000
   2.000   0.3307   0.01602   0.00558  -0.0171   0.7873   1.0000
   2.250   0.3543   0.01626   0.00580  -0.0158   0.7579   1.0000
   2.500   0.3772   0.01652   0.00606  -0.0143   0.7297   1.0000
   2.750   0.3999   0.01680   0.00636  -0.0128   0.7022   1.0000
   3.000   0.4227   0.01710   0.00669  -0.0114   0.6752   1.0000
   3.250   0.4456   0.01741   0.00707  -0.0100   0.6485   1.0000
   3.500   0.4686   0.01775   0.00760  -0.0087   0.6218   1.0000
   3.750   0.4917   0.01810   0.00805  -0.0074   0.5949   1.0000
   4.000   0.5149   0.01847   0.00854  -0.0061   0.5678   1.0000
   4.250   0.5381   0.01886   0.00908  -0.0049   0.5404   1.0000
   4.500   0.5611   0.01926   0.00964  -0.0036   0.5111   1.0000
   4.750   0.5842   0.01966   0.01026  -0.0023   0.4777   1.0000
   5.000   0.6058   0.02002   0.01078  -0.0006   0.4352   1.0000
   5.250   0.6188   0.02044   0.01062   0.0021   0.3050   1.0000
   5.500   0.6320   0.02296   0.01201   0.0033   0.1061   1.0000
   5.750   0.6484   0.02553   0.01411   0.0042   0.0608   1.0000
   6.000   0.6669   0.02744   0.01618   0.0055   0.0507   1.0000
   6.250   0.6857   0.02939   0.01836   0.0071   0.0458   1.0000
   6.500   0.7054   0.03127   0.02044   0.0085   0.0404   1.0000
   6.750   0.7247   0.03347   0.02281   0.0096   0.0349   1.0000
   7.000   0.7462   0.03582   0.02552   0.0109   0.0321   1.0000
   7.250   0.7667   0.03864   0.02874   0.0120   0.0307   1.0000
   7.500   0.7848   0.04182   0.03236   0.0131   0.0298   1.0000
   7.750   0.7995   0.04534   0.03634   0.0141   0.0292   1.0000
   8.000   0.8103   0.04916   0.04063   0.0149   0.0289   1.0000
   8.250   0.8165   0.05325   0.04519   0.0156   0.0289   1.0000
   8.500   0.8177   0.05755   0.04991   0.0159   0.0289   1.0000
   8.750   0.8135   0.06201   0.05471   0.0158   0.0292   1.0000
   9.000   0.8040   0.06658   0.05955   0.0151   0.0295   1.0000
   9.250   0.7887   0.07112   0.06425   0.0139   0.0298   1.0000
   9.500   0.7718   0.07641   0.06962   0.0105   0.0302   1.0000
   9.750   0.7561   0.08304   0.07629   0.0052   0.0307   1.0000
  10.000   0.7444   0.09023   0.08348  -0.0003   0.0313   1.0000
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