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E62 (5.62%) (e62-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: E62 (5.62%) (e62-il)
Reynolds number: 500,000
Max Cl/Cd: 156.78 at α=2.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e62-il-500000.txt
Download as CSV file: xf-e62-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: E62  (5.62%)                                    
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.3660   0.12498   0.12273  -0.0234   1.0000   0.0089
  -9.500  -0.3648   0.12260   0.12037  -0.0232   1.0000   0.0089
  -9.250  -0.3643   0.12029   0.11809  -0.0229   1.0000   0.0089
  -7.500  -0.3000   0.09671   0.09456  -0.0387   0.9892   0.0091
  -7.250  -0.2842   0.09274   0.09058  -0.0412   0.9872   0.0093
  -7.000  -0.2666   0.08919   0.08702  -0.0445   0.9855   0.0096
  -6.750  -0.2554   0.08621   0.08406  -0.0468   0.9802   0.0098
  -6.500  -0.2348   0.08243   0.08028  -0.0517   0.9762   0.0101
  -6.250  -0.2079   0.07821   0.07605  -0.0585   0.9738   0.0105
  -6.000  -0.1872   0.07453   0.07236  -0.0636   0.9690   0.0108
  -5.750  -0.1629   0.07052   0.06834  -0.0697   0.9643   0.0113
  -5.500  -0.1300   0.06602   0.06380  -0.0781   0.9619   0.0121
  -5.250  -0.0758   0.06053   0.05826  -0.0930   0.9602   0.0139
  -5.000  -0.0482   0.05658   0.05427  -0.0989   0.9526   0.0141
  -4.750  -0.0063   0.04879   0.04641  -0.1113   0.9502   0.0148
  -4.500   0.0228   0.04627   0.04386  -0.1150   0.9488   0.0161
  -4.250   0.0683   0.04196   0.03946  -0.1239   0.9479   0.0174
  -4.000   0.1030   0.03792   0.03531  -0.1294   0.9413   0.0187
  -3.750   0.1580   0.03283   0.02997  -0.1385   0.9397   0.0211
  -3.500   0.2084   0.02930   0.02611  -0.1446   0.9385   0.0219
  -3.250   0.2526   0.02357   0.02013  -0.1525   0.9377   0.0245
  -3.000   0.2936   0.02143   0.01779  -0.1564   0.9368   0.0282
  -2.750   0.3402   0.01831   0.01414  -0.1610   0.9362   0.0344
  -2.500   0.3784   0.01683   0.01258  -0.1639   0.9354   0.0386
  -2.000   0.4568   0.01134   0.00607  -0.1671   0.9302   0.0185
  -1.750   0.4928   0.01023   0.00481  -0.1686   0.9277   0.0187
  -1.500   0.5293   0.00970   0.00422  -0.1703   0.9254   0.0212
  -1.250   0.5686   0.00893   0.00337  -0.1727   0.9236   0.0210
  -1.000   0.6042   0.00845   0.00281  -0.1742   0.9197   0.0215
  -0.750   0.6369   0.00808   0.00236  -0.1750   0.9131   0.0269
  -0.500   0.6770   0.00731   0.00216  -0.1780   0.9092   0.2376
  -0.250   0.7096   0.00657   0.00224  -0.1795   0.9006   0.5634
   0.000   0.7338   0.00568   0.00216  -0.1783   0.8924   1.0000
   0.250   0.7664   0.00565   0.00203  -0.1792   0.8825   1.0000
   0.500   0.7963   0.00568   0.00197  -0.1795   0.8710   1.0000
   0.750   0.8258   0.00572   0.00193  -0.1797   0.8592   1.0000
   1.000   0.8546   0.00579   0.00192  -0.1798   0.8466   1.0000
   1.250   0.8827   0.00588   0.00193  -0.1798   0.8337   1.0000
   1.500   0.9102   0.00598   0.00196  -0.1796   0.8207   1.0000
   1.750   0.9372   0.00609   0.00203  -0.1793   0.8073   1.0000
   2.000   0.9637   0.00621   0.00209  -0.1789   0.7930   1.0000
   2.250   0.9898   0.00634   0.00218  -0.1784   0.7780   1.0000
   2.500   1.0155   0.00649   0.00227  -0.1779   0.7624   1.0000
   2.750   1.0410   0.00664   0.00238  -0.1773   0.7461   1.0000
   3.000   1.0660   0.00682   0.00253  -0.1766   0.7278   1.0000
   3.250   1.0903   0.00700   0.00266  -0.1758   0.7061   1.0000
   3.500   1.1109   0.00733   0.00280  -0.1741   0.6618   1.0000
   3.750   1.1271   0.00790   0.00301  -0.1715   0.5849   1.0000
   4.000   1.1417   0.00875   0.00337  -0.1688   0.4902   1.0000
   4.250   1.1567   0.00977   0.00389  -0.1665   0.3912   1.0000
   4.500   1.1727   0.01085   0.00446  -0.1645   0.2923   1.0000
   4.750   1.1885   0.01205   0.00510  -0.1625   0.1909   1.0000
   5.000   1.2038   0.01336   0.00585  -0.1605   0.0945   1.0000
   5.250   1.2188   0.01478   0.00679  -0.1583   0.0194   1.0000
   5.500   1.2409   0.01537   0.00742  -0.1572   0.0140   1.0000
   5.750   1.2631   0.01593   0.00808  -0.1561   0.0127   1.0000
   6.000   1.2843   0.01661   0.00885  -0.1548   0.0116   1.0000
   6.250   1.3040   0.01747   0.00980  -0.1532   0.0108   1.0000
   6.500   1.3216   0.01855   0.01100  -0.1512   0.0103   1.0000
   6.750   1.3376   0.01979   0.01237  -0.1489   0.0101   1.0000
   7.000   1.3534   0.02107   0.01379  -0.1467   0.0101   1.0000
   7.250   1.3697   0.02240   0.01523  -0.1446   0.0101   1.0000
   7.500   1.3872   0.02374   0.01668  -0.1427   0.0103   1.0000
   7.750   1.4061   0.02522   0.01829  -0.1410   0.0107   1.0000
   8.000   1.4269   0.02720   0.02046  -0.1395   0.0110   1.0000
   8.250   1.4476   0.02917   0.02257  -0.1384   0.0108   1.0000
   8.500   1.4755   0.03388   0.02777  -0.1375   0.0125   1.0000
   9.750   1.4250   0.05396   0.05047  -0.1081   0.0192   1.0000
  10.000   1.4046   0.05726   0.05396  -0.1023   0.0189   1.0000
  10.250   1.3818   0.06117   0.05808  -0.0972   0.0186   1.0000
  10.500   1.3567   0.06569   0.06280  -0.0929   0.0184   1.0000
  10.750   1.3291   0.07084   0.06815  -0.0896   0.0183   1.0000
  11.000   1.2969   0.07709   0.07462  -0.0874   0.0190   1.0000
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