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E61 (5.64%) (e61-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: E61 (5.64%) (e61-il)
Reynolds number: 100,000
Max Cl/Cd: 75.58 at α=6.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e61-il-100000.txt
Download as CSV file: xf-e61-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: E61  (5.64%)                                    
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -5.250  -0.3700   0.10529   0.10141  -0.0167   1.0000   0.0431
  -5.000  -0.3720   0.10100   0.09714  -0.0111   1.0000   0.0446
  -4.750  -0.3693   0.09835   0.09452  -0.0104   1.0000   0.0458
  -4.500  -0.3646   0.09581   0.09201  -0.0110   1.0000   0.0472
  -4.250  -0.3572   0.09322   0.08944  -0.0125   1.0000   0.0489
  -4.000  -0.3463   0.09050   0.08673  -0.0152   1.0000   0.0510
  -3.750  -0.3253   0.08765   0.08383  -0.0214   1.0000   0.0539
  -3.500  -0.2576   0.08376   0.07981  -0.0441   1.0000   0.0563
  -3.250  -0.2657   0.07991   0.07603  -0.0380   1.0000   0.0573
  -3.000  -0.2474   0.07634   0.07246  -0.0384   0.9958   0.0606
  -2.750  -0.1559   0.07022   0.06613  -0.0612   0.9858   0.0717
  -2.250  -0.0187   0.06150   0.05701  -0.0897   0.9737   0.0970
  -2.000   0.0091   0.05743   0.05294  -0.0921   0.9684   0.1020
  -1.750   0.0678   0.05367   0.04902  -0.1024   0.9637   0.1167
  -1.500   0.1414   0.04993   0.04500  -0.1159   0.9607   0.1420
  -1.250   0.1836   0.04757   0.04257  -0.1210   0.9553   0.1712
  -0.500   0.4312   0.03505   0.02760  -0.1571   0.9507   0.0785
  -0.250   0.4860   0.03374   0.02547  -0.1616   0.9482   0.0699
   0.000   0.5236   0.03259   0.02420  -0.1642   0.9425   0.0746
   0.250   0.5653   0.03197   0.02339  -0.1668   0.9378   0.0802
   0.500   0.6118   0.03133   0.02270  -0.1704   0.9345   0.0874
   0.750   0.6456   0.03106   0.02243  -0.1718   0.9268   0.1102
   1.000   0.6858   0.02915   0.02227  -0.1740   0.9236   1.0000
   1.250   0.7149   0.02954   0.02240  -0.1747   0.9146   1.0000
   1.500   0.7581   0.02972   0.02235  -0.1776   0.9096   1.0000
   1.750   0.7849   0.03008   0.02259  -0.1778   0.8997   1.0000
   2.000   0.8186   0.03032   0.02274  -0.1791   0.8917   1.0000
   2.250   0.8572   0.03036   0.02272  -0.1811   0.8846   1.0000
   2.500   0.8865   0.03060   0.02292  -0.1815   0.8748   1.0000
   2.750   0.9328   0.03027   0.02257  -0.1845   0.8697   1.0000
   3.000   0.9609   0.03042   0.02273  -0.1846   0.8587   1.0000
   3.250   0.9933   0.03039   0.02279  -0.1852   0.8489   1.0000
   3.500   1.0412   0.02963   0.02210  -0.1881   0.8433   1.0000
   3.750   1.0725   0.02944   0.02199  -0.1883   0.8321   1.0000
   4.250   1.1613   0.02745   0.02033  -0.1920   0.8168   1.0000
   4.500   1.1988   0.02659   0.01963  -0.1926   0.8059   1.0000
   4.750   1.2646   0.02425   0.01754  -0.1973   0.8011   1.0000
   5.000   1.3116   0.02289   0.01642  -0.1993   0.7882   1.0000
   5.250   1.3487   0.02195   0.01576  -0.1995   0.7702   1.0000
   5.500   1.3928   0.02070   0.01474  -0.2006   0.7487   1.0000
   5.750   1.4279   0.01994   0.01416  -0.2003   0.7202   1.0000
   6.000   1.4524   0.01967   0.01403  -0.1983   0.6845   1.0000
   6.250   1.4760   0.01953   0.01401  -0.1961   0.6387   1.0000
   6.500   1.4902   0.01979   0.01406  -0.1921   0.5673   1.0000
   6.750   1.4950   0.02083   0.01451  -0.1867   0.4683   1.0000
   7.000   1.4915   0.02255   0.01552  -0.1807   0.3680   1.0000
   7.250   1.4846   0.02459   0.01686  -0.1746   0.2738   1.0000
   7.500   1.4786   0.02703   0.01856  -0.1692   0.1743   1.0000
   7.750   1.4635   0.03113   0.02157  -0.1626   0.0673   1.0000
   8.000   1.4594   0.03421   0.02453  -0.1573   0.0472   1.0000
   8.250   1.4611   0.03671   0.02710  -0.1532   0.0405   1.0000
   8.500   1.4656   0.03945   0.02996  -0.1496   0.0371   1.0000
   8.750   1.4811   0.04201   0.03267  -0.1473   0.0344   1.0000
   9.000   1.5024   0.04465   0.03546  -0.1461   0.0311   1.0000
   9.250   1.5427   0.04921   0.04019  -0.1479   0.0275   1.0000
   9.500   1.5799   0.05392   0.04538  -0.1488   0.0270   1.0000
   9.750   1.6001   0.05886   0.05084  -0.1477   0.0273   1.0000
  10.000   1.6084   0.06406   0.05652  -0.1452   0.0277   1.0000
  10.250   1.6162   0.07113   0.06404  -0.1434   0.0284   1.0000
  10.500   1.6110   0.07308   0.06632  -0.1386   0.0288   1.0000
  10.750   1.6000   0.07538   0.06903  -0.1334   0.0295   1.0000
  11.000   1.5781   0.07924   0.07343  -0.1278   0.0306   1.0000
  11.250   1.5507   0.08422   0.07892  -0.1233   0.0317   1.0000
  11.500   1.5255   0.08970   0.08480  -0.1203   0.0325   1.0000
  11.750   1.5014   0.09540   0.09083  -0.1187   0.0331   1.0000
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