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MA409 (smoothed) (ma409sm-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: MA409 (smoothed) (ma409sm-il)
Reynolds number: 500,000
Max Cl/Cd: 121.64 at α=2.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ma409sm-il-500000.txt
Download as CSV file: xf-ma409sm-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: MA409 (smoothed)                                
  
 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
  ------ -------- --------- --------- -------- -------- --------
   0.250   0.4700   0.00528   0.00176  -0.1085   0.9080   1.0000
   0.500   0.4983   0.00526   0.00169  -0.1083   0.8969   1.0000
   0.750   0.5263   0.00524   0.00163  -0.1081   0.8843   1.0000
   1.000   0.5539   0.00525   0.00159  -0.1078   0.8699   1.0000
   1.250   0.5813   0.00528   0.00156  -0.1074   0.8535   1.0000
   1.500   0.6084   0.00534   0.00155  -0.1070   0.8354   1.0000
   1.750   0.6351   0.00543   0.00159  -0.1065   0.8144   1.0000
   2.000   0.6614   0.00554   0.00162  -0.1059   0.7914   1.0000
   2.500   0.7128   0.00586   0.00175  -0.1046   0.7358   1.0000
   2.750   0.7380   0.00607   0.00185  -0.1039   0.7033   1.0000
   3.000   0.7623   0.00634   0.00199  -0.1030   0.6640   1.0000
   3.250   0.7839   0.00680   0.00214  -0.1015   0.5958   1.0000
   3.500   0.8054   0.00735   0.00235  -0.1002   0.5230   1.0000
   3.750   0.8277   0.00788   0.00260  -0.0992   0.4598   1.0000
   4.000   0.8503   0.00844   0.00287  -0.0982   0.3937   1.0000
   4.250   0.8727   0.00908   0.00320  -0.0973   0.3212   1.0000
   4.500   0.8937   0.00992   0.00361  -0.0963   0.2307   1.0000
   4.750   0.9131   0.01105   0.00418  -0.0951   0.1217   1.0000
   5.000   0.9315   0.01242   0.00503  -0.0937   0.0347   1.0000
   5.250   0.9551   0.01308   0.00573  -0.0927   0.0255   1.0000
   5.500   0.9754   0.01427   0.00704  -0.0912   0.0207   1.0000
   5.750   0.9992   0.01483   0.00768  -0.0903   0.0197   1.0000
   6.000   1.0225   0.01545   0.00837  -0.0895   0.0182   1.0000
   6.250   1.0443   0.01628   0.00929  -0.0883   0.0171   1.0000
   6.500   1.0658   0.01719   0.01027  -0.0871   0.0162   1.0000
   6.750   1.0867   0.01820   0.01135  -0.0858   0.0153   1.0000
   7.000   1.1057   0.01966   0.01289  -0.0843   0.0142   1.0000
   7.250   1.1227   0.02237   0.01578  -0.0824   0.0129   1.0000
   7.500   1.1453   0.02330   0.01688  -0.0814   0.0123   1.0000
   7.750   1.1666   0.02489   0.01864  -0.0801   0.0117   1.0000
   8.000   1.1869   0.02684   0.02081  -0.0787   0.0112   1.0000
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