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AG14 (ag14-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: AG14 (ag14-il)
Reynolds number: 1,000,000
Max Cl/Cd: 93.8 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag14-il-1000000.txt
Download as CSV file: xf-ag14-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG14                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5686   0.10140   0.09983   0.0202   1.0000   0.0059
  -8.250  -0.5644   0.09772   0.09617   0.0185   1.0000   0.0060
  -8.000  -0.5594   0.09428   0.09273   0.0168   1.0000   0.0061
  -7.750  -0.5541   0.09086   0.08933   0.0150   1.0000   0.0062
  -7.500  -0.5486   0.08733   0.08582   0.0127   1.0000   0.0063
  -7.250  -0.5395   0.08342   0.08192   0.0089   1.0000   0.0064
  -7.000  -0.5270   0.07913   0.07763   0.0041   1.0000   0.0065
  -6.750  -0.5119   0.07452   0.07302  -0.0016   1.0000   0.0067
  -6.500  -0.4940   0.06956   0.06805  -0.0079   1.0000   0.0069
  -6.250  -0.4730   0.06423   0.06269  -0.0149   1.0000   0.0071
  -6.000  -0.4489   0.05848   0.05690  -0.0221   1.0000   0.0074
  -5.750  -0.4219   0.05230   0.05064  -0.0293   1.0000   0.0078
  -5.500  -0.3925   0.04568   0.04389  -0.0358   1.0000   0.0083
  -5.250  -0.3608   0.04068   0.03875  -0.0394   1.0000   0.0094
  -4.500  -0.2734   0.01669   0.01308  -0.0487   1.0000   0.0077
  -4.250  -0.2459   0.01451   0.01058  -0.0486   1.0000   0.0083
  -4.000  -0.2186   0.01274   0.00851  -0.0484   1.0000   0.0085
  -3.750  -0.1916   0.01143   0.00699  -0.0480   1.0000   0.0088
  -3.500  -0.1652   0.01098   0.00647  -0.0476   1.0000   0.0094
  -3.250  -0.1383   0.00966   0.00497  -0.0472   1.0000   0.0094
  -3.000  -0.1113   0.00855   0.00373  -0.0468   1.0000   0.0093
  -2.750  -0.0844   0.00788   0.00296  -0.0464   1.0000   0.0096
  -2.500  -0.0566   0.00699   0.00190  -0.0462   1.0000   0.0120
  -2.250  -0.0198   0.00670   0.00159  -0.0480   0.9961   0.0157
  -2.000   0.0168   0.00628   0.00122  -0.0498   0.9903   0.0377
  -1.750   0.0529   0.00595   0.00106  -0.0516   0.9808   0.0831
  -1.500   0.0868   0.00563   0.00092  -0.0529   0.9639   0.1451
  -1.250   0.1166   0.00534   0.00083  -0.0531   0.9348   0.2214
  -1.000   0.1417   0.00517   0.00076  -0.0523   0.8955   0.2944
  -0.750   0.1668   0.00504   0.00071  -0.0515   0.8538   0.3798
  -0.500   0.1926   0.00491   0.00069  -0.0510   0.8115   0.4779
  -0.250   0.2185   0.00465   0.00069  -0.0506   0.7688   0.6248
   0.000   0.2382   0.00398   0.00071  -0.0485   0.7278   0.9017
   0.250   0.2679   0.00404   0.00067  -0.0487   0.6836   1.0000
   0.500   0.2951   0.00424   0.00067  -0.0484   0.6401   1.0000
   0.750   0.3225   0.00444   0.00068  -0.0482   0.5987   1.0000
   1.000   0.3500   0.00463   0.00070  -0.0481   0.5590   1.0000
   1.250   0.3775   0.00483   0.00074  -0.0479   0.5200   1.0000
   1.500   0.4051   0.00503   0.00078  -0.0478   0.4834   1.0000
   1.750   0.4326   0.00524   0.00085  -0.0477   0.4480   1.0000
   2.000   0.4602   0.00543   0.00091  -0.0476   0.4155   1.0000
   2.250   0.4876   0.00564   0.00098  -0.0475   0.3823   1.0000
   2.500   0.5151   0.00585   0.00107  -0.0474   0.3529   1.0000
   2.750   0.5426   0.00607   0.00117  -0.0473   0.3218   1.0000
   3.000   0.5700   0.00628   0.00129  -0.0472   0.2953   1.0000
   3.250   0.5974   0.00652   0.00142  -0.0471   0.2657   1.0000
   3.500   0.6247   0.00675   0.00155  -0.0470   0.2403   1.0000
   3.750   0.6520   0.00700   0.00171  -0.0469   0.2156   1.0000
   4.000   0.6791   0.00728   0.00187  -0.0467   0.1886   1.0000
   4.250   0.7063   0.00753   0.00207  -0.0466   0.1667   1.0000
   4.500   0.7333   0.00782   0.00228  -0.0465   0.1432   1.0000
   4.750   0.7602   0.00814   0.00250  -0.0463   0.1221   1.0000
   5.000   0.7871   0.00843   0.00274  -0.0462   0.1038   1.0000
   5.250   0.8139   0.00876   0.00302  -0.0460   0.0869   1.0000
   5.500   0.8406   0.00911   0.00330  -0.0459   0.0717   1.0000
   5.750   0.8672   0.00946   0.00360  -0.0457   0.0579   1.0000
   6.000   0.8937   0.00982   0.00393  -0.0455   0.0458   1.0000
   6.250   0.9200   0.01022   0.00428  -0.0453   0.0347   1.0000
   6.500   0.9463   0.01062   0.00465  -0.0451   0.0267   1.0000
   6.750   0.9725   0.01104   0.00509  -0.0448   0.0198   1.0000
   7.000   0.9983   0.01157   0.00562  -0.0445   0.0132   1.0000
   7.250   1.0234   0.01234   0.00645  -0.0441   0.0085   1.0000
   7.500   1.0489   0.01290   0.00705  -0.0437   0.0067   1.0000
   7.750   1.0721   0.01419   0.00854  -0.0430   0.0051   1.0000
   8.000   1.0971   0.01484   0.00929  -0.0426   0.0049   1.0000
   8.250   1.1213   0.01565   0.01022  -0.0421   0.0046   1.0000
   8.500   1.1450   0.01655   0.01125  -0.0415   0.0043   1.0000
   8.750   1.1682   0.01751   0.01237  -0.0409   0.0041   1.0000
   9.000   1.1912   0.01844   0.01341  -0.0404   0.0037   1.0000
   9.250   1.2138   0.01939   0.01447  -0.0398   0.0034   1.0000
   9.500   1.2334   0.02099   0.01625  -0.0390   0.0031   1.0000
   9.750   1.2345   0.02672   0.02264  -0.0364   0.0027   1.0000
  10.000   1.2519   0.02841   0.02454  -0.0356   0.0027   1.0000
  10.250   1.2641   0.03096   0.02738  -0.0344   0.0026   1.0000
  10.500   1.2704   0.03425   0.03099  -0.0330   0.0026   1.0000
  10.750   1.2699   0.03813   0.03522  -0.0316   0.0026   1.0000
  11.000   1.2613   0.04250   0.03990  -0.0303   0.0026   1.0000
  11.250   1.2438   0.04659   0.04423  -0.0291   0.0026   1.0000
  11.500   1.2239   0.05261   0.05046  -0.0326   0.0026   1.0000
  11.750   1.2060   0.06172   0.05977  -0.0415   0.0026   1.0000
  12.000   1.1856   0.07307   0.07129  -0.0512   0.0027   1.0000
  12.250   1.1555   0.08575   0.08408  -0.0595   0.0027   1.0000
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