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AG14 (ag14-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: AG14 (ag14-il)
Reynolds number: 200,000
Max Cl/Cd: 62.17 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag14-il-200000.txt
Download as CSV file: xf-ag14-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
   0.250   0.2269   0.00734   0.00227  -0.0434   0.9971   1.0000
   0.500   0.2772   0.00737   0.00226  -0.0481   0.9805   1.0000
   0.750   0.3218   0.00735   0.00222  -0.0513   0.9574   1.0000
   1.250   0.3940   0.00733   0.00210  -0.0535   0.8880   1.0000
   1.500   0.4202   0.00742   0.00207  -0.0523   0.8414   1.0000
   1.750   0.4443   0.00761   0.00209  -0.0507   0.7918   1.0000
   2.000   0.4684   0.00786   0.00213  -0.0493   0.7402   1.0000
   2.250   0.4929   0.00816   0.00220  -0.0481   0.6883   1.0000
   2.500   0.5178   0.00849   0.00230  -0.0472   0.6371   1.0000
   2.750   0.5431   0.00885   0.00243  -0.0464   0.5871   1.0000
   3.000   0.5686   0.00921   0.00264  -0.0457   0.5384   1.0000
   3.250   0.5943   0.00959   0.00283  -0.0452   0.4923   1.0000
   3.500   0.6200   0.00999   0.00305  -0.0447   0.4483   1.0000
   3.750   0.6459   0.01039   0.00329  -0.0443   0.4057   1.0000
   4.000   0.6718   0.01081   0.00357  -0.0439   0.3644   1.0000
   4.250   0.6977   0.01126   0.00393  -0.0436   0.3246   1.0000
   4.500   0.7236   0.01173   0.00428  -0.0433   0.2856   1.0000
   4.750   0.7494   0.01223   0.00467  -0.0430   0.2474   1.0000
   5.000   0.7750   0.01280   0.00513  -0.0427   0.2125   1.0000
   5.250   0.8005   0.01342   0.00564  -0.0424   0.1789   1.0000
   5.500   0.8257   0.01412   0.00629  -0.0420   0.1507   1.0000
   5.750   0.8508   0.01484   0.00695  -0.0417   0.1255   1.0000
   6.000   0.8757   0.01567   0.00776  -0.0412   0.1047   1.0000
   6.250   0.9004   0.01649   0.00857  -0.0408   0.0861   1.0000
   6.500   0.9239   0.01768   0.00972  -0.0402   0.0700   1.0000
   6.750   0.9483   0.01863   0.01079  -0.0397   0.0552   1.0000
   7.000   0.9713   0.02003   0.01232  -0.0389   0.0435   1.0000
   7.250   0.9929   0.02183   0.01417  -0.0381   0.0341   1.0000
   7.500   1.0164   0.02336   0.01595  -0.0371   0.0288   1.0000
   7.750   1.0353   0.02614   0.01888  -0.0362   0.0240   1.0000
   8.000   1.0573   0.02816   0.02128  -0.0351   0.0219   1.0000
   8.250   1.0763   0.03117   0.02473  -0.0339   0.0205   1.0000
   8.500   1.0921   0.03485   0.02891  -0.0326   0.0198   1.0000
   8.750   1.1026   0.03949   0.03414  -0.0312   0.0197   1.0000
   9.000   1.1054   0.04522   0.04051  -0.0299   0.0201   1.0000
   9.250   1.1005   0.05151   0.04737  -0.0291   0.0207   1.0000
   9.500   1.0891   0.05768   0.05396  -0.0289   0.0212   1.0000
   9.750   1.0723   0.06331   0.05987  -0.0295   0.0216   1.0000
  10.000   1.0519   0.06881   0.06557  -0.0320   0.0218   1.0000
  10.250   1.0319   0.07667   0.07359  -0.0397   0.0220   1.0000
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