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AG11 (ag11-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: AG11 (ag11-il)
Reynolds number: 50,000
Max Cl/Cd: 33.02 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag11-il-50000.txt
Download as CSV file: xf-ag11-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG11                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.5272   0.10598   0.09946   0.0114   1.0000   0.1518
  -7.500  -0.5205   0.10187   0.09538   0.0108   1.0000   0.1570
  -7.250  -0.5173   0.09911   0.09268   0.0079   1.0000   0.1662
  -7.000  -0.5070   0.09484   0.08846   0.0086   1.0000   0.1748
  -6.750  -0.5046   0.09186   0.08555   0.0026   1.0000   0.1836
  -6.500  -0.4980   0.08924   0.08293  -0.0030   1.0000   0.1965
  -6.250  -0.4875   0.08504   0.07881  -0.0023   1.0000   0.2112
  -6.000  -0.4759   0.08080   0.07464   0.0005   1.0000   0.2276
  -5.750  -0.4647   0.07728   0.07117   0.0017   1.0000   0.2475
  -5.500  -0.4556   0.07388   0.06782   0.0006   1.0000   0.2712
  -5.250  -0.4458   0.07060   0.06460   0.0015   1.0000   0.2997
  -5.000  -0.4357   0.06744   0.06153   0.0046   1.0000   0.3333
  -4.750  -0.4277   0.06449   0.05864   0.0085   1.0000   0.3767
  -4.500  -0.1443   0.04397   0.03753   0.0139   1.0000   1.0000
  -4.250  -0.1308   0.04149   0.03510   0.0123   1.0000   1.0000
  -4.000  -0.1316   0.04007   0.03377   0.0141   1.0000   0.9938
  -3.750  -0.1820   0.04177   0.03567   0.0268   1.0000   0.9484
  -3.500  -0.2291   0.04274   0.03686   0.0361   1.0000   0.8909
  -3.250  -0.2743   0.04306   0.03741   0.0424   1.0000   0.8304
  -3.000  -0.1864   0.03539   0.02718  -0.0354   1.0000   0.1769
  -2.750  -0.1477   0.03249   0.02342  -0.0363   1.0000   0.1373
  -2.500  -0.1160   0.03028   0.02054  -0.0360   1.0000   0.1245
  -2.250  -0.0890   0.02779   0.01789  -0.0355   1.0000   0.1209
  -2.000  -0.0604   0.02577   0.01550  -0.0348   1.0000   0.1153
  -1.750  -0.0317   0.02426   0.01352  -0.0337   1.0000   0.1116
  -1.500  -0.0051   0.02257   0.01172  -0.0328   1.0000   0.1137
  -1.250   0.0208   0.02134   0.01034  -0.0316   1.0000   0.1237
  -1.000   0.0483   0.01999   0.00896  -0.0306   1.0000   0.1357
  -0.750   0.0747   0.01876   0.00779  -0.0298   1.0000   0.1675
  -0.500   0.1011   0.01463   0.00627  -0.0277   1.0000   1.0000
  -0.250   0.1251   0.01477   0.00573  -0.0267   1.0000   1.0000
   0.000   0.1480   0.01494   0.00557  -0.0260   1.0000   1.0000
   0.250   0.1705   0.01516   0.00553  -0.0255   1.0000   1.0000
   0.500   0.1928   0.01543   0.00563  -0.0251   1.0000   1.0000
   0.750   0.2147   0.01578   0.00584  -0.0249   1.0000   1.0000
   1.000   0.2360   0.01620   0.00616  -0.0248   1.0000   1.0000
   1.250   0.2566   0.01673   0.00661  -0.0249   1.0000   1.0000
   1.500   0.2764   0.01739   0.00722  -0.0252   1.0000   1.0000
   1.750   0.2952   0.01820   0.00802  -0.0259   1.0000   1.0000
   2.000   0.3511   0.01913   0.00898  -0.0335   0.9798   1.0000
   2.250   0.4393   0.01955   0.00955  -0.0454   0.9357   1.0000
   2.500   0.5130   0.01943   0.00965  -0.0527   0.8904   1.0000
   2.750   0.5590   0.01924   0.00955  -0.0535   0.8441   1.0000
   3.000   0.5875   0.01922   0.00952  -0.0510   0.7967   1.0000
   3.250   0.6099   0.01934   0.00957  -0.0474   0.7485   1.0000
   3.500   0.6308   0.01955   0.00964  -0.0437   0.7009   1.0000
   3.750   0.6514   0.01995   0.00995  -0.0406   0.6501   1.0000
   4.000   0.6729   0.02042   0.01020  -0.0377   0.6012   1.0000
   4.250   0.6947   0.02104   0.01065  -0.0353   0.5504   1.0000
   4.500   0.7170   0.02175   0.01111  -0.0332   0.5028   1.0000
   4.750   0.7395   0.02259   0.01174  -0.0314   0.4560   1.0000
   5.000   0.7623   0.02356   0.01262  -0.0298   0.4113   1.0000
   5.250   0.7853   0.02465   0.01352  -0.0285   0.3709   1.0000
   5.500   0.8087   0.02588   0.01464  -0.0273   0.3340   1.0000
   5.750   0.8320   0.02721   0.01584  -0.0263   0.3013   1.0000
   6.000   0.8562   0.02879   0.01740  -0.0254   0.2750   1.0000
   6.250   0.8794   0.03044   0.01923  -0.0246   0.2503   1.0000
   6.500   0.9031   0.03229   0.02113  -0.0239   0.2315   1.0000
   6.750   0.9266   0.03462   0.02370  -0.0233   0.2175   1.0000
   7.000   0.9487   0.03685   0.02622  -0.0226   0.2034   1.0000
   7.250   0.9709   0.03908   0.02856  -0.0220   0.1912   1.0000
   7.500   0.9874   0.04302   0.03334  -0.0216   0.1846   1.0000
   7.750   1.0091   0.04587   0.03625  -0.0209   0.1780   1.0000
   8.000   1.0166   0.05084   0.04207  -0.0208   0.1727   1.0000
   8.250   1.0388   0.05303   0.04413  -0.0199   0.1638   1.0000
   8.500   1.0415   0.05897   0.05075  -0.0202   0.1632   1.0000
   8.750   1.0425   0.06516   0.05739  -0.0207   0.1635   1.0000
   9.000   0.9952   0.07776   0.07062  -0.0273   0.1732   1.0000
   9.250   0.9778   0.08634   0.07932  -0.0318   0.1795   1.0000
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