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AG11 (ag11-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: AG11 (ag11-il)
Reynolds number: 50,000
Max Cl/Cd: 34.61 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ag11-il-50000-n5.txt
Download as CSV file: xf-ag11-il-50000-n5.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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5319   0.11433   0.10750   0.0119   1.0000   0.0807
  -8.250  -0.5349   0.11233   0.10560   0.0071   1.0000   0.0817
  -8.000  -0.5346   0.10984   0.10318   0.0006   1.0000   0.0822
  -7.750  -0.5173   0.10324   0.09661   0.0072   1.0000   0.0850
  -7.500  -0.5094   0.09965   0.09305   0.0068   1.0000   0.0887
  -7.250  -0.5042   0.09641   0.08986   0.0030   1.0000   0.0932
  -7.000  -0.4980   0.09384   0.08728  -0.0091   1.0000   0.0963
  -6.750  -0.4876   0.08865   0.08217  -0.0030   1.0000   0.1002
  -6.500  -0.4767   0.08514   0.07867  -0.0058   1.0000   0.1071
  -6.000  -0.4528   0.07750   0.07107  -0.0110   1.0000   0.1201
  -5.750  -0.4392   0.07357   0.06712  -0.0145   1.0000   0.1286
  -5.500  -0.4233   0.06996   0.06345  -0.0186   1.0000   0.1410
  -5.250  -0.4069   0.06648   0.05991  -0.0212   1.0000   0.1547
  -5.000  -0.3916   0.06272   0.05621  -0.0202   1.0000   0.1616
  -4.750  -0.3738   0.05926   0.05271  -0.0220   1.0000   0.1773
  -4.500  -0.3190   0.05116   0.04378  -0.0309   1.0000   0.0625
  -4.250  -0.2907   0.04694   0.03918  -0.0327   1.0000   0.0508
  -4.000  -0.2656   0.04326   0.03521  -0.0338   1.0000   0.0469
  -3.750  -0.2329   0.03979   0.03097  -0.0348   1.0000   0.0418
  -3.500  -0.2075   0.03696   0.02783  -0.0349   1.0000   0.0405
  -3.250  -0.1820   0.03411   0.02470  -0.0350   1.0000   0.0399
  -3.000  -0.1556   0.03158   0.02182  -0.0348   1.0000   0.0394
  -2.750  -0.1286   0.02941   0.01924  -0.0344   1.0000   0.0400
  -2.500  -0.1016   0.02751   0.01694  -0.0339   1.0000   0.0417
  -2.250  -0.0766   0.02555   0.01483  -0.0335   1.0000   0.0445
  -2.000  -0.0506   0.02398   0.01304  -0.0327   1.0000   0.0460
  -1.750  -0.0247   0.02257   0.01143  -0.0317   1.0000   0.0477
  -1.500   0.0007   0.02143   0.01008  -0.0306   1.0000   0.0514
  -1.250   0.0264   0.02038   0.00891  -0.0298   1.0000   0.0587
  -1.000   0.0519   0.01951   0.00785  -0.0291   1.0000   0.0666
  -0.750   0.0775   0.01864   0.00701  -0.0287   1.0000   0.0863
  -0.500   0.1031   0.01589   0.00623  -0.0292   1.0000   0.4762
  -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   0.2018   0.01541   0.00558  -0.0268   0.9928   1.0000
   0.750   0.2531   0.01563   0.00560  -0.0318   0.9699   1.0000
   1.000   0.3020   0.01579   0.00561  -0.0360   0.9445   1.0000
   1.250   0.3463   0.01590   0.00564  -0.0390   0.9159   1.0000
   1.500   0.3861   0.01599   0.00567  -0.0408   0.8840   1.0000
   1.750   0.4204   0.01609   0.00571  -0.0413   0.8483   1.0000
   2.000   0.4513   0.01621   0.00575  -0.0410   0.8112   1.0000
   2.250   0.4788   0.01636   0.00581  -0.0399   0.7724   1.0000
   2.500   0.5041   0.01657   0.00593  -0.0384   0.7320   1.0000
   2.750   0.5284   0.01683   0.00605  -0.0368   0.6908   1.0000
   3.000   0.5523   0.01714   0.00621  -0.0352   0.6487   1.0000
   3.250   0.5760   0.01750   0.00642  -0.0337   0.6059   1.0000
   3.500   0.5996   0.01791   0.00667  -0.0324   0.5628   1.0000
   3.750   0.6232   0.01837   0.00703  -0.0311   0.5198   1.0000
   4.000   0.6468   0.01888   0.00738  -0.0300   0.4774   1.0000
   4.250   0.6704   0.01945   0.00780  -0.0290   0.4362   1.0000
   4.500   0.6940   0.02006   0.00828  -0.0281   0.3968   1.0000
   4.750   0.7181   0.02075   0.00882  -0.0275   0.3594   1.0000
   5.000   0.7423   0.02147   0.00955  -0.0269   0.3228   1.0000
   5.250   0.7661   0.02227   0.01024  -0.0263   0.2898   1.0000
   5.500   0.7899   0.02312   0.01103  -0.0258   0.2597   1.0000
   5.750   0.8134   0.02403   0.01189  -0.0253   0.2323   1.0000
   6.000   0.8369   0.02500   0.01293  -0.0248   0.2074   1.0000
   6.250   0.8604   0.02607   0.01403  -0.0242   0.1870   1.0000
   6.500   0.8836   0.02721   0.01523  -0.0237   0.1691   1.0000
   6.750   0.9063   0.02837   0.01644  -0.0231   0.1533   1.0000
   7.000   0.9296   0.02974   0.01800  -0.0225   0.1416   1.0000
   7.250   0.9524   0.03121   0.01962  -0.0218   0.1315   1.0000
   7.500   0.9743   0.03260   0.02108  -0.0213   0.1216   1.0000
   7.750   0.9963   0.03422   0.02297  -0.0206   0.1130   1.0000
   8.000   1.0182   0.03622   0.02519  -0.0200   0.1075   1.0000
   8.250   1.0387   0.03830   0.02769  -0.0193   0.1011   1.0000
   8.500   1.0579   0.04021   0.02978  -0.0187   0.0950   1.0000
   8.750   1.0748   0.04285   0.03294  -0.0180   0.0896   1.0000
   9.000   1.0920   0.04537   0.03573  -0.0173   0.0861   1.0000
   9.250   1.1062   0.04849   0.03916  -0.0166   0.0832   1.0000
   9.500   1.1107   0.05272   0.04411  -0.0158   0.0797   1.0000
   9.750   1.1166   0.05601   0.04779  -0.0152   0.0758   1.0000
  10.000   1.1263   0.05860   0.05059  -0.0145   0.0728   1.0000
  10.250   1.1276   0.06267   0.05491  -0.0140   0.0714   1.0000
  10.500   1.1117   0.06826   0.06095  -0.0140   0.0709   1.0000
  10.750   1.0898   0.07385   0.06682  -0.0146   0.0708   1.0000
  11.000   1.0642   0.08003   0.07316  -0.0169   0.0709   1.0000
  11.250   1.0376   0.08786   0.08102  -0.0222   0.0713   1.0000
  11.500   1.0116   0.09740   0.09058  -0.0292   0.0716   1.0000
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