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AG12 (ag12-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: AG12 (ag12-il)
Reynolds number: 100,000
Max Cl/Cd: 47.12 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag12-il-100000.txt
Download as CSV file: xf-ag12-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG12                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5656   0.11248   0.10765   0.0179   1.0000   0.0736
  -8.500  -0.5699   0.10995   0.10520   0.0122   1.0000   0.0753
  -8.250  -0.5740   0.10717   0.10252   0.0054   1.0000   0.0758
  -8.000  -0.5521   0.10114   0.09644   0.0136   1.0000   0.0812
  -7.750  -0.5479   0.09774   0.09308   0.0116   1.0000   0.0852
  -7.500  -0.5500   0.09451   0.08995   0.0020   1.0000   0.0885
  -7.250  -0.5433   0.08919   0.08468  -0.0030   1.0000   0.0903
  -7.000  -0.5316   0.08584   0.08134   0.0025   1.0000   0.0939
  -6.750  -0.5205   0.08175   0.07726  -0.0025   1.0000   0.0995
  -6.500  -0.5073   0.07571   0.07122  -0.0136   1.0000   0.1042
  -6.250  -0.4951   0.07266   0.06819  -0.0107   1.0000   0.1086
  -6.000  -0.4756   0.06657   0.06200  -0.0217   1.0000   0.1179
  -5.750  -0.4611   0.06349   0.05896  -0.0199   1.0000   0.1232
  -5.500  -0.4315   0.05816   0.05316  -0.0337   1.0000   0.1434
  -5.250  -0.4187   0.05400   0.04924  -0.0301   1.0000   0.1460
  -5.000  -0.3965   0.05001   0.04514  -0.0329   1.0000   0.1595
  -4.750  -0.3406   0.03290   0.02622  -0.0452   1.0000   0.0602
  -4.500  -0.3123   0.02823   0.02097  -0.0460   1.0000   0.0550
  -4.250  -0.2838   0.02520   0.01739  -0.0463   1.0000   0.0563
  -4.000  -0.2551   0.02286   0.01447  -0.0461   1.0000   0.0601
  -3.750  -0.2260   0.02058   0.01169  -0.0455   1.0000   0.0619
  -3.500  -0.1988   0.01828   0.00931  -0.0451   1.0000   0.0675
  -3.250  -0.1718   0.01691   0.00791  -0.0446   1.0000   0.0841
  -3.000  -0.1448   0.01538   0.00646  -0.0441   1.0000   0.1135
  -2.750  -0.1182   0.01377   0.00532  -0.0439   1.0000   0.1969
  -2.500  -0.0929   0.01244   0.00478  -0.0437   1.0000   0.3567
  -2.250  -0.0691   0.01167   0.00453  -0.0427   1.0000   0.5005
  -2.000  -0.0471   0.01092   0.00430  -0.0409   1.0000   0.6413
  -1.750  -0.0251   0.00977   0.00393  -0.0378   1.0000   1.0000
  -1.500   0.0031   0.00978   0.00359  -0.0381   1.0000   1.0000
  -1.250   0.0301   0.00980   0.00333  -0.0381   1.0000   1.0000
  -1.000   0.0566   0.00983   0.00317  -0.0380   1.0000   1.0000
  -0.750   0.0828   0.00989   0.00307  -0.0378   1.0000   1.0000
  -0.500   0.1088   0.00996   0.00303  -0.0377   1.0000   1.0000
  -0.250   0.1345   0.01006   0.00304  -0.0376   1.0000   1.0000
   0.000   0.1600   0.01019   0.00308  -0.0375   1.0000   1.0000
   0.250   0.1852   0.01037   0.00321  -0.0375   1.0000   1.0000
   0.500   0.2102   0.01060   0.00341  -0.0376   1.0000   1.0000
   0.750   0.2346   0.01091   0.00371  -0.0379   1.0000   1.0000
   1.000   0.2812   0.01119   0.00402  -0.0426   0.9857   1.0000
   1.250   0.3427   0.01126   0.00414  -0.0493   0.9598   1.0000
   1.500   0.3942   0.01121   0.00417  -0.0534   0.9287   1.0000
   1.750   0.4326   0.01119   0.00419  -0.0545   0.8901   1.0000
   2.000   0.4620   0.01124   0.00420  -0.0534   0.8481   1.0000
   2.250   0.4867   0.01138   0.00425  -0.0515   0.8049   1.0000
   2.500   0.5101   0.01161   0.00434  -0.0495   0.7607   1.0000
   2.750   0.5335   0.01189   0.00454  -0.0477   0.7158   1.0000
   3.000   0.5572   0.01222   0.00471  -0.0461   0.6714   1.0000
   3.250   0.5813   0.01259   0.00492  -0.0446   0.6276   1.0000
   3.500   0.6056   0.01300   0.00517  -0.0434   0.5843   1.0000
   3.750   0.6301   0.01344   0.00546  -0.0423   0.5401   1.0000
   4.000   0.6548   0.01391   0.00585  -0.0413   0.4958   1.0000
   4.250   0.6794   0.01442   0.00623  -0.0404   0.4518   1.0000
   4.500   0.7040   0.01498   0.00664  -0.0396   0.4080   1.0000
   4.750   0.7284   0.01560   0.00710  -0.0388   0.3642   1.0000
   5.000   0.7529   0.01628   0.00768  -0.0382   0.3195   1.0000
   5.250   0.7771   0.01706   0.00839  -0.0375   0.2762   1.0000
   5.500   0.8011   0.01794   0.00914  -0.0369   0.2347   1.0000
   5.750   0.8249   0.01899   0.01005  -0.0362   0.1974   1.0000
   6.000   0.8485   0.02013   0.01109  -0.0356   0.1633   1.0000
   6.250   0.8718   0.02149   0.01233  -0.0349   0.1349   1.0000
   6.500   0.8953   0.02299   0.01387  -0.0341   0.1107   1.0000
   6.750   0.9182   0.02470   0.01561  -0.0334   0.0906   1.0000
   7.000   0.9417   0.02692   0.01804  -0.0324   0.0766   1.0000
   7.250   0.9640   0.02897   0.02027  -0.0316   0.0639   1.0000
   7.500   0.9858   0.03216   0.02351  -0.0308   0.0568   1.0000
   7.750   1.0063   0.03489   0.02693  -0.0295   0.0506   1.0000
   8.000   1.0252   0.03769   0.02986  -0.0289   0.0457   1.0000
   8.250   1.0373   0.04354   0.03621  -0.0279   0.0444   1.0000
   8.500   1.0471   0.04832   0.04166  -0.0268   0.0440   1.0000
   8.750   1.0520   0.05400   0.04787  -0.0260   0.0442   1.0000
   9.000   1.0529   0.06008   0.05435  -0.0256   0.0444   1.0000
   9.250   1.0509   0.06516   0.05987  -0.0253   0.0446   1.0000
   9.500   1.0444   0.06988   0.06495  -0.0255   0.0448   1.0000
   9.750   0.9785   0.08212   0.07788  -0.0350   0.0516   1.0000
  10.000   0.9583   0.09166   0.08744  -0.0434   0.0532   1.0000
  10.250   0.9481   0.09961   0.09537  -0.0481   0.0548   1.0000
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