XFOIL         Version 6.94
  
 Calculated polar for: AG19                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.2741   0.00478   0.00122  -0.0572  0.8931  1.0000
   0.500   0.3254   0.00500   0.00114  -0.0553  0.8323  1.0000
   1.000   0.3777   0.00530   0.00112  -0.0539  0.7666  1.0000
   1.500   0.4308   0.00567   0.00116  -0.0528  0.6948  1.0000
   2.000   0.4840   0.00613   0.00127  -0.0519  0.6092  1.0000
   2.500   0.5372   0.00673   0.00144  -0.0512  0.5083  1.0000
   3.000   0.5905   0.00741   0.00172  -0.0507  0.4101  1.0000
   3.500   0.6441   0.00809   0.00211  -0.0503  0.3298  1.0000
   4.000   0.6978   0.00878   0.00253  -0.0500  0.2624  1.0000
   4.500   0.7513   0.00948   0.00306  -0.0496  0.2025  1.0000
   5.000   0.8044   0.01027   0.00367  -0.0492  0.1497  1.0000
   5.500   0.8571   0.01113   0.00444  -0.0488  0.1064  1.0000
   6.000   0.9092   0.01215   0.00536  -0.0482  0.0734  1.0000
   6.500   0.9608   0.01322   0.00644  -0.0476  0.0478  1.0000
   7.000   1.0105   0.01478   0.00805  -0.0467  0.0268  1.0000
   8.000   1.1021   0.01976   0.01356  -0.0436  0.0107  1.0000
   8.500   1.1450   0.02273   0.01695  -0.0417  0.0088  1.0000
   9.000   1.1775   0.02762   0.02233  -0.0395  0.0070  1.0000
   9.500   1.2058   0.03297   0.02840  -0.0369  0.0065  1.0000
  10.000   1.2180   0.03996   0.03619  -0.0340  0.0061  1.0000
  10.500   1.2013   0.04858   0.04552  -0.0311  0.0060  1.0000
  11.000   1.1655   0.05897   0.05636  -0.0346  0.0060  1.0000
  11.500   1.1313   0.07698   0.07473  -0.0506  0.0062  1.0000
  12.000   1.0874   0.09895   0.09687  -0.0651  0.0066  1.0000