XFOIL         Version 6.94
  
 Calculated polar for: AG16                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.2463   0.00601   0.00182  -0.0551  0.8997  1.0000
   0.500   0.2980   0.00615   0.00171  -0.0532  0.8417  1.0000
   1.000   0.3499   0.00642   0.00167  -0.0516  0.7791  1.0000
   1.500   0.4027   0.00676   0.00173  -0.0503  0.7158  1.0000
   2.000   0.4561   0.00716   0.00184  -0.0493  0.6512  1.0000
   2.500   0.5099   0.00760   0.00202  -0.0486  0.5863  1.0000
   3.000   0.5636   0.00810   0.00229  -0.0479  0.5197  1.0000
   3.500   0.6173   0.00866   0.00262  -0.0473  0.4513  1.0000
   4.000   0.6707   0.00931   0.00306  -0.0468  0.3795  1.0000
   4.500   0.7238   0.01006   0.00359  -0.0463  0.3079  1.0000
   5.000   0.7764   0.01091   0.00426  -0.0459  0.2387  1.0000
   5.500   0.8286   0.01190   0.00505  -0.0454  0.1737  1.0000
   6.000   0.8796   0.01312   0.00608  -0.0449  0.1086  1.0000
   6.500   0.9288   0.01481   0.00757  -0.0440  0.0567  1.0000
   7.000   0.9763   0.01687   0.00963  -0.0428  0.0295  1.0000
   7.500   1.0180   0.02016   0.01313  -0.0409  0.0178  1.0000
   8.000   1.0620   0.02300   0.01634  -0.0389  0.0143  1.0000
   8.500   1.1010   0.02697   0.02069  -0.0367  0.0126  1.0000
   9.000   1.1316   0.03307   0.02744  -0.0339  0.0121  1.0000
   9.500   1.1503   0.04026   0.03547  -0.0308  0.0123  1.0000
  10.000   1.1461   0.04876   0.04468  -0.0278  0.0119  1.0000
  10.500   1.1172   0.05706   0.05351  -0.0254  0.0117  1.0000
  11.000   1.0864   0.06712   0.06398  -0.0305  0.0118  1.0000