XFOIL         Version 6.94
  
 Calculated polar for: AG17                                            
  
 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   0.2587   0.00595   0.00177  -0.0569  0.9222  1.0000
   0.500   0.3119   0.00605   0.00165  -0.0553  0.8669  1.0000
   1.000   0.3631   0.00628   0.00161  -0.0534  0.8048  1.0000
   1.500   0.4151   0.00662   0.00165  -0.0518  0.7390  1.0000
   2.000   0.4679   0.00702   0.00175  -0.0506  0.6694  1.0000
   2.500   0.5210   0.00750   0.00193  -0.0497  0.5944  1.0000
   3.000   0.5741   0.00806   0.00220  -0.0489  0.5152  1.0000
   3.500   0.6272   0.00870   0.00255  -0.0483  0.4338  1.0000
   4.000   0.6803   0.00943   0.00302  -0.0477  0.3559  1.0000
   4.500   0.7331   0.01023   0.00358  -0.0473  0.2834  1.0000
   5.000   0.7857   0.01111   0.00428  -0.0468  0.2163  1.0000
   5.500   0.8377   0.01213   0.00511  -0.0463  0.1549  1.0000
   6.000   0.8888   0.01335   0.00618  -0.0457  0.0999  1.0000
   6.500   0.9387   0.01488   0.00761  -0.0449  0.0586  1.0000
   7.000   0.9861   0.01696   0.00970  -0.0437  0.0328  1.0000
   7.500   1.0320   0.01938   0.01234  -0.0422  0.0207  1.0000
   8.000   1.0741   0.02263   0.01589  -0.0402  0.0153  1.0000
   8.500   1.1128   0.02672   0.02032  -0.0380  0.0131  1.0000
   9.500   1.1492   0.04305   0.03842  -0.0318  0.0123  1.0000
  10.000   1.1513   0.04952   0.04570  -0.0291  0.0117  1.0000
  10.500   1.1213   0.05792   0.05463  -0.0277  0.0117  1.0000
  11.000   1.0794   0.07138   0.06841  -0.0362  0.0122  1.0000