XFOIL         Version 6.94
  
 Calculated polar for: GOE 286 AIRFOIL                                 
  
 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   0.3870   0.01081   0.00309  -0.0654  0.5453  0.0419
   0.500   0.4438   0.01050   0.00259  -0.0654  0.4987  0.0417
   1.000   0.4998   0.01055   0.00233  -0.0654  0.4462  0.0421
   1.500   0.5557   0.01075   0.00227  -0.0653  0.4037  0.0480
   2.000   0.6086   0.00873   0.00250  -0.0652  0.3732  1.0000
   2.500   0.6643   0.00908   0.00265  -0.0652  0.3482  1.0000
   3.000   0.7198   0.00944   0.00285  -0.0652  0.3230  1.0000
   3.500   0.7749   0.00985   0.00311  -0.0652  0.2969  1.0000
   4.000   0.8294   0.01033   0.00340  -0.0651  0.2611  1.0000
   4.500   0.8745   0.01235   0.00450  -0.0643  0.1161  1.0000
   5.000   0.9270   0.01304   0.00506  -0.0640  0.0963  1.0000
   5.500   0.9739   0.01454   0.00602  -0.0630  0.0053  1.0000
   6.000   1.0260   0.01514   0.00667  -0.0626  0.0038  1.0000
   6.500   1.0772   0.01583   0.00745  -0.0619  0.0037  1.0000
   7.000   1.1274   0.01658   0.00838  -0.0612  0.0037  1.0000
   7.500   1.1761   0.01744   0.00938  -0.0602  0.0039  1.0000
   8.000   1.2229   0.01846   0.01052  -0.0589  0.0042  1.0000
   8.500   1.2670   0.01965   0.01190  -0.0572  0.0047  1.0000
   9.000   1.3072   0.02110   0.01358  -0.0550  0.0052  1.0000
   9.500   1.3381   0.02317   0.01597  -0.0517  0.0056  1.0000
  10.000   1.3562   0.02563   0.01869  -0.0468  0.0061  1.0000
  10.500   1.3711   0.02806   0.02131  -0.0421  0.0068  1.0000
  11.000   1.3823   0.03124   0.02473  -0.0383  0.0078  1.0000
  11.500   1.3855   0.03582   0.02961  -0.0356  0.0091  1.0000
  12.000   1.3791   0.04213   0.03622  -0.0340  0.0102  1.0000