XFOIL         Version 6.94
  
 Calculated polar for: GOE 561 AIRFOIL                                 
  
 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   0.5990   0.02889   0.02094  -0.0711  0.4754  0.0771
   0.500   0.5814   0.02893   0.02107  -0.0579  0.4706  0.0801
   1.000   0.5952   0.02826   0.02024  -0.0505  0.4658  0.0896