XFOIL         Version 6.94
  
 Calculated polar for: GOE 6K 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   0.4142   0.00826   0.00410  -0.1064  0.9198  0.9743
   0.500   0.5210   0.00797   0.00379  -0.1185  0.9195  1.0000
   1.000   0.5929   0.00686   0.00153  -0.1187  0.6820  1.0000
   1.500   0.5551   0.01010   0.00245  -0.0991  0.1599  1.0000