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 =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.3135   0.01244   0.00748  -0.0870  0.9197  1.0000
   0.500   0.3633   0.01263   0.00753  -0.0865  0.9118  1.0000
   1.000   0.4372   0.01245   0.00734  -0.0908  0.9088  1.0000
   1.500   0.6180   0.00811   0.00190  -0.1117  0.6534  1.0000
   2.000   0.5679   0.01223   0.00311  -0.0899  0.0158  1.0000