XFOIL         Version 6.94
  
 Calculated polar for: GOE 599 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   0.2232   0.00765   0.00326  -0.0556  0.9021  0.9651
   0.500   0.3103   0.00738   0.00289  -0.0615  0.8694  0.9925
   1.000   0.3764   0.00732   0.00271  -0.0635  0.8360  1.0000
   1.500   0.4255   0.00741   0.00264  -0.0619  0.7979  1.0000
   2.000   0.4725   0.00758   0.00266  -0.0597  0.7458  1.0000
   2.500   0.5023   0.00870   0.00254  -0.0538  0.4960  1.0000
   3.500   0.5736   0.01233   0.00398  -0.0475  0.0548  1.0000
   4.000   0.6217   0.01310   0.00438  -0.0463  0.0081  1.0000
   4.500   0.6716   0.01374   0.00502  -0.0451  0.0062  1.0000
   5.000   0.7212   0.01444   0.00592  -0.0438  0.0059  1.0000
   5.500   0.7706   0.01519   0.00690  -0.0425  0.0061  1.0000
   6.000   0.8187   0.01613   0.00810  -0.0408  0.0064  1.0000
   6.500   0.8642   0.01741   0.00964  -0.0386  0.0068  1.0000
   7.000   0.9051   0.01931   0.01186  -0.0356  0.0074  1.0000
   7.500   0.9408   0.02235   0.01529  -0.0317  0.0080  1.0000
   8.000   0.9780   0.02609   0.01946  -0.0284  0.0088  1.0000
   8.500   1.0151   0.03100   0.02504  -0.0247  0.0105  1.0000
   9.000   1.0123   0.04361   0.03893  -0.0166  0.0132  1.0000