XFOIL         Version 6.94
  
 Calculated polar for: GOE 574 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   0.4634   0.01044   0.00220  -0.0905  0.5491  0.0238
   0.500   0.5108   0.01057   0.00228  -0.0891  0.5286  0.0407
   1.500   0.7639   0.00941   0.00312  -0.1227  0.4945  1.0000
   2.000   0.8103   0.00967   0.00328  -0.1210  0.4858  1.0000
   2.500   0.8393   0.00997   0.00325  -0.1154  0.4206  1.0000
   3.000   0.8789   0.01030   0.00338  -0.1122  0.3850  1.0000
   3.500   0.8778   0.01312   0.00465  -0.1015  0.0790  1.0000
   4.000   0.9095   0.01393   0.00527  -0.0967  0.0056  1.0000
   4.500   0.9468   0.01438   0.00580  -0.0929  0.0059  1.0000
   5.000   0.9823   0.01489   0.00649  -0.0887  0.0069  1.0000
   5.500   1.0130   0.01547   0.00730  -0.0833  0.0085  1.0000
   6.000   1.0342   0.01642   0.00852  -0.0758  0.0105  1.0000
   6.500   1.0513   0.01762   0.00995  -0.0678  0.0140  1.0000
   7.000   1.0105   0.00995   0.00279  -0.0508  0.0160  1.0000
   7.500   1.0103   0.01169   0.00476  -0.0415  0.0201  1.0000