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 =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.4768   0.01167   0.00295  -0.0941  0.5894  0.0394
   0.500   0.5208   0.01131   0.00276  -0.0917  0.5670  0.1542
   1.000   0.7157   0.01033   0.00346  -0.1240  0.5415  1.0000
   1.500   0.7627   0.01063   0.00359  -0.1224  0.5284  1.0000
   2.000   0.8105   0.01093   0.00378  -0.1210  0.5180  1.0000
   2.500   0.8470   0.01114   0.00390  -0.1170  0.4918  1.0000
   3.000   0.8691   0.01145   0.00383  -0.1098  0.4292  1.0000
   3.500   0.9068   0.01178   0.00407  -0.1062  0.3987  1.0000
   4.000   0.8953   0.01480   0.00539  -0.0936  0.0790  1.0000
   4.500   0.9234   0.01579   0.00616  -0.0881  0.0078  1.0000
   5.000   0.9587   0.01630   0.00688  -0.0840  0.0078  1.0000
   5.500   0.9898   0.01688   0.00765  -0.0788  0.0083  1.0000
   6.000   1.0150   0.01775   0.00885  -0.0724  0.0094  1.0000
   6.500   1.0323   0.01907   0.01051  -0.0647  0.0104  1.0000
   7.000   1.0474   0.02055   0.01220  -0.0572  0.0117  1.0000
   7.500   1.0497   0.02275   0.01465  -0.0480  0.0136  1.0000
   8.000   1.0412   0.02590   0.01800  -0.0383  0.0158  1.0000
   8.500   1.0381   0.02989   0.02205  -0.0306  0.0177  1.0000