XFOIL         Version 6.94
  
 Calculated polar for: GOE 457 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   0.4782   0.00872   0.00173  -0.0945  0.6481  0.0739
   0.500   0.5157   0.00775   0.00191  -0.0912  0.5772  0.5945
   2.500   0.7424   0.00838   0.00252  -0.0915  0.4592  1.0000
   3.000   0.7920   0.00870   0.00270  -0.0901  0.4339  1.0000
   3.500   0.8360   0.00928   0.00278  -0.0877  0.3478  1.0000
   4.000   0.8764   0.01046   0.00329  -0.0850  0.2252  1.0000
   4.500   0.9129   0.01226   0.00428  -0.0819  0.0833  1.0000
   5.000   0.9534   0.01367   0.00528  -0.0792  0.0065  1.0000
   5.500   1.0013   0.01424   0.00591  -0.0775  0.0062  1.0000
   6.000   1.0484   0.01487   0.00664  -0.0756  0.0071  1.0000
   6.500   1.0948   0.01554   0.00744  -0.0736  0.0086  1.0000
   7.000   1.1384   0.01644   0.00852  -0.0710  0.0094  1.0000
   7.500   1.1792   0.01754   0.00983  -0.0679  0.0100  1.0000
   8.000   1.2136   0.01909   0.01159  -0.0639  0.0105  1.0000
   8.500   1.2352   0.02126   0.01388  -0.0579  0.0110  1.0000
   9.000   1.2596   0.02271   0.01550  -0.0521  0.0121  1.0000
   9.500   1.2628   0.02550   0.01834  -0.0437  0.0132  1.0000
  10.000   1.2774   0.02842   0.02132  -0.0378  0.0132  1.0000
  10.500   1.2957   0.03105   0.02420  -0.0328  0.0124  1.0000
  11.000   1.3222   0.03548   0.02870  -0.0299  0.0105  1.0000
  11.500   1.3606   0.04236   0.03568  -0.0298  0.0069  1.0000
  12.000   1.3229   0.04507   0.03895  -0.0211  0.0064  1.0000