XFOIL         Version 6.94
  
 Calculated polar for: GOE 566 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   0.2722   0.00596   0.00188  -0.0522  0.7856  0.8963
   0.500   0.4114   0.00606   0.00187  -0.0697  0.7567  1.0000
   1.000   0.4579   0.00614   0.00177  -0.0676  0.7256  1.0000
   2.000   0.5494   0.00660   0.00176  -0.0630  0.6300  1.0000
   2.500   0.5754   0.00823   0.00206  -0.0573  0.3610  1.0000
   3.000   0.6175   0.00908   0.00241  -0.0549  0.2671  1.0000
   3.500   0.6642   0.00956   0.00275  -0.0531  0.2344  1.0000
   4.000   0.7001   0.01108   0.00338  -0.0499  0.0834  1.0000
   4.500   0.7440   0.01193   0.00396  -0.0478  0.0328  1.0000
   5.000   0.7916   0.01245   0.00446  -0.0462  0.0265  1.0000
   5.500   0.8370   0.01324   0.00512  -0.0443  0.0071  1.0000
   6.000   0.8857   0.01368   0.00567  -0.0429  0.0050  1.0000
   6.500   0.9329   0.01430   0.00637  -0.0413  0.0048  1.0000
   7.000   0.9789   0.01507   0.00726  -0.0394  0.0048  1.0000
   7.500   1.0235   0.01596   0.00835  -0.0373  0.0049  1.0000
   8.000   1.0661   0.01705   0.00968  -0.0350  0.0051  1.0000
   8.500   1.1048   0.01843   0.01137  -0.0321  0.0054  1.0000
   9.000   1.1384   0.02015   0.01334  -0.0285  0.0057  1.0000
   9.500   1.1652   0.02221   0.01567  -0.0241  0.0062  1.0000
  10.000   1.1792   0.02480   0.01856  -0.0177  0.0068  1.0000
  10.500   1.1874   0.02815   0.02225  -0.0113  0.0074  1.0000
  11.000   1.1903   0.03319   0.02774  -0.0054  0.0083  1.0000
  11.500   1.1868   0.03897   0.03399  -0.0004  0.0088  1.0000