XFOIL         Version 6.94
  
 Calculated polar for: GOE 360 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   0.5588   0.01686   0.00949  -0.0993  0.5975  0.0691
   1.500   0.7261   0.01361   0.00530  -0.0974  0.5352  0.0582
   2.000   0.7799   0.01300   0.00458  -0.0969  0.5118  0.0564
   2.500   0.8332   0.01275   0.00424  -0.0965  0.4891  0.0607
   3.000   0.8864   0.01267   0.00410  -0.0962  0.4675  0.0664
   3.500   0.9398   0.01276   0.00415  -0.0959  0.4476  0.0928
   5.000   1.0881   0.01258   0.00502  -0.0935  0.3535  1.0000
   5.500   1.1370   0.01321   0.00545  -0.0928  0.3189  1.0000
   6.000   1.1859   0.01386   0.00600  -0.0920  0.2949  1.0000
   6.500   1.2323   0.01471   0.00665  -0.0910  0.2549  1.0000
   7.000   1.2425   0.01880   0.00945  -0.0856  0.0483  1.0000
   7.500   1.2739   0.02068   0.01123  -0.0825  0.0055  1.0000
   8.000   1.3105   0.02190   0.01260  -0.0801  0.0053  1.0000
   8.500   1.3397   0.02333   0.01422  -0.0766  0.0054  1.0000
   9.000   1.3647   0.02517   0.01627  -0.0732  0.0056  1.0000
   9.500   1.3840   0.02762   0.01898  -0.0698  0.0059  1.0000
  10.000   1.3960   0.03092   0.02254  -0.0666  0.0062  1.0000
  10.500   1.3980   0.03552   0.02742  -0.0637  0.0064  1.0000
  11.000   1.3962   0.04109   0.03324  -0.0621  0.0066  1.0000
  11.500   1.3946   0.04721   0.03964  -0.0614  0.0070  1.0000
  12.000   1.3810   0.05515   0.04784  -0.0613  0.0074  1.0000
  12.500   1.3631   0.06388   0.05682  -0.0617  0.0077  1.0000
  13.000   1.3448   0.07290   0.06604  -0.0623  0.0080  1.0000
  13.500   1.3306   0.08109   0.07432  -0.0624  0.0083  1.0000
  14.000   1.3362   0.08698   0.08045  -0.0622  0.0091  1.0000
  14.500   1.3589   0.08797   0.08138  -0.0568  0.0107  1.0000