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 =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.3853   0.00738   0.00291  -0.0768  0.8287  1.0000
   0.500   0.4274   0.00733   0.00269  -0.0739  0.8033  1.0000
   1.000   0.4714   0.00733   0.00253  -0.0711  0.7748  1.0000
   1.500   0.5166   0.00741   0.00244  -0.0686  0.7436  1.0000
   2.000   0.5616   0.00754   0.00240  -0.0660  0.7044  1.0000
   2.500   0.6069   0.00776   0.00245  -0.0635  0.6611  1.0000
   4.000   0.7092   0.01100   0.00353  -0.0509  0.2452  1.0000
   4.500   0.7517   0.01181   0.00401  -0.0486  0.1871  1.0000
   5.000   0.7879   0.01326   0.00482  -0.0454  0.0794  1.0000
   5.500   0.8288   0.01434   0.00566  -0.0428  0.0297  1.0000
   6.000   0.8717   0.01526   0.00652  -0.0405  0.0070  1.0000
   6.500   0.9153   0.01609   0.00747  -0.0382  0.0061  1.0000
   7.000   0.9582   0.01698   0.00852  -0.0358  0.0062  1.0000
   7.500   0.9991   0.01803   0.00977  -0.0331  0.0065  1.0000
   8.000   1.0366   0.01937   0.01133  -0.0298  0.0069  1.0000
   8.500   1.0687   0.02101   0.01323  -0.0259  0.0073  1.0000
   9.000   1.0957   0.02278   0.01522  -0.0212  0.0081  1.0000
   9.500   1.1054   0.02521   0.01785  -0.0140  0.0088  1.0000
  10.000   1.1146   0.02818   0.02097  -0.0077  0.0094  1.0000
  10.500   1.1375   0.03271   0.02584  -0.0032  0.0112  1.0000
  11.000   1.1895   0.04476   0.03889  -0.0032  0.0172  1.0000
  12.500   0.9855   0.06409   0.06033   0.0114  0.0210  1.0000
  13.000   0.9493   0.07589   0.07240   0.0072  0.0207  1.0000
  13.500   0.9153   0.08876   0.08548   0.0008  0.0204  1.0000
  14.000   0.8845   0.10205   0.09895  -0.0069  0.0202  1.0000
  15.000   0.8101   0.12938   0.12668  -0.0310  0.0143  1.0000