XFOIL         Version 6.94
  
 Calculated polar for: GOE 600 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   0.2167   0.00896   0.00353  -0.0476  0.7622  0.7846
   0.500   0.2684   0.00884   0.00361  -0.0461  0.7495  0.8514
   1.000   0.3223   0.00888   0.00379  -0.0447  0.7368  0.9269
   1.500   0.3958   0.00899   0.00384  -0.0481  0.7226  0.9708
   2.000   0.4837   0.00908   0.00395  -0.0549  0.7048  0.9929
   2.500   0.5474   0.00903   0.00376  -0.0567  0.6744  1.0000
   3.000   0.5919   0.00898   0.00346  -0.0542  0.6194  1.0000
   3.500   0.6337   0.00936   0.00334  -0.0514  0.5023  1.0000
   4.500   0.7147   0.01155   0.00436  -0.0467  0.2883  1.0000
   5.000   0.7571   0.01272   0.00504  -0.0449  0.1971  1.0000
   5.500   0.7928   0.01456   0.00614  -0.0423  0.0775  1.0000
   6.000   0.8346   0.01584   0.00714  -0.0404  0.0266  1.0000
   6.500   0.8778   0.01695   0.00819  -0.0385  0.0058  1.0000
   7.000   0.9226   0.01786   0.00921  -0.0369  0.0054  1.0000
   7.500   0.9651   0.01890   0.01039  -0.0349  0.0053  1.0000
   8.000   1.0047   0.02009   0.01176  -0.0326  0.0054  1.0000
   8.500   1.0408   0.02142   0.01327  -0.0297  0.0056  1.0000
   9.000   1.0698   0.02294   0.01499  -0.0258  0.0058  1.0000
   9.500   1.0938   0.02491   0.01717  -0.0218  0.0061  1.0000
  10.000   1.1110   0.02751   0.01999  -0.0177  0.0064  1.0000
  10.500   1.1214   0.03087   0.02355  -0.0138  0.0067  1.0000
  11.000   1.1264   0.03503   0.02787  -0.0102  0.0070  1.0000
  11.500   1.1416   0.03857   0.03163  -0.0079  0.0073  1.0000
  12.000   1.1518   0.04301   0.03636  -0.0050  0.0081  1.0000
  12.500   1.1625   0.04800   0.04158  -0.0019  0.0088  1.0000
  14.500   1.0937   0.08532   0.08107   0.0002  0.0137  1.0000