XFOIL         Version 6.94
  
 Calculated polar for: GOE 613 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   0.4753   0.00959   0.00267  -0.0922  0.7295  0.0471
   0.500   0.5234   0.00931   0.00226  -0.0903  0.6976  0.0530
   1.000   0.5680   0.00876   0.00203  -0.0879  0.6593  0.2189
   1.500   0.6787   0.00739   0.00230  -0.1003  0.5879  1.0000
   2.000   0.7159   0.00789   0.00245  -0.0964  0.5365  1.0000
   2.500   0.7564   0.00835   0.00267  -0.0932  0.4987  1.0000
   3.000   0.7979   0.00878   0.00292  -0.0903  0.4633  1.0000
   3.500   0.8239   0.00987   0.00324  -0.0846  0.3389  1.0000
   4.000   0.8610   0.01064   0.00363  -0.0811  0.2806  1.0000
   4.500   0.8669   0.01369   0.00511  -0.0726  0.0070  1.0000
   5.000   0.9112   0.01416   0.00563  -0.0704  0.0047  1.0000
   5.500   0.9547   0.01469   0.00622  -0.0681  0.0047  1.0000
   6.000   0.9974   0.01528   0.00691  -0.0656  0.0049  1.0000
   6.500   1.0383   0.01595   0.00769  -0.0628  0.0055  1.0000
   7.000   1.0752   0.01674   0.00862  -0.0592  0.0062  1.0000
   7.500   1.1081   0.01774   0.00989  -0.0549  0.0070  1.0000
   8.000   1.1340   0.01923   0.01161  -0.0497  0.0081  1.0000
   8.500   1.1616   0.02057   0.01310  -0.0450  0.0096  1.0000
   9.000   1.1801   0.02249   0.01525  -0.0392  0.0121  1.0000
   9.500   1.1753   0.02598   0.01898  -0.0314  0.0139  1.0000