XFOIL         Version 6.94
  
 Calculated polar for: GOE 199 (L.F.G. 5406) 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   0.5013   0.01215   0.00534  -0.0874  0.7107  0.0247
   1.500   0.6531   0.00327  -0.00392  -0.0843  0.5637  0.0449
   2.000   0.7067   0.00309  -0.00433  -0.0840  0.5181  0.0431
   2.500   0.7601   0.00301  -0.00453  -0.0839  0.4739  0.0445
   3.000   0.8125   0.00310  -0.00466  -0.0837  0.4156  0.0495
   3.500   0.8634   0.00334  -0.00471  -0.0833  0.3446  0.0702
   5.000   1.0099   0.00372  -0.00371  -0.0819  0.1839  1.0000
   5.500   1.0571   0.00427  -0.00349  -0.0811  0.1342  1.0000
   6.000   1.0946   0.00573  -0.00260  -0.0789  0.0202  1.0000
   6.500   1.1409   0.00638  -0.00195  -0.0778  0.0042  1.0000
   7.000   1.1872   0.00697  -0.00125  -0.0766  0.0043  1.0000
   7.500   1.2313   0.00773  -0.00032  -0.0750  0.0047  1.0000
   8.000   1.2737   0.00855   0.00065  -0.0731  0.0054  1.0000
   8.500   1.3112   0.00969   0.00201  -0.0704  0.0063  1.0000
   9.000   1.3449   0.01091   0.00344  -0.0671  0.0074  1.0000
   9.500   1.3638   0.01272   0.00550  -0.0616  0.0084  1.0000
  10.000   1.3610   0.01521   0.00830  -0.0535  0.0093  1.0000
  10.500   1.3438   0.01946   0.01285  -0.0460  0.0098  1.0000
  11.000   1.3257   0.02532   0.01896  -0.0407  0.0102  1.0000
  11.500   1.3058   0.03300   0.02698  -0.0362  0.0116  1.0000
  12.000   1.2946   0.04105   0.03509  -0.0314  0.0125  1.0000