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 =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.500   0.5599   0.01343   0.00605  -0.0890  0.7219  0.0391
   2.000   0.7245   0.01183   0.00398  -0.0878  0.5992  0.0521
   3.500   0.8833   0.01078   0.00395  -0.0868  0.4282  1.0000
   4.500   0.9826   0.01283   0.00497  -0.0857  0.2772  1.0000
   5.000   1.0322   0.01378   0.00550  -0.0852  0.2262  1.0000
   5.500   1.0824   0.01460   0.00607  -0.0848  0.1884  1.0000
   6.000   1.1331   0.01530   0.00668  -0.0843  0.1648  1.0000
   6.500   1.1698   0.01786   0.00843  -0.0821  0.0302  1.0000
   7.000   1.2159   0.01902   0.00954  -0.0809  0.0059  1.0000
   7.500   1.2624   0.02004   0.01073  -0.0796  0.0052  1.0000
   8.000   1.3070   0.02116   0.01205  -0.0780  0.0054  1.0000
   8.500   1.3483   0.02251   0.01364  -0.0759  0.0057  1.0000
   9.000   1.3833   0.02429   0.01570  -0.0729  0.0061  1.0000
   9.500   1.4137   0.02615   0.01783  -0.0692  0.0067  1.0000
  10.000   1.4266   0.02875   0.02074  -0.0630  0.0074  1.0000
  10.500   1.4228   0.03258   0.02485  -0.0561  0.0079  1.0000
  11.000   1.4241   0.03678   0.02933  -0.0515  0.0085  1.0000
  11.500   1.4141   0.04276   0.03563  -0.0476  0.0093  1.0000
  12.000   1.4015   0.04958   0.04262  -0.0444  0.0100  1.0000