XFOIL         Version 6.94
  
 Calculated polar for: GOE 174 (ALBATROS 5020) 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   0.5934   0.01043   0.00285  -0.1123  0.5861  0.0288
   0.500   0.6484   0.01027   0.00266  -0.1121  0.5651  0.0344
   1.000   0.7031   0.01012   0.00247  -0.1118  0.5409  0.0471
   1.500   0.7575   0.01010   0.00238  -0.1114  0.5136  0.0591
   3.500   0.9656   0.00973   0.00327  -0.1090  0.3849  1.0000
   4.000   1.0155   0.01039   0.00359  -0.1082  0.3386  1.0000
   4.500   1.0630   0.01131   0.00411  -0.1070  0.2690  1.0000
   5.000   1.1027   0.01314   0.00510  -0.1049  0.1349  1.0000
   5.500   1.1417   0.01505   0.00651  -0.1024  0.0392  1.0000
   6.000   1.1871   0.01613   0.00749  -0.1008  0.0048  1.0000
   6.500   1.2350   0.01688   0.00835  -0.0995  0.0052  1.0000
   7.000   1.2818   0.01769   0.00929  -0.0980  0.0060  1.0000
   7.500   1.3267   0.01863   0.01038  -0.0961  0.0071  1.0000
   8.000   1.3678   0.01984   0.01180  -0.0936  0.0082  1.0000
   8.500   1.4044   0.02130   0.01345  -0.0904  0.0089  1.0000
   9.000   1.4289   0.02341   0.01579  -0.0855  0.0097  1.0000
   9.500   1.4309   0.02642   0.01902  -0.0777  0.0102  1.0000
  10.000   1.4417   0.02923   0.02207  -0.0723  0.0111  1.0000
  10.500   1.4331   0.03426   0.02734  -0.0671  0.0121  1.0000
  11.000   1.4326   0.03937   0.03269  -0.0637  0.0134  1.0000
  11.500   1.4315   0.04475   0.03811  -0.0588  0.0153  1.0000