XFOIL         Version 6.94
  
 Calculated polar for: GOE 328 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   0.4112   0.01181   0.00479  -0.0815  0.6794  0.0438
   0.500   0.4681   0.01083   0.00365  -0.0813  0.6578  0.0438
   1.000   0.5249   0.01032   0.00301  -0.0813  0.6339  0.0460
   1.500   0.5814   0.00994   0.00262  -0.0814  0.6064  0.0511
   2.000   0.6379   0.00974   0.00237  -0.0816  0.5736  0.0587
   2.500   0.6866   0.00766   0.00237  -0.0804  0.5337  1.0000
   3.000   0.7394   0.00836   0.00249  -0.0800  0.4478  1.0000
   3.500   0.7923   0.00910   0.00277  -0.0798  0.3781  1.0000
   4.000   0.8465   0.00964   0.00305  -0.0797  0.3281  1.0000
   4.500   0.8967   0.01077   0.00357  -0.0793  0.2307  1.0000
   5.000   0.9387   0.01321   0.00499  -0.0780  0.0367  1.0000
   5.500   0.9894   0.01414   0.00576  -0.0774  0.0052  1.0000
   6.000   1.0413   0.01482   0.00651  -0.0768  0.0048  1.0000
   6.500   1.0921   0.01559   0.00738  -0.0761  0.0049  1.0000
   7.000   1.1413   0.01651   0.00842  -0.0751  0.0051  1.0000
   7.500   1.1882   0.01764   0.00971  -0.0738  0.0055  1.0000
   8.000   1.2303   0.01919   0.01145  -0.0718  0.0058  1.0000
   8.500   1.2724   0.02055   0.01294  -0.0698  0.0065  1.0000
   9.000   1.3029   0.02278   0.01537  -0.0663  0.0073  1.0000
   9.500   1.3212   0.02550   0.01823  -0.0612  0.0079  1.0000
  10.000   1.3334   0.02822   0.02122  -0.0553  0.0090  1.0000
  10.500   1.3393   0.03224   0.02537  -0.0500  0.0100  1.0000
  11.000   1.2395   0.02583   0.01962  -0.0361  0.0101  1.0000