XFOIL         Version 6.94
  
 Calculated polar for: GOE 442 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
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.4793   0.00935   0.00231  -0.0921  0.7792  0.0722
   0.500   0.5499   0.00735   0.00233  -0.0953  0.6933  1.0000
   1.000   0.5943   0.00794   0.00233  -0.0924  0.6109  1.0000
   1.500   0.6409   0.00850   0.00243  -0.0902  0.5590  1.0000
   2.000   0.6897   0.00901   0.00263  -0.0886  0.5245  1.0000
   2.500   0.7390   0.00948   0.00288  -0.0871  0.4943  1.0000
   3.000   0.7885   0.00987   0.00313  -0.0857  0.4647  1.0000
   3.500   0.8385   0.01027   0.00346  -0.0844  0.4406  1.0000
   4.000   0.8864   0.01064   0.00368  -0.0827  0.3997  1.0000
   4.500   0.9254   0.01170   0.00397  -0.0798  0.2647  1.0000
   5.000   0.9671   0.01297   0.00466  -0.0776  0.1652  1.0000
   5.500   1.0003   0.01524   0.00593  -0.0744  0.0345  1.0000
   6.000   1.0428   0.01649   0.00701  -0.0719  0.0051  1.0000
   6.500   1.0897   0.01716   0.00794  -0.0703  0.0047  1.0000
   7.000   1.1362   0.01786   0.00881  -0.0685  0.0049  1.0000
   7.500   1.1805   0.01876   0.00985  -0.0663  0.0053  1.0000
   8.000   1.2219   0.01991   0.01125  -0.0636  0.0057  1.0000
   8.500   1.2594   0.02137   0.01309  -0.0602  0.0062  1.0000
   9.000   1.2909   0.02318   0.01533  -0.0562  0.0070  1.0000
   9.500   1.3123   0.02528   0.01777  -0.0506  0.0082  1.0000
  10.000   1.3190   0.02814   0.02096  -0.0433  0.0095  1.0000
  10.500   1.3162   0.03206   0.02517  -0.0363  0.0106  1.0000
  11.000   1.3122   0.03712   0.03048  -0.0307  0.0114  1.0000
  11.500   1.3316   0.04087   0.03454  -0.0269  0.0138  1.0000
  12.000   1.3457   0.04888   0.04312  -0.0230  0.0175  1.0000
  12.500   1.3431   0.06479   0.06000  -0.0201  0.0269  1.0000