XFOIL         Version 6.94
  
 Calculated polar for: GOE 585 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   0.3758   0.00725   0.00172  -0.0752  0.7731  0.4283
   1.000   0.5282   0.00629   0.00177  -0.0832  0.6942  1.0000
   1.500   0.5747   0.00656   0.00181  -0.0809  0.6502  1.0000
   2.000   0.6195   0.00692   0.00189  -0.0782  0.5878  1.0000
   2.500   0.6606   0.00754   0.00207  -0.0750  0.4909  1.0000
   3.000   0.6958   0.00874   0.00246  -0.0710  0.3536  1.0000
   3.500   0.7427   0.00923   0.00278  -0.0691  0.3143  1.0000
   4.000   0.7906   0.00969   0.00309  -0.0674  0.2830  1.0000
   4.500   0.8341   0.01054   0.00346  -0.0652  0.2044  1.0000
   5.000   0.8810   0.01116   0.00389  -0.0635  0.1678  1.0000
   5.500   0.9260   0.01198   0.00450  -0.0616  0.1174  1.0000
   6.000   0.9595   0.01408   0.00588  -0.0579  0.0064  1.0000
   6.500   1.0076   0.01464   0.00648  -0.0564  0.0053  1.0000
   7.000   1.0541   0.01535   0.00729  -0.0545  0.0051  1.0000
   7.500   1.0993   0.01618   0.00828  -0.0525  0.0051  1.0000
   8.000   1.1428   0.01715   0.00948  -0.0502  0.0052  1.0000
   8.500   1.1832   0.01837   0.01102  -0.0474  0.0054  1.0000
   9.000   1.2156   0.02021   0.01317  -0.0436  0.0050  1.0000
   9.500   1.2394   0.02246   0.01570  -0.0386  0.0044  1.0000
  10.000   1.2440   0.02540   0.01897  -0.0309  0.0047  1.0000
  10.500   1.2463   0.02868   0.02252  -0.0239  0.0047  1.0000
  11.000   1.2584   0.03142   0.02547  -0.0193  0.0050  1.0000
  11.500   1.2675   0.03470   0.02898  -0.0153  0.0054  1.0000
  12.000   1.2679   0.03962   0.03435  -0.0114  0.0066  1.0000
  12.500   1.2626   0.04579   0.04094  -0.0087  0.0076  1.0000
  13.000   1.2398   0.05498   0.05061  -0.0076  0.0088  1.0000