XFOIL         Version 6.94
  
 Calculated polar for: GOE 591 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   0.6036   0.00854   0.00338  -0.1186  0.6845  1.0000
   0.500   0.6485   0.00871   0.00328  -0.1162  0.6615  1.0000
   1.000   0.6933   0.00892   0.00327  -0.1138  0.6380  1.0000
   1.500   0.7379   0.00917   0.00332  -0.1114  0.6135  1.0000
   2.000   0.7822   0.00946   0.00340  -0.1090  0.5887  1.0000
   2.500   0.8275   0.00978   0.00356  -0.1068  0.5677  1.0000
   3.000   0.8730   0.01014   0.00377  -0.1047  0.5479  1.0000
   4.500   0.9834   0.01151   0.00444  -0.0934  0.4178  1.0000
   5.000   1.0159   0.01225   0.00486  -0.0891  0.3588  1.0000
   5.500   1.0406   0.01339   0.00553  -0.0836  0.2772  1.0000
   6.000   1.0619   0.01468   0.00642  -0.0777  0.2087  1.0000
   6.500   1.0655   0.01677   0.00788  -0.0691  0.1096  1.0000
   7.500   1.0949   0.02035   0.01108  -0.0567  0.0062  1.0000
   8.000   1.1230   0.02157   0.01244  -0.0529  0.0063  1.0000
   8.500   1.1475   0.02307   0.01413  -0.0489  0.0066  1.0000
   9.000   1.1708   0.02477   0.01600  -0.0451  0.0070  1.0000
   9.500   1.1912   0.02679   0.01823  -0.0414  0.0074  1.0000
  10.000   1.2054   0.02944   0.02111  -0.0377  0.0081  1.0000
  10.500   1.2121   0.03290   0.02481  -0.0340  0.0085  1.0000
  11.000   1.2242   0.03621   0.02834  -0.0313  0.0094  1.0000
  11.500   1.2194   0.04136   0.03371  -0.0287  0.0101  1.0000
  12.000   1.2134   0.04712   0.03965  -0.0269  0.0107  1.0000
  12.500   1.2136   0.05266   0.04543  -0.0258  0.0118  1.0000
  13.000   1.2042   0.05939   0.05223  -0.0248  0.0126  1.0000
  13.500   1.2167   0.06382   0.05687  -0.0236  0.0145  1.0000
  14.000   1.1005   0.07785   0.07183  -0.0270  0.0143  1.0000