XFOIL         Version 6.94
  
 Calculated polar for: GOE 264 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   0.5775   0.01179   0.00411  -0.1151  0.6649  0.0734
   0.500   0.6332   0.01151   0.00363  -0.1150  0.6283  0.0801
   1.000   0.6890   0.01141   0.00343  -0.1150  0.5970  0.0895
   2.000   0.8013   0.01137   0.00344  -0.1154  0.5379  0.1644
   2.500   0.8573   0.01135   0.00361  -0.1158  0.5096  0.2769
   3.500   0.9538   0.01108   0.00389  -0.1131  0.3896  1.0000
   4.000   1.0060   0.01183   0.00428  -0.1127  0.3269  1.0000
   4.500   1.0570   0.01276   0.00484  -0.1122  0.2561  1.0000
   5.000   1.0952   0.01584   0.00669  -0.1104  0.0304  1.0000
   5.500   1.1451   0.01690   0.00776  -0.1094  0.0079  1.0000
   6.000   1.1952   0.01784   0.00889  -0.1082  0.0086  1.0000
   6.500   1.2421   0.01918   0.01050  -0.1065  0.0099  1.0000
   7.000   1.2834   0.02110   0.01273  -0.1038  0.0117  1.0000
   7.500   1.3105   0.02415   0.01605  -0.0989  0.0139  1.0000
   8.000   1.3268   0.02817   0.02028  -0.0922  0.0165  1.0000
  10.000   1.4307   0.05158   0.04653  -0.0664  0.0504  1.0000
  10.500   1.4091   0.05885   0.05424  -0.0588  0.0448  1.0000
  11.000   1.3899   0.06736   0.06294  -0.0528  0.0412  1.0000
  11.500   1.3337   0.07353   0.06950  -0.0440  0.0410  1.0000
  12.000   1.2836   0.08214   0.07843  -0.0400  0.0412  1.0000
  12.500   1.2344   0.09197   0.08854  -0.0390  0.0413  1.0000
  13.000   1.1869   0.10230   0.09911  -0.0401  0.0415  1.0000
  13.500   1.1399   0.11184   0.10884  -0.0419  0.0416  1.0000
  14.000   1.0926   0.12106   0.11825  -0.0450  0.0417  1.0000