XFOIL         Version 6.94
  
 Calculated polar for: AG36                                            
  
 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   0.4037   0.00737   0.00189  -0.0461  0.6556  1.0000
   0.500   0.4565   0.00780   0.00198  -0.0453  0.6114  1.0000
   1.000   0.5102   0.00824   0.00222  -0.0449  0.5775  1.0000
   1.500   0.5643   0.00865   0.00250  -0.0445  0.5473  1.0000
   2.000   0.6187   0.00898   0.00279  -0.0443  0.5161  1.0000
   2.500   0.6727   0.00933   0.00309  -0.0440  0.4825  1.0000
   3.000   0.7268   0.00964   0.00338  -0.0437  0.4407  1.0000
   3.500   0.7797   0.01011   0.00376  -0.0433  0.3887  1.0000
   4.000   0.8312   0.01081   0.00429  -0.0428  0.3283  1.0000
   4.500   0.8818   0.01169   0.00498  -0.0424  0.2651  1.0000
   5.000   0.9313   0.01274   0.00581  -0.0419  0.1981  1.0000
   5.500   0.9799   0.01391   0.00680  -0.0413  0.1421  1.0000
   6.000   1.0268   0.01530   0.00800  -0.0405  0.0877  1.0000
   6.500   1.0678   0.01761   0.01005  -0.0389  0.0395  1.0000
   7.500   1.1433   0.02247   0.01518  -0.0346  0.0217  1.0000
   8.500   1.2066   0.02848   0.02163  -0.0291  0.0167  1.0000
   9.000   1.2302   0.03372   0.02713  -0.0259  0.0156  1.0000
   9.500   1.2497   0.03818   0.03216  -0.0223  0.0152  1.0000
  10.000   1.2501   0.04323   0.03781  -0.0173  0.0148  1.0000
  10.500   1.2350   0.04915   0.04426  -0.0132  0.0145  1.0000
  11.000   1.2105   0.05710   0.05271  -0.0128  0.0143  1.0000
  11.500   1.1787   0.06778   0.06386  -0.0167  0.0143  1.0000
  12.000   1.1409   0.08153   0.07801  -0.0245  0.0145  1.0000
  12.500   1.0989   0.09776   0.09455  -0.0343  0.0149  1.0000
  13.000   1.0558   0.11626   0.11329  -0.0454  0.0153  1.0000