XFOIL         Version 6.94
  
 Calculated polar for: AG04                                            
  
 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   0.2048   0.00597   0.00176  -0.0418  0.9295  1.0000
   0.500   0.2580   0.00606   0.00163  -0.0402  0.8683  1.0000
   1.000   0.3078   0.00630   0.00157  -0.0379  0.7996  1.0000
   1.500   0.3590   0.00665   0.00159  -0.0362  0.7278  1.0000
   2.000   0.4113   0.00706   0.00169  -0.0349  0.6547  1.0000
   2.500   0.4641   0.00753   0.00186  -0.0340  0.5782  1.0000
   3.000   0.5171   0.00808   0.00212  -0.0332  0.4988  1.0000
   3.500   0.5701   0.00871   0.00245  -0.0326  0.4195  1.0000
   4.000   0.6232   0.00939   0.00289  -0.0321  0.3452  1.0000
   4.500   0.6762   0.01013   0.00342  -0.0316  0.2782  1.0000
   5.000   0.7289   0.01094   0.00407  -0.0312  0.2171  1.0000
   5.500   0.7812   0.01185   0.00484  -0.0307  0.1599  1.0000
   6.000   0.8327   0.01297   0.00581  -0.0302  0.1055  1.0000
   6.500   0.8825   0.01450   0.00722  -0.0294  0.0625  1.0000
   7.000   0.9311   0.01630   0.00908  -0.0284  0.0385  1.0000
   9.000   1.0932   0.03096   0.02503  -0.0210  0.0137  1.0000
   9.500   1.1171   0.03744   0.03243  -0.0182  0.0132  1.0000
  10.000   1.1192   0.04570   0.04160  -0.0154  0.0130  1.0000
  10.500   1.0962   0.05405   0.05059  -0.0134  0.0128  1.0000
  11.000   1.0594   0.06532   0.06227  -0.0197  0.0131  1.0000
  11.500   1.0208   0.08224   0.07940  -0.0332  0.0137  1.0000