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 =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.1969   0.00496   0.00126  -0.0387  0.8601  1.0000
   0.500   0.2481   0.00520   0.00118  -0.0370  0.7919  1.0000
   1.000   0.3006   0.00553   0.00116  -0.0358  0.7228  1.0000
   1.500   0.3541   0.00589   0.00121  -0.0349  0.6534  1.0000
   2.000   0.4080   0.00630   0.00131  -0.0342  0.5824  1.0000
   2.500   0.4620   0.00675   0.00148  -0.0337  0.5082  1.0000
   3.000   0.5161   0.00726   0.00172  -0.0333  0.4335  1.0000
   3.500   0.5701   0.00783   0.00203  -0.0329  0.3620  1.0000
   4.000   0.6240   0.00843   0.00242  -0.0326  0.2970  1.0000
   4.500   0.6778   0.00908   0.00287  -0.0323  0.2380  1.0000
   5.000   0.7312   0.00980   0.00343  -0.0320  0.1837  1.0000
   5.500   0.7842   0.01062   0.00409  -0.0316  0.1330  1.0000
   6.000   0.8364   0.01159   0.00491  -0.0312  0.0853  1.0000
   6.500   0.8877   0.01278   0.00602  -0.0305  0.0496  1.0000
   7.000   0.9380   0.01419   0.00745  -0.0297  0.0295  1.0000
   7.500   0.9852   0.01627   0.00966  -0.0285  0.0189  1.0000
   8.000   1.0312   0.01840   0.01200  -0.0273  0.0135  1.0000
   8.500   1.0763   0.02060   0.01446  -0.0258  0.0109  1.0000
   9.000   1.1107   0.02521   0.01951  -0.0234  0.0096  1.0000
   9.500   1.1490   0.02837   0.02317  -0.0215  0.0083  1.0000
  10.000   1.1787   0.03246   0.02772  -0.0194  0.0075  1.0000
  10.500   1.1916   0.03837   0.03419  -0.0168  0.0071  1.0000
  11.000   1.1763   0.04589   0.04231  -0.0135  0.0070  1.0000
  11.500   1.1469   0.05504   0.05193  -0.0161  0.0070  1.0000
  12.000   1.1147   0.07004   0.06740  -0.0284  0.0072  1.0000