XFOIL         Version 6.94
  
 Calculated polar for: EPPLER 904 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   0.2794   0.01208   0.00726  -0.0666  0.9381  0.9855
   0.500   0.3705   0.01150   0.00672  -0.0747  0.9337  0.9901
   1.000   0.4586   0.01057   0.00591  -0.0813  0.9224  0.9931
   1.500   0.5306   0.01008   0.00558  -0.0851  0.9117  0.9973
   2.000   0.5920   0.00975   0.00552  -0.0865  0.8986  1.0000
   2.500   0.6285   0.00881   0.00451  -0.0797  0.8316  1.0000
   3.000   0.6060   0.01139   0.00431  -0.0626  0.2434  1.0000