XFOIL         Version 6.94
  
 Calculated polar for: agora10                                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.045 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
  -3.000  -0.0759   0.03692   0.02756  -0.0153  0.8297  0.3391
  -2.750  -0.0497   0.03635   0.02712  -0.0147  0.8128  0.3491
  -2.500  -0.0264   0.03667   0.02739  -0.0128  0.7936  0.3590
  -2.250  -0.0049   0.03744   0.02796  -0.0101  0.7695  0.3689
  -2.000   0.0176   0.03768   0.02833  -0.0079  0.7453  0.3807
  -1.750   0.0405   0.03839   0.02903  -0.0057  0.7195  0.3917
  -1.500   0.0649   0.03912   0.02975  -0.0039  0.6877  0.4012
  -1.250   0.0822   0.04049   0.03126  -0.0002  0.6446  0.4126
  -1.000   0.0962   0.04302   0.03379   0.0045  0.5839  0.4270
  -0.750   0.1098   0.04691   0.03783   0.0082  0.4973  0.4462
  -0.250   0.1729   0.04510   0.03815   0.0075  0.3765  1.0008
   0.000   0.2072   0.04418   0.03668   0.0082  0.3537  1.0008
   0.250   0.2332   0.04298   0.03509   0.0112  0.3387  1.0008
   0.500   0.2557   0.04119   0.03293   0.0155  0.3284  1.0008
   0.750   0.2803   0.03999   0.03139   0.0187  0.3183  1.0008
   1.000   0.3039   0.03875   0.02975   0.0223  0.3090  1.0008
   1.250   0.3276   0.03783   0.02837   0.0256  0.2990  1.0008
   1.500   0.3557   0.03816   0.02839   0.0267  0.2867  1.0008
   1.750   0.3845   0.03885   0.02886   0.0274  0.2776  1.0008
   2.000   0.4148   0.04002   0.02990   0.0273  0.2697  1.0008
   2.250   0.4455   0.04151   0.03132   0.0268  0.2612  1.0008
   2.500   0.4772   0.04342   0.03322   0.0256  0.2530  1.0008
   2.750   0.5080   0.04537   0.03516   0.0247  0.2478  1.0008
   3.000   0.5631   0.05642   0.04723   0.0074  0.2442  1.0008
   3.250   0.6070   0.08407   0.07585  -0.0343  0.2631  1.0008