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 =     0.040 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
  -3.000  -0.0585   0.03624   0.02692  -0.0210  0.8602  0.3667
  -2.750  -0.0345   0.03608   0.02673  -0.0195  0.8392  0.3761
  -2.500  -0.0116   0.03639   0.02700  -0.0175  0.8175  0.3879
  -2.250   0.0106   0.03706   0.02750  -0.0151  0.7928  0.4003
  -2.000   0.0312   0.03751   0.02805  -0.0123  0.7681  0.4119
  -1.750   0.0541   0.03820   0.02874  -0.0100  0.7415  0.4232
  -1.500   0.0770   0.03890   0.02953  -0.0078  0.7082  0.4355
  -1.250   0.0942   0.04041   0.03119  -0.0039  0.6669  0.4516
  -1.000   0.1080   0.04302   0.03400   0.0006  0.6024  0.4720
  -0.750   0.1210   0.04762   0.03883   0.0043  0.5152  0.5042
  -0.500   0.1377   0.04956   0.04189   0.0065  0.4463  0.6395
  -0.250   0.1721   0.04970   0.04206   0.0063  0.4027  1.0008
   0.000   0.2212   0.04906   0.04118   0.0013  0.3708  1.0008
   0.250   0.2505   0.04847   0.04027   0.0028  0.3537  1.0008
   0.750   0.3021   0.04702   0.03829   0.0080  0.3298  1.0008
   1.000   0.3281   0.04638   0.03741   0.0105  0.3205  1.0008
   1.250   0.3539   0.04573   0.03653   0.0129  0.3105  1.0008
   1.500   0.3777   0.04467   0.03518   0.0163  0.3016  1.0008
   1.750   0.4048   0.04480   0.03509   0.0179  0.2933  1.0008
   2.000   0.4374   0.04684   0.03709   0.0163  0.2856  1.0008
   2.250   0.4708   0.04949   0.03975   0.0138  0.2772  1.0008
   2.500   0.5109   0.05536   0.04586   0.0058  0.2688  1.0008
   3.000   0.5838   0.08509   0.07647  -0.0377  0.2871  1.0008
   3.250   0.5973   0.09249   0.08395  -0.0453  0.2990  1.0008