XFOIL         Version 6.94
  
 Calculated polar for: RAF 28 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   0.3088   0.00790   0.00312  -0.0656  0.8052  0.9779
   0.500   0.3936   0.00798   0.00308  -0.0718  0.7803  0.9937
   1.000   0.4558   0.00793   0.00292  -0.0735  0.7514  1.0000
   1.500   0.4963   0.00793   0.00278  -0.0704  0.7205  1.0000
   2.000   0.5282   0.00807   0.00251  -0.0649  0.6299  1.0000
   2.500   0.5629   0.00855   0.00253  -0.0604  0.5312  1.0000
   3.000   0.5944   0.00944   0.00272  -0.0556  0.3879  1.0000
   3.500   0.6252   0.01078   0.00327  -0.0512  0.2189  1.0000
   4.000   0.6525   0.01265   0.00423  -0.0463  0.0389  1.0000
   4.500   0.6946   0.01332   0.00487  -0.0436  0.0215  1.0000
   5.000   0.7362   0.01408   0.00564  -0.0407  0.0080  1.0000
   5.500   0.7781   0.01485   0.00655  -0.0377  0.0081  1.0000
   6.000   0.8171   0.01589   0.00779  -0.0342  0.0085  1.0000
   6.500   0.8537   0.01711   0.00923  -0.0303  0.0093  1.0000
   7.000   0.8793   0.01918   0.01150  -0.0246  0.0104  1.0000
   7.500   0.9041   0.02207   0.01456  -0.0187  0.0120  1.0000
   8.000   0.9628   0.02989   0.02290  -0.0171  0.0188  1.0000
   8.500   1.0010   0.04321   0.03797  -0.0088  0.0498  1.0000
   9.500   0.9840   0.05487   0.05074   0.0053  0.0415  1.0000