XFOIL         Version 6.94
  
 Calculated polar for: USA 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   0.3840   0.01102   0.00388  -0.0918  0.7349  0.4225
   0.500   0.4337   0.01091   0.00379  -0.0904  0.7196  0.4343
   1.000   0.4824   0.01088   0.00375  -0.0887  0.7041  0.4463
   1.500   0.5323   0.01087   0.00374  -0.0873  0.6876  0.4606
   3.500   0.6625   0.01094   0.00356  -0.0661  0.5139  0.5263
   5.000   0.9135   0.01353   0.00600  -0.0868  0.2712  1.0000
   5.500   0.8863   0.01636   0.00762  -0.0715  0.0598  1.0000
   6.000   0.9011   0.01769   0.00872  -0.0638  0.0072  1.0000
   6.500   0.9298   0.01848   0.00959  -0.0588  0.0067  1.0000
   7.000   0.9569   0.01940   0.01062  -0.0537  0.0066  1.0000
   7.500   0.9818   0.02048   0.01183  -0.0484  0.0066  1.0000
   8.000   1.0044   0.02178   0.01328  -0.0431  0.0067  1.0000
   8.500   1.0246   0.02332   0.01498  -0.0378  0.0069  1.0000
   9.000   1.0397   0.02531   0.01715  -0.0324  0.0072  1.0000
   9.500   1.0493   0.02786   0.01988  -0.0270  0.0075  1.0000
  10.000   1.0541   0.03100   0.02318  -0.0219  0.0077  1.0000
  10.500   1.0556   0.03472   0.02704  -0.0173  0.0080  1.0000
  11.000   1.0548   0.03897   0.03138  -0.0131  0.0082  1.0000
  11.500   1.0717   0.04199   0.03462  -0.0102  0.0090  1.0000
  12.000   1.0884   0.04543   0.03820  -0.0065  0.0100  1.0000
  12.500   1.1820   0.04774   0.04054  -0.0045  0.0130  1.0000
  13.000   1.0755   0.04523   0.03875   0.0069  0.0131  1.0000