XFOIL         Version 6.96
  
 Calculated polar for: USA 28 AIRFOIL
                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -5.000  -0.1200   0.00856   0.00263  -0.1030   0.8092   0.0682
  -4.500  -0.0756   0.00794   0.00225  -0.1007   0.7920   0.1391
  -4.000  -0.0269   0.00782   0.00208  -0.0990   0.7736   0.1552
  -3.500   0.0208   0.00774   0.00191  -0.0971   0.7525   0.1647
  -3.000   0.0676   0.00762   0.00173  -0.0950   0.7278   0.1759
  -2.500   0.1140   0.00757   0.00158  -0.0928   0.7022   0.1857
  -2.000   0.1608   0.00744   0.00145  -0.0907   0.6809   0.2102
  -1.500   0.2067   0.00717   0.00141  -0.0886   0.6646   0.2935
  -1.000   0.2566   0.00718   0.00145  -0.0872   0.6508   0.3291
  -0.500   0.3071   0.00724   0.00148  -0.0859   0.6382   0.3447
   0.000   0.3575   0.00732   0.00152  -0.0846   0.6275   0.3560
   0.500   0.4086   0.00737   0.00156  -0.0834   0.6167   0.3651
   1.000   0.4584   0.00742   0.00161  -0.0819   0.6049   0.3740
   1.500   0.5072   0.00749   0.00167  -0.0803   0.5919   0.3828
   2.000   0.5541   0.00755   0.00173  -0.0782   0.5748   0.3913
   2.500   0.5957   0.00763   0.00177  -0.0750   0.5475   0.3988
   3.000   0.6292   0.00794   0.00188  -0.0701   0.4832   0.4061
   3.500   0.6666   0.00833   0.00210  -0.0662   0.4466   0.4127
   4.500   0.7499   0.00886   0.00260  -0.0602   0.4139   0.4308
   5.500   0.8348   0.00928   0.00315  -0.0547   0.3934   0.4618
   6.000   0.8769   0.00935   0.00341  -0.0519   0.3809   0.5118