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 =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -5.000  -0.1262   0.00973   0.00355  -0.1026   0.8674   0.1416
  -4.500  -0.0665   0.00953   0.00326  -0.1035   0.8502   0.1706
  -4.000  -0.0111   0.00935   0.00298  -0.1033   0.8319   0.1868
  -3.500   0.0391   0.00916   0.00273  -0.1020   0.8136   0.2017
  -3.000   0.0876   0.00894   0.00253  -0.1003   0.7953   0.2263
  -2.500   0.1345   0.00868   0.00241  -0.0983   0.7759   0.2779
  -2.000   0.1827   0.00868   0.00244  -0.0965   0.7560   0.3308
  -1.500   0.2320   0.00875   0.00244  -0.0948   0.7376   0.3527
  -1.000   0.2808   0.00882   0.00242  -0.0931   0.7197   0.3675
  -0.500   0.3299   0.00891   0.00241  -0.0914   0.7023   0.3797
   0.000   0.3793   0.00895   0.00242  -0.0898   0.6865   0.3906
   0.500   0.4291   0.00901   0.00244  -0.0884   0.6721   0.4008
   1.000   0.4793   0.00913   0.00249  -0.0870   0.6581   0.4104
   1.500   0.5280   0.00918   0.00255  -0.0854   0.6429   0.4207
   2.000   0.5760   0.00926   0.00263  -0.0836   0.6273   0.4304
   2.500   0.6221   0.00935   0.00273  -0.0814   0.6098   0.4403
   3.000   0.6647   0.00937   0.00282  -0.0784   0.5882   0.4499
   3.500   0.7034   0.00944   0.00292  -0.0746   0.5601   0.4639
   4.000   0.7356   0.00954   0.00300  -0.0694   0.5204   0.4782
   4.500   0.7635   0.00978   0.00318  -0.0634   0.4810   0.5015