XFOIL         Version 6.94
  
 Calculated polar for: USA 48 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   0.3356   0.00979   0.00379  -0.0731  0.6865  0.7362
   0.500   0.3829   0.00988   0.00379  -0.0710  0.6642  0.7547
   1.000   0.4297   0.00995   0.00379  -0.0689  0.6423  0.7719
   1.500   0.4750   0.01002   0.00379  -0.0665  0.6198  0.7893
   2.000   0.5204   0.01005   0.00380  -0.0641  0.5990  0.8082
   2.500   0.5646   0.01010   0.00385  -0.0615  0.5785  0.8313
   3.000   0.6096   0.01017   0.00400  -0.0592  0.5585  0.8557
   4.000   0.7133   0.01076   0.00446  -0.0578  0.4668  0.9234
   4.500   0.7828   0.01198   0.00508  -0.0621  0.3493  0.9724
   5.000   0.8221   0.01310   0.00575  -0.0602  0.2751  1.0000
   5.500   0.8234   0.01463   0.00658  -0.0506  0.1596  1.0000
   6.000   0.8424   0.01582   0.00749  -0.0445  0.1202  1.0000
   6.500   0.8705   0.01672   0.00830  -0.0400  0.1031  1.0000
   7.000   0.9034   0.01747   0.00904  -0.0364  0.0900  1.0000
   7.500   0.9140   0.01937   0.01059  -0.0298  0.0220  1.0000
   8.000   0.9379   0.02074   0.01195  -0.0254  0.0054  1.0000
   8.500   0.9657   0.02201   0.01333  -0.0218  0.0050  1.0000
   9.000   0.9921   0.02347   0.01493  -0.0184  0.0050  1.0000
   9.500   1.0162   0.02518   0.01679  -0.0150  0.0049  1.0000
  10.000   1.0383   0.02717   0.01895  -0.0118  0.0050  1.0000
  10.500   1.0581   0.02946   0.02142  -0.0089  0.0052  1.0000
  11.000   1.0749   0.03218   0.02434  -0.0063  0.0053  1.0000
  11.500   1.0870   0.03549   0.02786  -0.0039  0.0055  1.0000
  12.000   1.0935   0.03958   0.03217  -0.0019  0.0057  1.0000
  12.500   1.0940   0.04454   0.03735  -0.0005  0.0059  1.0000
  13.000   1.0891   0.05043   0.04346   0.0003  0.0061  1.0000
  13.500   1.0808   0.05708   0.05030   0.0004  0.0063  1.0000
  14.000   1.0711   0.06418   0.05757   0.0000  0.0064  1.0000
  14.500   1.0653   0.07098   0.06451  -0.0005  0.0066  1.0000
  15.000   1.0712   0.07668   0.07044  -0.0012  0.0070  1.0000
  15.500   1.0722   0.08277   0.07672  -0.0016  0.0074  1.0000
  16.000   1.0780   0.08786   0.08195  -0.0010  0.0079  1.0000
  16.500   1.0966   0.09065   0.08478   0.0014  0.0086  1.0000
  17.000   1.1085   0.09520   0.08979   0.0039  0.0103  1.0000
  18.500   1.0426   0.12908   0.12514  -0.0052  0.0157  1.0000