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 =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   1.500   0.4721   0.00880   0.00289  -0.0661  0.5779  0.7421
   2.000   0.5172   0.00894   0.00296  -0.0637  0.5568  0.7552
   2.500   0.5626   0.00910   0.00306  -0.0615  0.5375  0.7681
   3.000   0.6063   0.00918   0.00322  -0.0589  0.5180  0.7855
   3.500   0.6259   0.00949   0.00326  -0.0514  0.4495  0.8005
   4.000   0.6537   0.00994   0.00349  -0.0458  0.3893  0.8178
   4.500   0.6647   0.01110   0.00412  -0.0374  0.2727  0.8404
   5.000   0.7057   0.01241   0.00509  -0.0358  0.1565  0.9014
   5.500   0.8208   0.01414   0.00649  -0.0505  0.0997  0.9996
   6.000   0.8403   0.01541   0.00741  -0.0445  0.0287  1.0000
   6.500   0.8688   0.01633   0.00824  -0.0400  0.0044  1.0000
   7.000   0.9024   0.01708   0.00904  -0.0366  0.0040  1.0000
   7.500   0.9356   0.01791   0.00994  -0.0332  0.0040  1.0000
   8.000   0.9676   0.01887   0.01098  -0.0298  0.0041  1.0000
   8.500   0.9976   0.01999   0.01219  -0.0263  0.0043  1.0000
   9.000   1.0254   0.02132   0.01365  -0.0228  0.0045  1.0000
   9.500   1.0499   0.02294   0.01541  -0.0192  0.0047  1.0000
  10.000   1.0706   0.02493   0.01755  -0.0156  0.0050  1.0000
  10.500   1.0932   0.02696   0.01971  -0.0126  0.0054  1.0000
  11.500   1.1148   0.03338   0.02649  -0.0065  0.0063  1.0000
  12.500   1.1278   0.04174   0.03522  -0.0027  0.0074  1.0000
  13.000   1.1215   0.04764   0.04132  -0.0015  0.0080  1.0000
  13.500   1.1085   0.05447   0.04831  -0.0006  0.0084  1.0000
  14.000   1.1175   0.05931   0.05337  -0.0002  0.0094  1.0000
  14.500   1.1168   0.06447   0.05861   0.0019  0.0105  1.0000