XFOIL         Version 6.96
  
 Calculated polar for: USA 5 AIRFOIL
                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.045 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -1.500  -0.2383   0.04650   0.03998   0.0052   1.0000   0.3673
   0.500   0.1295   0.03207   0.02134  -0.0299   0.9490   0.2945
   1.500   0.3960   0.03016   0.01971  -0.0640   0.8609   1.0000
   2.000   0.5068   0.03036   0.01946  -0.0749   0.8149   1.0000
   2.500   0.5617   0.03160   0.02074  -0.0765   0.7702   1.0000
   3.000   0.6549   0.03019   0.01915  -0.0815   0.7266   1.0000
   3.500   0.7398   0.02945   0.01818  -0.0849   0.6865   1.0000
   4.000   0.7820   0.02987   0.01848  -0.0824   0.6441   1.0000
   4.500   0.8626   0.02997   0.01833  -0.0853   0.6069   1.0000
   5.000   0.8761   0.03342   0.02187  -0.0809   0.5810   1.0000
   5.500   0.8933   0.03624   0.02478  -0.0767   0.5546   1.0000
   6.000   0.9338   0.03798   0.02658  -0.0748   0.5323   1.0000
   7.000   1.0438   0.02793   0.01476  -0.0593   0.3431   1.0000
   7.500   1.0594   0.02922   0.01600  -0.0512   0.2605   1.0000
   8.000   1.0474   0.03287   0.01844  -0.0411   0.0928   1.0000
   8.500   1.0369   0.03751   0.02282  -0.0328   0.0701   1.0000
   9.000   1.0295   0.04194   0.02767  -0.0258   0.0663   1.0000
   9.500   1.0291   0.04667   0.03289  -0.0199   0.0654   1.0000
  10.000   1.0979   0.05354   0.04083  -0.0154   0.0702   1.0000