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 =     0.030 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -2.000  -0.3399   0.05144   0.04466   0.0444   1.0000   0.6536
   0.000  -0.0302   0.03438   0.02303  -0.0092   1.0000   0.3125
   0.500   0.0373   0.03249   0.02036  -0.0139   1.0000   0.3522
   1.000   0.0823   0.03271   0.02049  -0.0162   1.0000   0.4278
   1.500   0.1764   0.03467   0.02326  -0.0298   1.0000   1.0000
   2.000   0.3055   0.03871   0.02598  -0.0492   0.9284   1.0000
   2.500   0.3777   0.04170   0.02850  -0.0569   0.8854   1.0000
   3.000   0.4158   0.04402   0.03060  -0.0585   0.8471   1.0000
   3.500   0.5124   0.04794   0.03462  -0.0691   0.8094   1.0000
   4.000   0.5285   0.05148   0.03807  -0.0675   0.7804   1.0000
   4.500   0.5399   0.05471   0.04128  -0.0658   0.7522   1.0000
   5.000   0.6406   0.05744   0.04393  -0.0728   0.7096   1.0000
   5.500   0.6472   0.06169   0.04812  -0.0698   0.6845   1.0000
   6.000   0.6529   0.06399   0.05040  -0.0661   0.6521   1.0000
   6.500   0.6842   0.06768   0.05413  -0.0660   0.6315   1.0000
   7.000   0.7721   0.07408   0.06062  -0.0689   0.5986   1.0000
   7.500   0.7391   0.07810   0.06469  -0.0652   0.5814   1.0000
   8.000   0.7245   0.08393   0.07056  -0.0637   0.5612   1.0000
   8.500   0.7200   0.09085   0.07755  -0.0638   0.5529   1.0000
   9.000   0.7145   0.09958   0.08635  -0.0655   0.5683   1.0000