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 =     0.035 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -1.500  -0.2404   0.04668   0.03933   0.0060   1.0000   0.4186
   0.000  -0.0466   0.03366   0.02327  -0.0064   1.0000   0.2647
   0.500   0.0001   0.03300   0.02188  -0.0073   1.0000   0.2976
   1.000   0.0794   0.03226   0.02070  -0.0163   0.9930   0.3772
   1.500   0.3023   0.03446   0.02304  -0.0521   0.9236   1.0000
   2.000   0.3769   0.03654   0.02452  -0.0592   0.8746   1.0000
   2.500   0.4495   0.03848   0.02613  -0.0655   0.8271   1.0000
   3.000   0.5369   0.04164   0.02932  -0.0738   0.7935   1.0000
   3.500   0.5387   0.04376   0.03136  -0.0690   0.7586   1.0000
   4.000   0.6260   0.04458   0.03209  -0.0742   0.7174   1.0000
   4.500   0.6632   0.04732   0.03473  -0.0730   0.6784   1.0000
   5.000   0.6929   0.04919   0.03654  -0.0707   0.6397   1.0000
   5.500   0.7804   0.04925   0.03656  -0.0730   0.6133   1.0000
   6.000   0.7970   0.05500   0.04232  -0.0710   0.5941   1.0000
   6.500   0.7641   0.06147   0.04887  -0.0667   0.5800   1.0000
   7.000   0.7541   0.06663   0.05412  -0.0640   0.5579   1.0000
   7.500   0.8128   0.06839   0.05600  -0.0629   0.5224   1.0000
   8.000   0.9084   0.06003   0.04757  -0.0536   0.4426   1.0000
   8.500   0.8759   0.07144   0.05908  -0.0540   0.4313   1.0000
   9.000   0.8474   0.08250   0.07021  -0.0554   0.4289   1.0000
   9.500   0.8218   0.09258   0.08035  -0.0572   0.4316   1.0000
  10.000   0.7981   0.10222   0.09007  -0.0594   0.4413   1.0000