XFOIL         Version 6.96
  
 Calculated polar for: USA 22 AIRFOIL
                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.040 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -2.000  -0.0486   0.03526   0.02641  -0.0545   1.0000   0.2140
  -1.000   0.1313   0.03143   0.02116  -0.0721   0.9657   0.3461
   1.000   0.4909   0.03144   0.02046  -0.0985   0.8212   1.0000
   2.000   0.6036   0.03466   0.02321  -0.0993   0.7602   1.0000
   3.000   0.7061   0.03814   0.02651  -0.0974   0.7023   1.0000
   4.000   0.7972   0.04227   0.03070  -0.0938   0.6442   1.0000
   5.000   0.8833   0.04699   0.03566  -0.0895   0.5877   1.0000
   6.000   0.9024   0.05947   0.04851  -0.0875   0.5359   1.0000
   7.000   0.8777   0.07731   0.06650  -0.0884   0.5071   1.0000
  10.000   0.8111   0.13047   0.12038  -0.1015   0.5831   1.0000
  11.000   0.8434   0.14294   0.13329  -0.1014   0.5383   1.0000
  12.000   0.8844   0.15703   0.14797  -0.1026   0.4984   1.0000
  13.000   0.9022   0.16857   0.16004  -0.1038   0.4639   1.0000
  15.000   1.0283   0.17167   0.16447  -0.0800   0.1080   1.0000