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 =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -2.000   0.0164   0.03045   0.02140  -0.0657   0.9807   0.1890
  -1.000   0.2440   0.02483   0.01569  -0.0869   0.9124   0.4849
   0.000   0.4244   0.02350   0.01433  -0.0969   0.8395   1.0000
   1.000   0.5449   0.02494   0.01490  -0.0967   0.7729   1.0000
   2.000   0.6504   0.02684   0.01633  -0.0939   0.7152   1.0000
   3.000   0.7522   0.02882   0.01806  -0.0900   0.6613   1.0000
   4.000   0.8438   0.03174   0.02102  -0.0859   0.6047   1.0000
   5.000   0.9315   0.03534   0.02487  -0.0818   0.5491   1.0000
   6.000   1.0238   0.03897   0.02885  -0.0777   0.4997   1.0000
   7.000   1.0911   0.04639   0.03700  -0.0744   0.4558   1.0000
   8.000   1.0333   0.06995   0.06113  -0.0759   0.4239   1.0000
  11.000   1.0867   0.09954   0.09239  -0.0665   0.2963   1.0000
  12.000   1.0107   0.12962   0.12245  -0.0780   0.2902   1.0000
  14.000   1.2160   0.10001   0.09255  -0.0386   0.0626   1.0000
  15.000   1.1689   0.12643   0.11992  -0.0495   0.0638   1.0000