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 =     0.035 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -2.000  -0.0563   0.03717   0.02819  -0.0526   1.0000   0.2396
  -1.000   0.0406   0.03417   0.02394  -0.0575   1.0000   0.3087
   0.000   0.2585   0.03078   0.02062  -0.0796   0.9346   1.0000
   1.000   0.4205   0.03518   0.02359  -0.0929   0.8606   1.0000
   2.000   0.5475   0.03955   0.02738  -0.0985   0.7973   1.0000
   3.000   0.6546   0.04433   0.03196  -0.0998   0.7372   1.0000
   4.000   0.7443   0.04992   0.03761  -0.0980   0.6771   1.0000
   5.000   0.7955   0.05856   0.04643  -0.0950   0.6206   1.0000
   6.000   0.8212   0.07027   0.05833  -0.0930   0.5767   1.0000
   7.000   0.8307   0.08508   0.07339  -0.0938   0.5572   1.0000
   9.000   0.7522   0.11934   0.10802  -0.1028   0.6818   1.0000
  10.000   0.8025   0.13291   0.12207  -0.1041   0.6329   1.0000