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 =     0.060 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -2.000   0.0996   0.02617   0.01707  -0.0778   0.9465   0.1888
  -1.000   0.3013   0.02095   0.01254  -0.0922   0.8823   0.5210
   0.000   0.4502   0.01960   0.01102  -0.0948   0.8087   1.0000
   1.000   0.5566   0.02075   0.01133  -0.0916   0.7410   1.0000
   2.000   0.6582   0.02244   0.01248  -0.0883   0.6836   1.0000
   3.000   0.7586   0.02423   0.01396  -0.0848   0.6307   1.0000
   4.000   0.8551   0.02644   0.01616  -0.0813   0.5768   1.0000
   5.000   0.9507   0.02897   0.01885  -0.0779   0.5245   1.0000
   6.000   1.0435   0.03220   0.02244  -0.0745   0.4737   1.0000
   7.000   1.1277   0.03712   0.02805  -0.0713   0.4296   1.0000
   8.000   1.2296   0.03548   0.02659  -0.0656   0.3671   1.0000
   9.000   1.3007   0.03253   0.02429  -0.0568   0.2849   1.0000
  11.000   1.2461   0.05418   0.04547  -0.0419   0.0707   1.0000
  12.000   1.2038   0.07324   0.06494  -0.0453   0.0631   1.0000
  13.000   1.1964   0.08696   0.07901  -0.0444   0.0552   1.0000
  14.000   1.2630   0.09249   0.08520  -0.0320   0.0520   1.0000
  15.000   1.2174   0.11618   0.10981  -0.0409   0.0527   1.0000