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 =     0.030 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -1.000   0.0425   0.03498   0.02413  -0.0566   1.0000   0.3604
   1.000   0.3348   0.03768   0.02542  -0.0829   0.9216   1.0000
   2.000   0.4662   0.04365   0.03059  -0.0927   0.8551   1.0000
   3.000   0.5778   0.05003   0.03665  -0.0978   0.7936   1.0000
   4.000   0.6608   0.05743   0.04402  -0.0988   0.7339   1.0000
   5.000   0.7121   0.06678   0.05350  -0.0975   0.6811   1.0000
   6.000   0.7517   0.07833   0.06527  -0.0975   0.6445   1.0000
   9.000   0.7228   0.12008   0.10773  -0.1033   0.7559   1.0000
  10.000   0.7685   0.13291   0.12103  -0.1044   0.6988   1.0000
  11.000   0.8224   0.14785   0.13653  -0.1068   0.6461   1.0000
  12.000   0.8466   0.15882   0.14805  -0.1065   0.5957   1.0000
  13.000   0.8815   0.17179   0.16166  -0.1082   0.5512   1.0000
  14.000   0.9151   0.18515   0.17588  -0.1104   0.5071   1.0000
  15.000   0.9632   0.20226   0.19360  -0.1050   0.3112   1.0000