XFOIL         Version 6.94
  
 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 =     0.045 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
  -2.000  -0.0462   0.03383   0.02525  -0.0554  1.0000  0.1957
  -1.000   0.1984   0.02765   0.01789  -0.0813  0.9356  0.4326
   0.000   0.3892   0.02609   0.01657  -0.0950  0.8602  1.0000
   1.000   0.5310   0.02771   0.01723  -0.0989  0.7949  1.0000
   2.000   0.6404   0.02992   0.01898  -0.0972  0.7361  1.0000
   3.000   0.7362   0.03286   0.02174  -0.0937  0.6787  1.0000
   4.000   0.8260   0.03632   0.02528  -0.0895  0.6214  1.0000
   5.000   0.9087   0.04077   0.03000  -0.0851  0.5649  1.0000
   6.000   0.9921   0.04580   0.03544  -0.0809  0.5150  1.0000
   7.000   1.0096   0.05957   0.04977  -0.0795  0.4721  1.0000
   8.000   0.8981   0.08890   0.07904  -0.0868  0.4629  1.0000
   8.100   0.8935   0.09121   0.08137  -0.0875  0.4647  1.0000
   8.200   0.8919   0.09328   0.08346  -0.0880  0.4663  1.0000
   8.300   0.8924   0.09519   0.08540  -0.0886  0.4676  1.0000
   8.400   0.8951   0.09699   0.08725  -0.0890  0.4687  1.0000
   8.500   0.9003   0.09870   0.08900  -0.0895  0.4698  1.0000
   8.000   0.8981   0.08890   0.07904  -0.0868  0.4629  1.0000
   8.100   0.8935   0.09121   0.08137  -0.0875  0.4647  1.0000
   8.200   0.8919   0.09328   0.08346  -0.0880  0.4663  1.0000
   8.300   0.8924   0.09519   0.08540  -0.0886  0.4676  1.0000
   8.400   0.8951   0.09699   0.08725  -0.0890  0.4687  1.0000
   8.500   0.9003   0.09870   0.08900  -0.0895  0.4698  1.0000