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 =     0.040 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
  -2.000  -0.0488   0.03527   0.02642  -0.0545  1.0000  0.2140
  -1.000   0.1312   0.03143   0.02116  -0.0721  0.9657  0.3460
   0.000   0.3341   0.02870   0.01883  -0.0895  0.8900  1.0000
   1.000   0.4908   0.03144   0.02046  -0.0985  0.8212  1.0000
   2.000   0.6036   0.03466   0.02321  -0.0993  0.7602  1.0000
   3.000   0.7061   0.03814   0.02651  -0.0974  0.7023  1.0000
   4.000   0.7972   0.04227   0.03070  -0.0938  0.6442  1.0000
   5.000   0.8833   0.04698   0.03566  -0.0895  0.5877  1.0000
   6.000   0.9025   0.05947   0.04851  -0.0875  0.5359  1.0000
   7.000   0.8777   0.07731   0.06650  -0.0884  0.5071  1.0000
   9.000   0.7768   0.11827   0.10777  -0.1018  0.6280  1.0000
  10.000   0.8112   0.13048   0.12039  -0.1015  0.5831  1.0000