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 =     0.025 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
  -2.000  -0.0762   0.04097   0.03124  -0.0470  1.0000  0.3277
  -1.000   0.0440   0.03519   0.02371  -0.0552  1.0000  0.4286
   0.000   0.1315   0.03210   0.02097  -0.0567  1.0000  1.0000
   1.000   0.2005   0.03721   0.02437  -0.0596  1.0000  1.0000
   2.000   0.3459   0.04515   0.03106  -0.0768  0.9473  1.0000
   3.000   0.4680   0.05337   0.03874  -0.0883  0.8848  1.0000
   4.000   0.5597   0.06220   0.04743  -0.0942  0.8276  1.0000
   5.000   0.6337   0.07239   0.05770  -0.0980  0.7790  1.0000
   6.000   0.6263   0.08428   0.06975  -0.0976  0.8112  1.0000
   7.000   0.5138   0.08798   0.07334  -0.0784  1.0000  1.0000
   8.000   0.6146   0.10530   0.09109  -0.0943  0.9388  1.0000
   9.000   0.6810   0.11938   0.10563  -0.1008  0.8602  1.0000
  10.000   0.7430   0.13413   0.12089  -0.1055  0.7928  1.0000