XFOIL         Version 6.94
  
 Calculated polar for: NACA M22 AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   1.000   0.4142   0.01766   0.00985  -0.0252  0.5474  0.0371
   2.500   0.5749   0.00698  -0.00125  -0.0254  0.5175  0.0670
   3.000   0.6298   0.00616  -0.00212  -0.0253  0.5086  0.0586
   4.500   0.7917   0.00505  -0.00350  -0.0251  0.3944  0.0605
   5.000   0.8760   0.00651  -0.00225  -0.0369  0.0121  1.0000
   5.500   0.9243   0.00701  -0.00164  -0.0363  0.0095  1.0000
   6.000   0.9715   0.00766  -0.00063  -0.0354  0.0099  1.0000
   6.500   1.0151   0.00873   0.00080  -0.0344  0.0108  1.0000
   7.000   1.0463   0.01090   0.00337  -0.0327  0.0121  1.0000
   7.500   1.0482   0.01463   0.00739  -0.0297  0.0129  1.0000
   8.000   1.0477   0.01772   0.01065  -0.0260  0.0144  1.0000
   8.500   1.0467   0.02210   0.01522  -0.0221  0.0171  1.0000