XFOIL         Version 6.94
  
 Calculated polar for: NACA 6-H-20 AIRFOIL                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.500   0.3502   0.03078   0.02454  -0.0271  0.5663  0.2658
   1.000   0.2988   0.02666   0.02104  -0.0034  0.5682  0.5029