XFOIL         Version 6.96
  
 Calculated polar for: AH 79-100 A AIRFOIL
                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.010 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -2.000  -0.1154   0.04033   0.02379  -0.0326   1.0000   1.0000
  -1.000  -0.0473   0.03845   0.01958  -0.0422   1.0000   1.0000
   0.000   0.0603   0.03975   0.01842  -0.0503   1.0000   1.0000
   1.000   0.1530   0.04215   0.01928  -0.0540   1.0000   1.0000
   2.000   0.2358   0.04548   0.02172  -0.0562   1.0000   1.0000
   3.000   0.3104   0.04980   0.02565  -0.0582   1.0000   1.0000
   4.000   0.3778   0.05524   0.03107  -0.0603   1.0000   1.0000
   5.000   0.4383   0.06193   0.03808  -0.0628   1.0000   1.0000
   6.000   0.4918   0.07003   0.04677  -0.0661   1.0000   1.0000
   7.000   0.5388   0.07970   0.05724  -0.0702   1.0000   1.0000
   8.000   0.5795   0.09109   0.06949  -0.0753   1.0000   1.0000
   9.000   0.6145   0.10421   0.08361  -0.0813   1.0000   1.0000
  10.000   0.6458   0.11886   0.09920  -0.0882   1.0000   1.0000