XFOIL         Version 6.96
  
 Calculated polar for: NASA/AMES 63A108 MOD C AIRFOIL
                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.030 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -2.000  -0.1483   0.02053   0.00863  -0.0101   1.0000   1.0000
  -1.000  -0.0726   0.01958   0.00692  -0.0068   1.0000   1.0000
   0.000   0.0058   0.01944   0.00636  -0.0028   1.0000   1.0000
   1.000   0.0843   0.01991   0.00674   0.0011   1.0000   1.0000
   2.000   0.1641   0.02104   0.00812   0.0039   1.0000   1.0000
   3.000   0.2413   0.02330   0.01106   0.0045   1.0000   1.0000
   4.000   0.4846   0.02619   0.01447  -0.0047   0.5231   1.0000
   5.000   0.5811   0.03468   0.02225   0.0018   0.3193   1.0000
   6.000   0.6801   0.04756   0.03635   0.0007   0.2547   1.0000
   7.000   0.7347   0.06900   0.05934  -0.0097   0.2528   1.0000
   9.000   0.6318   0.12033   0.11060  -0.0659   0.5794   1.0000
  10.000   0.6656   0.13513   0.12537  -0.0634   0.4878   1.0000