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 =     0.023 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -1.000  -0.0730   0.02250   0.00772  -0.0067   1.0000   1.0000
   0.000   0.0064   0.02235   0.00710  -0.0029   1.0000   1.0000
   1.000   0.0851   0.02277   0.00746   0.0009   1.0000   1.0000
   2.000   0.1635   0.02385   0.00884   0.0040   1.0000   1.0000
   3.000   0.2394   0.02601   0.01173   0.0051   1.0000   1.0000
   5.000   0.5973   0.04063   0.02788  -0.0061   0.4049   1.0000
   6.000   0.6851   0.05554   0.04377  -0.0086   0.3315   1.0000
   7.000   0.7171   0.07829   0.06741  -0.0233   0.3353   1.0000
   9.000   0.5464   0.11520   0.10357  -0.0676   0.8127   1.0000
  10.000   0.6077   0.13338   0.12190  -0.0712   0.7022   1.0000