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 =     0.021 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -2.000  -0.1490   0.02414   0.00971  -0.0101   1.0000   1.0000
  -1.000  -0.0731   0.02327   0.00793  -0.0066   1.0000   1.0000
   0.000   0.0064   0.02311   0.00730  -0.0029   1.0000   1.0000
   1.000   0.0853   0.02353   0.00764   0.0009   1.0000   1.0000
   2.000   0.1635   0.02460   0.00903   0.0040   1.0000   1.0000
   3.000   0.2390   0.02674   0.01192   0.0053   1.0000   1.0000
   4.000   0.2860   0.03296   0.01922  -0.0001   1.0000   1.0000
   5.000   0.6020   0.04244   0.02960  -0.0093   0.4289   1.0000
   6.000   0.6826   0.05839   0.04656  -0.0137   0.3567   1.0000
   7.000   0.7006   0.08127   0.06998  -0.0295   0.3647   1.0000
   9.000   0.5253   0.11342   0.10126  -0.0655   0.8764   1.0000
  10.000   0.5983   0.13373   0.12179  -0.0730   0.7638   1.0000