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 =     0.052 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -1.000  -0.0715   0.01510   0.00567  -0.0070   1.0000   1.0000
   0.000   0.0038   0.01498   0.00520  -0.0026   1.0000   1.0000
   1.000   0.0825   0.01554   0.00566   0.0014   1.0000   1.0000
   2.000   0.1664   0.01679   0.00710   0.0032   1.0000   1.0000
   3.000   0.3835   0.01814   0.00969  -0.0187   0.9119   1.0000
   4.000   0.4596   0.02043   0.00957   0.0023   0.3599   1.0000
   5.000   0.5645   0.02697   0.01541   0.0049   0.2108   1.0000
   6.000   0.6702   0.03567   0.02477   0.0058   0.1588   1.0000
  10.000   0.7082   0.12841   0.12108  -0.0456   0.2524   1.0000