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 =     0.025 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -2.000  -0.1487   0.02228   0.00915  -0.0101   1.0000   1.0000
  -1.000  -0.0729   0.02138   0.00741  -0.0067   1.0000   1.0000
   0.000   0.0062   0.02123   0.00682  -0.0029   1.0000   1.0000
   1.000   0.0848   0.02167   0.00718   0.0010   1.0000   1.0000
   2.000   0.1637   0.02277   0.00856   0.0040   1.0000   1.0000
   3.000   0.2401   0.02497   0.01147   0.0049   1.0000   1.0000
   4.000   0.5025   0.02961   0.01774  -0.0127   0.5827   1.0000
   5.000   0.5902   0.03804   0.02528  -0.0018   0.3707   1.0000
   6.000   0.6835   0.05248   0.04102  -0.0047   0.3007   1.0000
   7.000   0.7194   0.07542   0.06506  -0.0193   0.3048   1.0000
   9.000   0.5802   0.11794   0.10704  -0.0689   0.7226   1.0000
  10.000   0.6462   0.13720   0.12643  -0.0708   0.6144   1.0000