XFOIL         Version 6.96
  
 Calculated polar for: NASA/AMES A-03 AIRFOIL
                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.040 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
   0.000  -0.0347   0.02241   0.01081   0.0013   1.0000   1.0000
   1.000   0.0448   0.02346   0.01149   0.0034   1.0000   1.0000
   2.000   0.1283   0.02520   0.01320   0.0039   1.0000   1.0000
   3.000   0.3230   0.03004   0.01870  -0.0175   0.9374   1.0000
   4.000   0.5624   0.02485   0.01449  -0.0192   0.6218   1.0000
   5.000   0.6601   0.02638   0.01491  -0.0100   0.4857   1.0000
   6.000   0.7621   0.03208   0.02082  -0.0086   0.4115   1.0000
   7.000   0.8550   0.04035   0.02992  -0.0084   0.3526   1.0000
   8.000   0.9267   0.05185   0.04235  -0.0074   0.3002   1.0000
  11.000   0.7153   0.13951   0.13081  -0.0545   0.4368   1.0000
  12.000   0.7340   0.15312   0.14461  -0.0536   0.3640   1.0000
  13.000   0.7495   0.16748   0.15905  -0.0554   0.3087   1.0000
  14.000   0.7755   0.18329   0.17495  -0.0589   0.2671   1.0000
  15.000   0.7980   0.19769   0.18942  -0.0642   0.2344   1.0000
   9.000   0.8867   0.07541   0.06711  -0.0125   0.2815   1.0000
  10.000   0.6609   0.12328   0.11426  -0.0554   0.5235   1.0000