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 =     0.035 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
   1.000   0.0451   0.02451   0.01169   0.0036   1.0000   1.0000
   2.000   0.1273   0.02622   0.01335   0.0043   1.0000   1.0000
   3.000   0.2043   0.02919   0.01669   0.0034   1.0000   1.0000
   4.000   0.5584   0.02928   0.01837  -0.0248   0.6530   1.0000
   5.000   0.6656   0.02878   0.01724  -0.0142   0.5212   1.0000
   6.000   0.7651   0.03473   0.02337  -0.0120   0.4422   1.0000
   7.000   0.8510   0.04417   0.03366  -0.0118   0.3818   1.0000
   8.000   0.9091   0.05740   0.04780  -0.0111   0.3308   1.0000
  10.000   0.6324   0.12445   0.11465  -0.0579   0.5950   1.0000
  11.000   0.6767   0.13921   0.12974  -0.0578   0.5075   1.0000
  12.000   0.7237   0.15570   0.14653  -0.0574   0.4210   1.0000
  13.000   0.7419   0.16901   0.15996  -0.0581   0.3532   1.0000
  14.000   0.7837   0.18743   0.17854  -0.0602   0.3004   1.0000
  15.000   0.7955   0.19915   0.19029  -0.0656   0.2655   1.0000
   0.000  -0.0326   0.02349   0.01105   0.0012   1.0000   1.0000
   9.000   0.6852   0.10486   0.09510  -0.0464   0.4736   1.0000