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 =     0.045 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
   1.000   0.0446   0.02260   0.01134   0.0032   1.0000   1.0000
   2.000   0.1293   0.02436   0.01308   0.0035   1.0000   1.0000
   3.000   0.3957   0.02887   0.01848  -0.0272   0.8860   1.0000
   4.000   0.5606   0.02204   0.01196  -0.0150   0.5894   1.0000
   5.000   0.6551   0.02498   0.01367  -0.0077   0.4565   1.0000
   6.000   0.7585   0.03025   0.01913  -0.0065   0.3877   1.0000
   7.000   0.8567   0.03743   0.02688  -0.0060   0.3308   1.0000
   8.000   0.9297   0.04837   0.03913  -0.0058   0.2777   1.0000
   9.000   0.9556   0.06411   0.05604  -0.0048   0.2361   1.0000
  10.000   0.8668   0.09344   0.08577  -0.0141   0.2399   1.0000
  12.000   0.7360   0.15075   0.14274  -0.0511   0.3240   1.0000
  13.000   0.7798   0.16922   0.16139  -0.0516   0.2702   1.0000
  14.000   0.7868   0.18336   0.17557  -0.0566   0.2387   1.0000
  15.000   0.8269   0.20287   0.19519  -0.0604   0.2095   1.0000
   0.000  -0.0365   0.02152   0.01061   0.0014   1.0000   1.0000
  11.000   0.7225   0.13681   0.12866  -0.0511   0.3900   1.0000