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 =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
   2.000   0.1303   0.02367   0.01300   0.0030   1.0000   1.0000
   3.000   0.4474   0.02689   0.01722  -0.0321   0.8526   1.0000
   4.000   0.5559   0.02031   0.01029  -0.0117   0.5559   1.0000
   5.000   0.6510   0.02392   0.01277  -0.0063   0.4325   1.0000
   6.000   0.7555   0.02880   0.01784  -0.0053   0.3681   1.0000
   7.000   0.8550   0.03548   0.02513  -0.0050   0.3133   1.0000
   8.000   0.9372   0.04473   0.03543  -0.0043   0.2591   1.0000
   9.000   0.9946   0.05649   0.04813  -0.0021   0.2087   1.0000
  10.000   0.9368   0.07941   0.07225  -0.0037   0.1922   1.0000
  13.000   0.7609   0.16455   0.15707  -0.0516   0.2467   1.0000
  14.000   0.7792   0.18010   0.17269  -0.0565   0.2163   1.0000
  15.000   0.8157   0.19856   0.19123  -0.0608   0.1937   1.0000
  12.000   0.7734   0.15135   0.14391  -0.0465   0.2855   1.0000
  11.000   0.7434   0.13490   0.12726  -0.0471   0.3454   1.0000
   1.000   0.0446   0.02187   0.01122   0.0030   1.0000   1.0000
   0.000  -0.0379   0.02077   0.01045   0.0015   1.0000   1.0000