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 =     0.055 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
   1.000   0.0447   0.02126   0.01113   0.0027   1.0000   1.0000
   2.000   0.1795   0.02375   0.01364  -0.0063   0.9800   1.0000
   3.000   0.4603   0.02509   0.01590  -0.0314   0.8301   1.0000
   4.000   0.5507   0.01932   0.00928  -0.0094   0.5238   1.0000
   5.000   0.6477   0.02304   0.01205  -0.0052   0.4122   1.0000
   6.000   0.7529   0.02761   0.01678  -0.0045   0.3520   1.0000
   7.000   0.8540   0.03375   0.02342  -0.0042   0.2993   1.0000
   8.000   0.9401   0.04212   0.03288  -0.0037   0.2452   1.0000
   9.000   1.0059   0.05209   0.04373  -0.0014   0.1929   1.0000
  10.000   1.0396   0.06356   0.05609   0.0021   0.1508   1.0000
  14.000   0.7829   0.17915   0.17210  -0.0555   0.1997   1.0000
  15.000   0.8182   0.19756   0.19059  -0.0599   0.1789   1.0000
  11.000   0.7472   0.13231   0.12507  -0.0445   0.3140   1.0000
  12.000   0.7511   0.14752   0.14034  -0.0467   0.2643   1.0000
  13.000   0.7962   0.16663   0.15963  -0.0470   0.2246   1.0000
   0.000  -0.0390   0.02013   0.01032   0.0015   1.0000   1.0000