  
       XFOIL         Version 6.99
  
 Calculated polar for: NACA-like airfoil m=11%, p=40%, t=10%           
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.504 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
   0.000   1.2060   0.01175   0.00462  -0.2611   0.5932   0.0763
  10.000   2.0295   0.02325   0.01772  -0.2216   0.4280   0.9996
  11.000   2.0637   0.02716   0.02192  -0.2121   0.3986   0.9996
  12.000   2.0501   0.03573   0.03039  -0.1985   0.3204   0.9996
  13.000   1.9405   0.05552   0.04953  -0.1799   0.1872   0.9996
  14.000   1.8213   0.08064   0.07428  -0.1681   0.0838   0.9996
  15.000   1.7469   0.10271   0.09630  -0.1633   0.0356   0.9996
  16.000   1.7227   0.11871   0.11257  -0.1619   0.0271   0.9996
  17.000   1.7118   0.13300   0.12714  -0.1624   0.0231   0.9996
  18.000   1.7076   0.14624   0.14064  -0.1643   0.0201   0.9996
  19.000   1.7107   0.15825   0.15287  -0.1672   0.0174   0.9996
  20.000   1.7204   0.16884   0.16366  -0.1705   0.0150   0.9996
   0.000   1.2060   0.01175   0.00462  -0.2611   0.5932   0.0763
  10.000   2.0295   0.02325   0.01772  -0.2216   0.4280   0.9996
  11.000   2.0637   0.02716   0.02192  -0.2121   0.3986   0.9996
  12.000   2.0501   0.03573   0.03039  -0.1985   0.3204   0.9996
  13.000   1.9405   0.05552   0.04953  -0.1799   0.1872   0.9996
  14.000   1.8213   0.08064   0.07428  -0.1681   0.0838   0.9996
  15.000   1.7469   0.10271   0.09630  -0.1633   0.0356   0.9996
  16.000   1.7227   0.11871   0.11257  -0.1619   0.0271   0.9996
  17.000   1.7118   0.13300   0.12714  -0.1624   0.0231   0.9996
  18.000   1.7076   0.14624   0.14064  -0.1643   0.0201   0.9996
  19.000   1.7107   0.15825   0.15287  -0.1672   0.0174   0.9996
  20.000   1.7204   0.16884   0.16366  -0.1705   0.0150   0.9996
