XFOIL         Version 6.94
  
 Calculated polar for: NACA 16009                                      
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.0000   0.01332   0.00848   0.0000  1.0000  1.0000
   0.500   0.0653   0.01329   0.00846  -0.0037  0.9951  1.0000
   1.000   0.1436   0.01322   0.00846  -0.0099  0.9873  1.0000
   1.500   0.3189   0.01164   0.00723  -0.0344  0.9626  0.9955
   2.000   0.4043   0.01027   0.00610  -0.0396  0.9215  0.9952
   2.500   0.4516   0.01108   0.00500  -0.0356  0.4279  0.9970
   3.000   0.5003   0.01440   0.00649  -0.0356  0.0247  1.0000
   3.500   0.5412   0.01612   0.00837  -0.0324  0.0200  1.0000
   4.000   0.5868   0.01885   0.01148  -0.0296  0.0209  1.0000
   5.000   0.6823   0.03088   0.02492  -0.0224  0.0395  1.0000
   5.500   0.7077   0.03381   0.02841  -0.0169  0.0287  1.0000
   6.000   0.7085   0.04387   0.03906  -0.0091  0.0264  1.0000
   6.500   0.7290   0.04632   0.04222  -0.0010  0.0210  1.0000
   7.000   0.7306   0.05168   0.04790   0.0064  0.0193  1.0000
   7.500   0.7240   0.05698   0.05342   0.0140  0.0184  1.0000
   8.000   0.7035   0.06320   0.05985   0.0220  0.0176  1.0000