  
       XFOIL         Version 6.99
  
 Calculated polar for: NACA 9509                                       
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.175 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
   0.000   0.9860   0.02301   0.01549  -0.2524   0.8056   0.2456
   0.000   0.9860   0.02301   0.01549  -0.2524   0.8056   0.2456
   1.000   1.0963   0.02283   0.01570  -0.2528   0.7894   0.4140
   1.500   1.1701   0.02140   0.01489  -0.2560   0.7860   0.6656
   2.000   1.2172   0.02129   0.01489  -0.2542   0.7770   1.0000
   2.500   1.2702   0.02153   0.01504  -0.2538   0.7682   1.0000
   4.000   1.4526   0.02117   0.01464  -0.2568   0.7459   1.0000
   4.500   1.5054   0.02141   0.01496  -0.2564   0.7362   1.0000
