  
       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 =     0.541 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
   0.500   1.1761   0.00909   0.00326  -0.2709   0.7526   0.2891
   1.000   1.2307   0.00913   0.00341  -0.2711   0.7435   0.3589
   1.500   1.2849   0.00917   0.00362  -0.2712   0.7342   0.4527
   2.000   1.3338   0.00861   0.00391  -0.2701   0.7246   1.0000
   2.500   1.3864   0.00885   0.00408  -0.2698   0.7146   1.0000
   3.000   1.4382   0.00911   0.00431  -0.2693   0.7044   1.0000
   3.500   1.4891   0.00938   0.00459  -0.2687   0.6938   1.0000
   4.000   1.5392   0.00968   0.00490  -0.2679   0.6829   1.0000
   4.500   1.5883   0.01000   0.00526  -0.2669   0.6715   1.0000
   5.000   1.6357   0.01035   0.00564  -0.2655   0.6586   1.0000
