  
       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 =     0.420 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
   0.000   1.1174   0.01019   0.00395  -0.2706   0.7758   0.2461
   0.000   1.1174   0.01019   0.00395  -0.2706   0.7758   0.2461
   0.500   1.1721   0.01020   0.00407  -0.2707   0.7668   0.3105
   1.000   1.2266   0.01022   0.00424  -0.2709   0.7576   0.3882
   1.500   1.2809   0.01020   0.00447  -0.2710   0.7483   0.5060
   2.000   1.3291   0.00972   0.00467  -0.2697   0.7388   1.0000
   2.500   1.3821   0.00996   0.00485  -0.2694   0.7290   1.0000
   3.000   1.4344   0.01022   0.00508  -0.2690   0.7189   1.0000
   3.500   1.4861   0.01050   0.00535  -0.2685   0.7087   1.0000
   4.000   1.5371   0.01080   0.00567  -0.2679   0.6981   1.0000
   4.500   1.5876   0.01113   0.00602  -0.2672   0.6873   1.0000
   5.000   1.6363   0.01149   0.00645  -0.2661   0.6758   1.0000
