XFOIL         Version 6.94
  
 Calculated polar for: BOEING BACXXX AIRFOIL                           
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.055 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
  -1.000  -0.0329   0.02176   0.01239  -0.0083  1.0000  1.0000
   0.000  -0.0469   0.02015   0.01048   0.0062  1.0000  1.0000
   1.000   0.0235   0.02108   0.01095   0.0085  1.0000  1.0000
   2.000   0.1070   0.02358   0.01334   0.0067  0.9956  1.0000
   3.000   0.3991   0.02510   0.01543  -0.0283  0.8675  1.0000
   5.000   0.7415   0.02268   0.01244  -0.0364  0.5009  1.0000
   6.000   0.8346   0.02679   0.01641  -0.0327  0.4014  1.0000
   7.000   0.9152   0.03146   0.02136  -0.0273  0.3206  1.0000
   8.000   0.9908   0.03689   0.02723  -0.0214  0.2521  1.0000
   9.000   1.0538   0.04411   0.03512  -0.0144  0.1964  1.0000
  10.000   1.0727   0.05622   0.04857  -0.0045  0.1670  1.0000