XFOIL         Version 6.94
  
 Calculated polar for: NACA 1408                                       
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.500   0.2222   0.00543   0.00177  -0.0363  0.8859  1.0000
   1.000   0.2715   0.00548   0.00164  -0.0344  0.8323  1.0000
   1.500   0.3209   0.00569   0.00159  -0.0326  0.7654  1.0000
   2.000   0.3689   0.00611   0.00160  -0.0306  0.6671  1.0000
   2.500   0.4176   0.00661   0.00172  -0.0290  0.5735  1.0000
   3.000   0.4667   0.00717   0.00193  -0.0276  0.4813  1.0000
   3.500   0.5161   0.00784   0.00224  -0.0264  0.3800  1.0000
   4.000   0.5637   0.00894   0.00273  -0.0252  0.2313  1.0000
   4.500   0.6090   0.01067   0.00365  -0.0238  0.0700  1.0000
   5.000   0.6595   0.01159   0.00455  -0.0228  0.0541  1.0000
   5.500   0.7089   0.01267   0.00568  -0.0217  0.0464  1.0000
   6.000   0.7594   0.01352   0.00662  -0.0207  0.0402  1.0000
   6.500   0.8076   0.01478   0.00800  -0.0195  0.0330  1.0000
   7.000   0.8547   0.01630   0.00964  -0.0181  0.0260  1.0000
   7.500   0.9005   0.01810   0.01158  -0.0166  0.0210  1.0000
   8.000   0.9477   0.01956   0.01323  -0.0152  0.0175  1.0000
   8.500   0.9882   0.02266   0.01666  -0.0131  0.0151  1.0000
   9.000   1.0281   0.02572   0.02018  -0.0109  0.0135  1.0000
   9.500   1.0630   0.02905   0.02393  -0.0085  0.0124  1.0000
  10.000   1.0774   0.03537   0.03088  -0.0049  0.0115  1.0000
  10.500   1.0672   0.04342   0.03977  -0.0001  0.0111  1.0000
  11.000   1.0307   0.05158   0.04848   0.0042  0.0111  1.0000
  11.500   0.9861   0.06358   0.06088  -0.0014  0.0113  1.0000