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 =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.1555   0.00696   0.00271  -0.0361  0.9636  1.0000
   0.500   0.2249   0.00676   0.00248  -0.0385  0.9330  1.0000
   1.000   0.2790   0.00664   0.00229  -0.0373  0.8908  1.0000
   1.500   0.3276   0.00666   0.00217  -0.0349  0.8361  1.0000
   2.000   0.3754   0.00687   0.00213  -0.0324  0.7637  1.0000
   2.500   0.4226   0.00729   0.00219  -0.0300  0.6711  1.0000
   3.000   0.4700   0.00787   0.00236  -0.0279  0.5704  1.0000
   3.500   0.5183   0.00854   0.00267  -0.0263  0.4695  1.0000
   4.000   0.5658   0.00944   0.00311  -0.0249  0.3415  1.0000
   4.500   0.6081   0.01148   0.00404  -0.0232  0.1125  1.0000
   5.000   0.6548   0.01307   0.00536  -0.0217  0.0673  1.0000
   5.500   0.7031   0.01434   0.00668  -0.0203  0.0573  1.0000
   6.000   0.7519   0.01552   0.00797  -0.0191  0.0492  1.0000
   6.500   0.7987   0.01720   0.00975  -0.0176  0.0405  1.0000
   7.000   0.8446   0.01938   0.01209  -0.0159  0.0325  1.0000
   7.500   0.8895   0.02235   0.01537  -0.0141  0.0269  1.0000
   8.000   0.9319   0.02573   0.01896  -0.0125  0.0225  1.0000
   8.500   0.9688   0.03043   0.02449  -0.0096  0.0206  1.0000
   9.000   0.9951   0.03622   0.03110  -0.0063  0.0187  1.0000
   9.500   0.9899   0.04672   0.04266  -0.0015  0.0189  1.0000
  10.000   0.9556   0.05781   0.05446   0.0024  0.0199  1.0000
  10.500   0.9113   0.06801   0.06501   0.0005  0.0205  1.0000
  11.000   0.8660   0.08569   0.08293  -0.0146  0.0211  1.0000