XFOIL         Version 6.94
  
 Calculated polar for: BOEING 707 .99 SPAN AIRFOIL                     
  
 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
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.2319   0.00928   0.00225  -0.0454  0.7822  0.0318
   0.500   0.2823   0.00902   0.00187  -0.0439  0.7567  0.0401
   1.000   0.3162   0.00635   0.00191  -0.0391  0.7327  0.9178
   1.500   0.4216   0.00655   0.00200  -0.0494  0.6949  0.9773
   2.000   0.5078   0.00684   0.00199  -0.0560  0.6290  0.9947
   2.500   0.5633   0.00735   0.00197  -0.0565  0.4989  1.0000
   3.000   0.5867   0.00877   0.00240  -0.0507  0.2817  1.0000
   3.500   0.6232   0.00957   0.00279  -0.0473  0.1943  1.0000
   4.500   0.6957   0.01152   0.00401  -0.0404  0.0443  1.0000
   5.000   0.7373   0.01218   0.00471  -0.0378  0.0397  1.0000
   5.500   0.7780   0.01296   0.00556  -0.0351  0.0356  1.0000
   6.000   0.8168   0.01390   0.00652  -0.0322  0.0290  1.0000
   6.500   0.8638   0.01417   0.00689  -0.0306  0.0217  1.0000
   7.000   0.9086   0.01464   0.00740  -0.0285  0.0134  1.0000
   7.500   0.9454   0.01581   0.00866  -0.0251  0.0120  1.0000
   8.000   0.9771   0.01735   0.01039  -0.0209  0.0113  1.0000
   8.500   1.0023   0.01938   0.01257  -0.0159  0.0104  1.0000
   9.000   1.0253   0.02197   0.01541  -0.0107  0.0102  1.0000
   9.500   1.0524   0.02542   0.01921  -0.0065  0.0101  1.0000
  10.000   1.0775   0.03120   0.02555  -0.0027  0.0102  1.0000
  11.500   0.9685   0.04691   0.04324   0.0205  0.0128  1.0000