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 =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.2384   0.01048   0.00294  -0.0481  0.8218  0.0428
   0.500   0.2683   0.00829   0.00276  -0.0433  0.8031  0.6993
   1.000   0.3837   0.00794   0.00298  -0.0536  0.7718  0.9742
   1.500   0.4859   0.00795   0.00283  -0.0632  0.7360  0.9976
   2.000   0.5322   0.00784   0.00265  -0.0616  0.7028  1.0000
   2.500   0.5697   0.00787   0.00251  -0.0578  0.6562  1.0000
   3.000   0.6082   0.00810   0.00254  -0.0542  0.5940  1.0000
   3.500   0.6389   0.00888   0.00269  -0.0491  0.4436  1.0000
   4.000   0.6669   0.01027   0.00331  -0.0442  0.2829  1.0000
   4.500   0.7045   0.01117   0.00392  -0.0410  0.2075  1.0000
   5.000   0.7341   0.01280   0.00477  -0.0367  0.0681  1.0000
   5.500   0.7732   0.01372   0.00562  -0.0337  0.0489  1.0000
   6.000   0.8124   0.01466   0.00664  -0.0308  0.0423  1.0000
   6.500   0.8445   0.01620   0.00823  -0.0269  0.0326  1.0000
   7.000   0.8793   0.01757   0.00977  -0.0233  0.0221  1.0000
   7.500   0.9126   0.01911   0.01145  -0.0193  0.0176  1.0000
   8.000   0.9442   0.02096   0.01342  -0.0153  0.0158  1.0000
   8.500   0.9775   0.02409   0.01672  -0.0118  0.0152  1.0000
   9.000   1.0172   0.02805   0.02113  -0.0093  0.0140  1.0000
   9.500   1.0476   0.03434   0.02814  -0.0055  0.0143  1.0000
  10.000   1.0584   0.04077   0.03520   0.0000  0.0151  1.0000
  10.500   0.9809   0.04234   0.03790   0.0127  0.0179  1.0000