XFOIL         Version 6.94
  
 Calculated polar for: BOEING 707 .54 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   0.2252   0.00812   0.00249  -0.0411  0.8143  0.6736
   0.500   0.3217   0.00753   0.00251  -0.0466  0.7709  0.9785
   1.000   0.4168   0.00753   0.00215  -0.0550  0.7138  1.0000
   1.500   0.4585   0.00755   0.00198  -0.0524  0.6775  1.0000
   2.000   0.5006   0.00771   0.00193  -0.0498  0.6238  1.0000
   2.500   0.5444   0.00796   0.00202  -0.0475  0.5701  1.0000
   3.000   0.5843   0.00855   0.00223  -0.0445  0.4645  1.0000
   3.500   0.6066   0.01084   0.00305  -0.0392  0.1829  1.0000
   4.000   0.6414   0.01233   0.00385  -0.0359  0.0486  1.0000
   4.500   0.6866   0.01290   0.00443  -0.0341  0.0289  1.0000
   5.000   0.7305   0.01367   0.00518  -0.0321  0.0195  1.0000
   5.500   0.7754   0.01442   0.00609  -0.0302  0.0150  1.0000
   6.000   0.8201   0.01525   0.00720  -0.0283  0.0084  1.0000
   6.500   0.8602   0.01663   0.00878  -0.0256  0.0072  1.0000
   7.000   0.8931   0.01879   0.01120  -0.0218  0.0068  1.0000
   7.500   0.9224   0.02190   0.01454  -0.0174  0.0069  1.0000
   8.000   0.9582   0.02612   0.01913  -0.0142  0.0073  1.0000
   8.500   0.9907   0.03160   0.02526  -0.0109  0.0079  1.0000
   9.000   1.0068   0.03855   0.03285  -0.0064  0.0085  1.0000
   9.500   0.9574   0.04233   0.03798   0.0049  0.0123  1.0000