XFOIL         Version 6.94
  
 Calculated polar for: KC-135 BL351.6 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   0.2177   0.00773   0.00288  -0.0390  0.8269  0.8414
   0.500   0.3228   0.00774   0.00293  -0.0477  0.8062  0.9743
   1.000   0.4275   0.00775   0.00280  -0.0582  0.7879  0.9977
   1.500   0.4770   0.00771   0.00266  -0.0573  0.7683  1.0000
   2.000   0.5214   0.00777   0.00263  -0.0551  0.7490  1.0000
   2.500   0.5666   0.00785   0.00272  -0.0530  0.7270  1.0000
   3.000   0.6128   0.00795   0.00281  -0.0510  0.7028  1.0000
   3.500   0.6590   0.00808   0.00295  -0.0489  0.6730  1.0000
   4.000   0.7042   0.00825   0.00313  -0.0465  0.6300  1.0000
   4.500   0.7416   0.00879   0.00326  -0.0426  0.5104  1.0000
   5.000   0.7682   0.01045   0.00399  -0.0374  0.3229  1.0000
   5.500   0.8068   0.01153   0.00480  -0.0346  0.2494  1.0000
   6.000   0.8433   0.01287   0.00566  -0.0317  0.1427  1.0000
   6.500   0.8701   0.01541   0.00760  -0.0270  0.0372  1.0000
   7.000   0.9033   0.01738   0.00975  -0.0231  0.0281  1.0000
   7.500   0.9341   0.01972   0.01220  -0.0192  0.0236  1.0000
   8.000   0.9673   0.02378   0.01655  -0.0157  0.0217  1.0000
   8.500   1.0087   0.02746   0.02061  -0.0135  0.0211  1.0000
   9.000   1.0440   0.03142   0.02506  -0.0106  0.0200  1.0000
   9.500   1.0673   0.03651   0.03077  -0.0065  0.0195  1.0000
  10.000   1.0701   0.04340   0.03836  -0.0007  0.0201  1.0000
  10.500   0.9541   0.03322   0.02901   0.0128  0.0208  1.0000
  11.000   0.9079   0.04217   0.03835   0.0169  0.0214  1.0000