XFOIL         Version 6.96
  
 Calculated polar for: LOCKHEED C-141 BL958.89 AIRFOIL
                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.045 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -3.000  -0.3279   0.04136   0.03179  -0.0261   1.0000   0.2160
  -2.000  -0.2163   0.03393   0.02278  -0.0269   1.0000   0.1925
  -1.000  -0.1110   0.03013   0.01802  -0.0252   1.0000   0.2028
   0.000  -0.0072   0.02768   0.01567  -0.0234   1.0000   0.2547
   1.000   0.0816   0.02524   0.01515  -0.0185   1.0000   1.0000
   2.000   0.1631   0.02764   0.01695  -0.0169   1.0000   1.0000
   3.000   0.2393   0.03089   0.02009  -0.0162   1.0000   1.0000
   4.000   0.4637   0.03564   0.02533  -0.0410   0.9017   1.0000
   5.000   0.7710   0.02356   0.01366  -0.0492   0.5571   1.0000
   6.000   0.8880   0.03038   0.01811  -0.0468   0.3170   1.0000
   7.000   0.9897   0.03817   0.02635  -0.0446   0.2557   1.0000
   8.000   1.0593   0.04791   0.03705  -0.0387   0.2220   1.0000
   9.000   1.0856   0.06031   0.05065  -0.0293   0.1967   1.0000
  12.000   0.7585   0.15265   0.14411  -0.0599   0.3474   1.0000


  a = 10.000      CL =  1.0374
      Cm = -0.0189     CD =  0.07860   =>   CDf =  0.00863    CDp =  0.06997