XFOIL         Version 6.94
  
 Calculated polar for: PROPFAN CRUISE MISSILE WING 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   0.2038   0.00701   0.00263  -0.0488  0.9164  0.9839
   0.500   0.2669   0.00696   0.00249  -0.0507  0.8951  1.0000
   1.000   0.3179   0.00700   0.00243  -0.0496  0.8737  1.0000
   1.500   0.3700   0.00710   0.00252  -0.0487  0.8495  1.0000
   2.000   0.4216   0.00716   0.00254  -0.0475  0.8194  1.0000
   2.500   0.4674   0.00721   0.00225  -0.0441  0.7326  1.0000
   3.000   0.5167   0.00764   0.00236  -0.0423  0.6320  1.0000
   3.500   0.5540   0.00997   0.00300  -0.0395  0.2478  1.0000
   4.000   0.5960   0.01269   0.00478  -0.0377  0.0390  1.0000
   4.500   0.6444   0.01419   0.00643  -0.0364  0.0357  1.0000
   5.000   0.6886   0.01695   0.00929  -0.0345  0.0323  1.0000
   5.500   0.7388   0.01944   0.01195  -0.0333  0.0317  1.0000
   6.000   0.7899   0.02262   0.01547  -0.0321  0.0310  1.0000
   6.500   0.8382   0.02849   0.02189  -0.0304  0.0327  1.0000
   7.000   0.8848   0.02869   0.02253  -0.0291  0.0244  1.0000
   8.000   0.8571   0.03500   0.03132  -0.0116  0.0266  1.0000
   8.500   0.8440   0.04546   0.04229  -0.0069  0.0269  1.0000
   9.000   0.8103   0.05400   0.05115  -0.0020  0.0268  1.0000
   9.500   0.7585   0.06322   0.06058  -0.0011  0.0278  1.0000
  10.000   0.7102   0.07548   0.07292  -0.0048  0.0308  1.0000