XFOIL         Version 6.94
  
 Calculated polar for: KENNEDY AND MARSDEN AIRFOIL                     
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.8817   0.02492   0.01711  -0.1894  0.4380  0.1180
   0.500   0.4588   0.05100   0.04463  -0.1385  0.4292  0.0775
   1.000   0.5440   0.04947   0.04313  -0.1417  0.4293  0.0849
   1.500   0.5316   0.05773   0.05144  -0.1436  0.4230  0.0885
   2.000   0.5749   0.06139   0.05521  -0.1459  0.4239  0.0990
   2.500   0.7016   0.06171   0.05656  -0.1607  0.4241  0.3467