XFOIL         Version 6.94
  
 Calculated polar for: GOE 440 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
 ------- -------- --------- --------- -------- ------- -------
   1.000   1.2530   0.01613   0.00803  -0.2309  0.4387  0.5150
   1.500   1.2922   0.01590   0.00854  -0.2280  0.4296  1.0000
   2.000   1.2995   0.01710   0.00907  -0.2204  0.3346  1.0000
   2.500   1.2650   0.02240   0.01295  -0.2078  0.1446  1.0000
   3.000   1.2731   0.02562   0.01589  -0.2022  0.0216  1.0000
   3.500   1.3072   0.02707   0.01734  -0.2003  0.0291  1.0000
   4.000   1.3399   0.02866   0.01893  -0.1985  0.0353  1.0000
   5.000   1.5009   0.02476   0.01608  -0.2058  0.2287  1.0000
   5.500   1.4576   0.03202   0.02262  -0.1952  0.1149  1.0000