XFOIL         Version 6.94
  
 Calculated polar for: GOE 14K 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
 ------- -------- --------- --------- -------- ------- -------
   0.500   0.3351   0.02216   0.01725  -0.0824  0.8804  1.0000
   1.000   0.3954   0.02189   0.01695  -0.0840  0.8760  1.0000
   2.500   0.6560   0.01397   0.00673  -0.0953  0.3788  1.0000
   3.000   0.6395   0.01610   0.00772  -0.0812  0.1319  1.0000
   3.500   0.6573   0.01718   0.00835  -0.0740  0.0424  1.0000
   4.000   0.6854   0.01787   0.00896  -0.0687  0.0119  1.0000
   4.500   0.7177   0.01841   0.00960  -0.0640  0.0128  1.0000
   5.000   0.7498   0.01901   0.01035  -0.0593  0.0146  1.0000
   5.500   0.7822   0.01972   0.01122  -0.0544  0.0171  1.0000
   6.000   0.8111   0.02078   0.01245  -0.0490  0.0178  1.0000
   6.500   0.8343   0.02233   0.01411  -0.0426  0.0183  1.0000
   7.000   0.8587   0.02431   0.01611  -0.0366  0.0189  1.0000
   7.500   0.9060   0.02634   0.01822  -0.0344  0.0200  1.0000
   8.500   1.1165   0.04135   0.03486  -0.0505  0.0228  1.0000
   9.000   1.0954   0.04552   0.03968  -0.0360  0.0218  1.0000