XFOIL         Version 6.94
  
 Calculated polar for: GOE 7K 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   0.4571   0.01980   0.01465  -0.1030  0.8309  1.0000
   1.500   0.5306   0.01906   0.01383  -0.1066  0.8280  1.0000
   2.000   0.6045   0.01827   0.01300  -0.1104  0.8261  1.0000
   2.500   0.6754   0.01767   0.01240  -0.1138  0.8248  1.0000
   3.000   0.8308   0.01024   0.00446  -0.1282  0.7242  1.0000
   3.500   0.8136   0.01140   0.00471  -0.1121  0.5495  1.0000
   4.000   0.7688   0.01412   0.00638  -0.0921  0.3527  1.0000
   4.500   0.7414   0.01760   0.00841  -0.0776  0.0586  1.0000
   5.000   0.7708   0.01873   0.00942  -0.0730  0.0126  1.0000