XFOIL         Version 6.94
  
 Calculated polar for: GOE 395 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   0.7138   0.01331   0.00627  -0.1561  0.7680  0.1003
   0.500   0.7692   0.01324   0.00628  -0.1561  0.7615  0.1159
   1.000   0.8245   0.01305   0.00618  -0.1558  0.7531  0.1307
   1.500   0.8800   0.01274   0.00587  -0.1554  0.7453  0.1404
   2.000   0.9347   0.01258   0.00581  -0.1550  0.7354  0.1524
   2.500   0.9897   0.01222   0.00556  -0.1544  0.7233  0.1725
   3.000   1.0438   0.01146   0.00569  -0.1540  0.7095  0.5502
   3.500   1.0895   0.00938   0.00389  -0.1497  0.6153  1.0000
   4.000   1.1289   0.01071   0.00412  -0.1465  0.4555  1.0000
   4.500   1.1736   0.01195   0.00492  -0.1448  0.3717  1.0000
   5.000   1.2105   0.01399   0.00610  -0.1422  0.2296  1.0000
   5.500   1.2380   0.01699   0.00803  -0.1385  0.0594  1.0000
   6.000   1.2744   0.01886   0.00963  -0.1356  0.0066  1.0000
   6.500   1.3172   0.01995   0.01084  -0.1334  0.0067  1.0000
   7.000   1.3562   0.02127   0.01236  -0.1306  0.0071  1.0000
   7.500   1.3928   0.02266   0.01395  -0.1275  0.0080  1.0000
   8.000   1.4185   0.02458   0.01611  -0.1227  0.0088  1.0000
   8.500   1.4417   0.02674   0.01849  -0.1181  0.0099  1.0000
   9.000   1.4507   0.03023   0.02216  -0.1125  0.0110  1.0000
   9.500   1.4625   0.03382   0.02598  -0.1076  0.0125  1.0000
  10.000   1.4729   0.03797   0.03027  -0.1023  0.0144  1.0000
  10.500   1.3656   0.03470   0.02771  -0.0837  0.0149  1.0000