XFOIL         Version 6.94
  
 Calculated polar for: GOE 495 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   0.6162   0.00857   0.00313  -0.1465  0.8528  0.5681
   0.500   0.6688   0.00777   0.00308  -0.1451  0.8363  1.0000
   1.000   0.7228   0.00786   0.00300  -0.1442  0.8157  1.0000
   2.000   0.8294   0.00811   0.00301  -0.1421  0.7708  1.0000
   2.500   0.8723   0.00813   0.00268  -0.1381  0.6814  1.0000
   3.000   0.9217   0.00851   0.00286  -0.1365  0.6292  1.0000
   3.500   0.9602   0.00961   0.00320  -0.1328  0.4968  1.0000
   4.000   0.9841   0.01241   0.00442  -0.1276  0.2036  1.0000
   4.500   1.0185   0.01468   0.00575  -0.1244  0.0409  1.0000
   5.000   1.0667   0.01540   0.00652  -0.1230  0.0325  1.0000
   5.500   1.1121   0.01639   0.00740  -0.1211  0.0105  1.0000
   6.000   1.1582   0.01728   0.00839  -0.1192  0.0077  1.0000
   6.500   1.2020   0.01839   0.00984  -0.1168  0.0075  1.0000
   7.000   1.2425   0.01979   0.01160  -0.1138  0.0076  1.0000
   7.500   1.2762   0.02181   0.01398  -0.1097  0.0077  1.0000
   8.000   1.2913   0.02562   0.01820  -0.1031  0.0056  1.0000
   8.500   1.3140   0.02973   0.02266  -0.0977  0.0055  1.0000
   9.000   1.3440   0.03395   0.02730  -0.0937  0.0059  1.0000
   9.500   1.3743   0.03893   0.03286  -0.0896  0.0071  1.0000
  10.000   1.3736   0.05085   0.04590  -0.0822  0.0096  1.0000