XFOIL         Version 6.94
  
 Calculated polar for: GOE 287 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   0.4305   0.01271   0.00432  -0.0729  0.6037  0.0419
   1.000   0.5453   0.01217   0.00383  -0.0738  0.5849  0.0625
   1.500   0.6030   0.01187   0.00359  -0.0742  0.5753  0.0801
   2.000   0.6539   0.00983   0.00359  -0.0734  0.5658  1.0000
   2.500   0.7112   0.00999   0.00364  -0.0738  0.5556  1.0000
   3.000   0.7677   0.01029   0.00378  -0.0741  0.5450  1.0000
   3.500   0.8244   0.01034   0.00383  -0.0744  0.5284  1.0000
   4.000   0.8779   0.01030   0.00342  -0.0740  0.4234  1.0000
   4.500   0.9249   0.01218   0.00430  -0.0738  0.2611  1.0000
   5.000   0.9759   0.01328   0.00507  -0.0738  0.2071  1.0000
   5.500   1.0152   0.01609   0.00691  -0.0728  0.0264  1.0000
   6.000   1.0651   0.01706   0.00790  -0.0722  0.0075  1.0000
   6.500   1.1143   0.01804   0.00906  -0.0714  0.0077  1.0000
   7.000   1.1607   0.01928   0.01054  -0.0701  0.0084  1.0000
   7.500   1.2011   0.02109   0.01267  -0.0681  0.0096  1.0000
   8.000   1.2292   0.02378   0.01566  -0.0647  0.0109  1.0000
   8.500   1.2313   0.02770   0.01978  -0.0587  0.0123  1.0000
   9.000   1.2356   0.03205   0.02425  -0.0533  0.0148  1.0000
   9.500   1.1548   0.02527   0.01799  -0.0406  0.0156  1.0000