XFOIL         Version 6.94
  
 Calculated polar for: GOE 402 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   0.5557   0.01204   0.00465  -0.0990  0.6363  0.0135
   0.500   0.6090   0.01151   0.00372  -0.0978  0.6135  0.0253
   1.000   0.6621   0.01189   0.00398  -0.0971  0.5883  0.0384
   1.500   0.7138   0.01191   0.00382  -0.0963  0.5590  0.0544
   2.000   0.7647   0.01162   0.00341  -0.0951  0.5262  0.0529
   2.500   0.8154   0.01136   0.00303  -0.0940  0.4921  0.0581
   3.000   0.8707   0.00988   0.00314  -0.0943  0.4562  1.0000
   3.500   0.9209   0.01037   0.00335  -0.0933  0.4242  1.0000
   4.000   0.9711   0.01091   0.00369  -0.0923  0.4008  1.0000
   5.500   1.0838   0.01612   0.00656  -0.0849  0.0079  1.0000
   6.000   1.1311   0.01685   0.00728  -0.0835  0.0052  1.0000
   6.500   1.1777   0.01762   0.00827  -0.0819  0.0055  1.0000
   7.000   1.2239   0.01839   0.00922  -0.0802  0.0061  1.0000
   7.500   1.2663   0.01949   0.01062  -0.0779  0.0068  1.0000
   8.000   1.3038   0.02098   0.01249  -0.0748  0.0078  1.0000
   8.500   1.3345   0.02286   0.01474  -0.0707  0.0095  1.0000
   9.000   1.3480   0.02542   0.01765  -0.0643  0.0112  1.0000
   9.500   1.3402   0.02925   0.02177  -0.0565  0.0126  1.0000
  10.000   1.3305   0.03423   0.02697  -0.0510  0.0136  1.0000