XFOIL         Version 6.94
  
 Calculated polar for: GOE 342 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.6960   0.01180   0.00566  -0.1493  0.7675  0.0291
   2.000   0.9199   0.00966   0.00313  -0.1490  0.7031  0.0447
   2.500   0.9750   0.00929   0.00282  -0.1488  0.6734  0.0482
   3.000   1.0280   0.00923   0.00264  -0.1481  0.6097  0.0516
   3.500   1.0651   0.01105   0.00322  -0.1449  0.3991  0.0617
   4.000   1.1070   0.01091   0.00409  -0.1430  0.3066  1.0000
   4.500   1.1542   0.01190   0.00459  -0.1418  0.2287  1.0000
   5.000   1.1927   0.01395   0.00582  -0.1393  0.0893  1.0000
   5.500   1.2327   0.01580   0.00714  -0.1369  0.0050  1.0000
   6.000   1.2810   0.01655   0.00793  -0.1356  0.0046  1.0000
   6.500   1.3279   0.01740   0.00890  -0.1340  0.0050  1.0000
   7.000   1.3738   0.01829   0.00993  -0.1322  0.0059  1.0000
   7.500   1.4170   0.01940   0.01122  -0.1299  0.0068  1.0000
   8.000   1.4580   0.02066   0.01267  -0.1271  0.0081  1.0000
   8.500   1.4918   0.02246   0.01470  -0.1232  0.0086  1.0000
   9.000   1.5173   0.02458   0.01702  -0.1180  0.0093  1.0000
   9.500   1.5250   0.02754   0.02018  -0.1104  0.0100  1.0000
  10.000   1.5336   0.03083   0.02364  -0.1037  0.0107  1.0000
  10.500   1.5432   0.03521   0.02825  -0.0972  0.0125  1.0000
  11.000   1.4879   0.03320   0.02677  -0.0803  0.0153  1.0000