XFOIL         Version 6.94
  
 Calculated polar for: GOE 165 (MVA MK.11) 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   0.7222   0.00995   0.00294  -0.1614  0.5369  0.3787
   0.500   0.7649   0.01133   0.00373  -0.1585  0.4170  0.4030
   1.000   0.8161   0.01176   0.00397  -0.1577  0.3825  0.4099
   1.500   0.8683   0.01215   0.00429  -0.1570  0.3620  0.4162
   2.000   0.9203   0.01255   0.00457  -0.1564  0.3461  0.4228
   2.500   0.9730   0.01288   0.00485  -0.1559  0.3342  0.4292
   3.000   1.0197   0.01354   0.00503  -0.1546  0.2604  0.4375
   3.500   1.0510   0.01600   0.00638  -0.1507  0.0657  0.4429
   4.000   1.0981   0.01682   0.00709  -0.1491  0.0354  0.4478
   4.500   1.1460   0.01752   0.00775  -0.1475  0.0089  0.4516
   5.000   1.1946   0.01810   0.00886  -0.1457  0.0105  0.4564
   5.500   1.2439   0.01852   0.00938  -0.1442  0.0141  0.4610
   6.000   1.2900   0.01928   0.01030  -0.1420  0.0151  0.4660
   6.500   1.3302   0.02054   0.01180  -0.1387  0.0149  0.4707
   7.000   1.3484   0.02325   0.01474  -0.1322  0.0127  0.4752
   7.500   1.3420   0.02663   0.01830  -0.1221  0.0104  0.4790
   8.000   1.3293   0.03093   0.02271  -0.1125  0.0081  0.4826