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 =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   1.000   0.8111   0.01306   0.00505  -0.1560  0.4474  0.4331
   1.500   0.8591   0.01373   0.00543  -0.1546  0.4056  0.4395
   2.000   0.9102   0.01425   0.00580  -0.1538  0.3839  0.4489
   2.500   0.9602   0.01471   0.00617  -0.1529  0.3659  0.4582
   3.000   1.0093   0.01506   0.00633  -0.1519  0.3273  0.4657
   3.500   1.0555   0.01573   0.00651  -0.1504  0.2603  0.4708
   4.000   1.0867   0.01806   0.00775  -0.1467  0.0931  0.4745
   4.500   1.1303   0.01909   0.00868  -0.1446  0.0372  0.4797
   5.000   1.1768   0.01991   0.00994  -0.1424  0.0119  0.4850
   5.500   1.2232   0.02058   0.01073  -0.1405  0.0124  0.4902
   6.000   1.2699   0.02117   0.01147  -0.1386  0.0153  0.4952
   6.500   1.3151   0.02184   0.01239  -0.1362  0.0183  0.5014
   7.000   1.3519   0.02320   0.01404  -0.1323  0.0183  0.5083
   7.500   1.3747   0.02518   0.01633  -0.1263  0.0177  0.5136
   8.000   1.3860   0.02778   0.01922  -0.1191  0.0174  0.5193
   8.500   1.3869   0.03110   0.02279  -0.1113  0.0167  0.5252
   9.000   1.3722   0.03601   0.02783  -0.1036  0.0156  0.5304
   9.500   1.3707   0.04072   0.03281  -0.0985  0.0141  0.5369
  10.000   1.3697   0.04611   0.03833  -0.0940  0.0118  0.5450
  10.500   1.3804   0.05148   0.04384  -0.0898  0.0072  0.5561