XFOIL         Version 6.94
  
 Calculated polar for: GOE 164 (MVA MK.10) 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
 ------- -------- --------- --------- -------- ------- -------
   0.500   0.8041   0.01114   0.00368  -0.1574  0.6310  0.2358
   1.000   0.8547   0.01130   0.00360  -0.1560  0.5926  0.2446
   1.500   0.9048   0.01140   0.00361  -0.1547  0.5522  0.2575
   2.000   0.9548   0.01165   0.00374  -0.1534  0.5132  0.2725
   2.500   1.0039   0.01199   0.00397  -0.1521  0.4727  0.2967
   3.000   1.0527   0.01233   0.00431  -0.1509  0.4285  0.3602
   3.500   1.0964   0.01182   0.00465  -0.1485  0.3955  1.0000
   4.000   1.1391   0.01283   0.00512  -0.1462  0.3353  1.0000
   4.500   1.1781   0.01423   0.00592  -0.1435  0.2238  1.0000
   5.000   1.2152   0.01592   0.00707  -0.1405  0.1495  1.0000
   5.500   1.2552   0.01726   0.00805  -0.1380  0.0833  1.0000
   6.000   1.2960   0.01843   0.00918  -0.1354  0.0623  1.0000
   6.500   1.3304   0.02010   0.01068  -0.1317  0.0070  1.0000
   7.000   1.3715   0.02106   0.01180  -0.1290  0.0077  1.0000
   7.500   1.4078   0.02220   0.01315  -0.1253  0.0088  1.0000
   8.000   1.4403   0.02350   0.01471  -0.1210  0.0102  1.0000
   8.500   1.4621   0.02554   0.01705  -0.1154  0.0106  1.0000
   9.000   1.4712   0.02845   0.02025  -0.1085  0.0108  1.0000
   9.500   1.4734   0.03200   0.02404  -0.1018  0.0111  1.0000
  10.000   1.4712   0.03636   0.02867  -0.0960  0.0115  1.0000
  10.500   1.4629   0.04201   0.03457  -0.0915  0.0120  1.0000
  11.000   1.4544   0.04826   0.04099  -0.0877  0.0126  1.0000
  11.500   1.4623   0.05292   0.04557  -0.0828  0.0134  1.0000