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 =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.7515   0.00953   0.00263  -0.1583  0.6033  0.1726
   0.500   0.8031   0.01001   0.00278  -0.1568  0.5634  0.2104
   1.000   0.8542   0.01023   0.00286  -0.1557  0.5244  0.2220
   1.500   0.9057   0.01045   0.00294  -0.1547  0.4855  0.2314
   2.000   0.9559   0.01081   0.00313  -0.1536  0.4381  0.2429
   2.500   1.0051   0.01126   0.00337  -0.1525  0.3887  0.2589
   3.000   1.0553   0.01167   0.00371  -0.1515  0.3596  0.2889
   3.500   1.1051   0.01197   0.00419  -0.1507  0.3279  0.4421
   5.000   1.2203   0.01546   0.00699  -0.1420  0.0671  1.0000
   5.500   1.2610   0.01677   0.00806  -0.1393  0.0057  1.0000
   6.000   1.3073   0.01747   0.00883  -0.1375  0.0060  1.0000
   6.500   1.3523   0.01823   0.00970  -0.1354  0.0072  1.0000
   7.000   1.3945   0.01917   0.01080  -0.1328  0.0077  1.0000
   7.500   1.4300   0.02053   0.01236  -0.1290  0.0074  1.0000
   8.000   1.4531   0.02246   0.01452  -0.1233  0.0070  1.0000
   8.500   1.4609   0.02527   0.01755  -0.1156  0.0068  1.0000
   9.000   1.4677   0.02818   0.02067  -0.1086  0.0065  1.0000
   9.500   1.4699   0.03171   0.02441  -0.1020  0.0062  1.0000
  10.000   1.4664   0.03622   0.02914  -0.0964  0.0058  1.0000
  10.500   1.4569   0.04199   0.03504  -0.0920  0.0054  1.0000