XFOIL         Version 6.94
  
 Calculated polar for: GOE 134 (MVA H.12) 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   0.4064   0.01071   0.00330  -0.0773  0.6700  0.0268
   0.500   0.4627   0.01053   0.00304  -0.0774  0.6411  0.0313
   1.000   0.5189   0.01034   0.00280  -0.0777  0.6120  0.0436
   1.500   0.5753   0.01021   0.00258  -0.0779  0.5812  0.0464
   2.000   0.6316   0.01024   0.00249  -0.0780  0.5488  0.0515
   2.500   0.6799   0.00822   0.00261  -0.0769  0.5195  1.0000
   3.000   0.7354   0.00856   0.00273  -0.0770  0.4894  1.0000
   3.500   0.7906   0.00893   0.00292  -0.0771  0.4578  1.0000
   4.000   0.8428   0.00956   0.00318  -0.0768  0.3871  1.0000
   4.500   0.8940   0.01040   0.00356  -0.0765  0.3083  1.0000
   5.000   0.9447   0.01132   0.00406  -0.0761  0.2376  1.0000
   5.500   0.9889   0.01312   0.00507  -0.0750  0.1064  1.0000
   6.000   1.0314   0.01511   0.00655  -0.0733  0.0049  1.0000
   6.500   1.0825   0.01580   0.00734  -0.0727  0.0052  1.0000
   7.000   1.1327   0.01655   0.00821  -0.0718  0.0060  1.0000
   7.500   1.1810   0.01746   0.00927  -0.0707  0.0069  1.0000
   8.000   1.2274   0.01852   0.01053  -0.0691  0.0083  1.0000
   8.500   1.2716   0.01971   0.01192  -0.0671  0.0097  1.0000
   9.000   1.3057   0.02168   0.01410  -0.0636  0.0105  1.0000
   9.500   1.3330   0.02383   0.01647  -0.0591  0.0120  1.0000
  10.000   1.3433   0.02646   0.01928  -0.0520  0.0131  1.0000
  10.500   1.3425   0.03015   0.02315  -0.0443  0.0149  1.0000
  11.000   1.3630   0.03406   0.02712  -0.0393  0.0180  1.0000
  11.500   1.3328   0.02968   0.02343  -0.0270  0.0239  1.0000