XFOIL         Version 6.94
  
 Calculated polar for: GOE 336 (MVA H.44) 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
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.4373   0.01200   0.00362  -0.0704  0.5880  0.0456
   0.500   0.4908   0.01189   0.00346  -0.0698  0.5738  0.0591
   1.000   0.5449   0.01189   0.00359  -0.0694  0.5600  0.1173
   1.500   0.5983   0.01191   0.00364  -0.0690  0.5468  0.1522
   2.000   0.6512   0.01185   0.00370  -0.0685  0.5353  0.2024
   3.500   0.8681   0.01099   0.00409  -0.0801  0.4836  1.0000
   4.000   0.9128   0.01108   0.00400  -0.0778  0.4287  1.0000
   4.500   0.9564   0.01163   0.00418  -0.0756  0.3560  1.0000
   5.000   0.9819   0.01402   0.00546  -0.0710  0.1634  1.0000
   5.500   1.0084   0.01622   0.00696  -0.0665  0.0479  1.0000
   6.000   1.0448   0.01741   0.00804  -0.0632  0.0079  1.0000
   6.500   1.0850   0.01822   0.00895  -0.0605  0.0080  1.0000
   7.000   1.1232   0.01912   0.01000  -0.0575  0.0087  1.0000
   7.500   1.1573   0.02025   0.01131  -0.0538  0.0094  1.0000
   8.000   1.1905   0.02135   0.01255  -0.0501  0.0106  1.0000
   8.500   1.2127   0.02282   0.01423  -0.0448  0.0116  1.0000
   9.000   1.2386   0.02417   0.01574  -0.0404  0.0133  1.0000
   9.500   1.2629   0.02586   0.01761  -0.0364  0.0150  1.0000
  10.000   1.2846   0.02800   0.01995  -0.0328  0.0171  1.0000
  10.500   1.3017   0.03086   0.02302  -0.0297  0.0193  1.0000
  11.000   1.3101   0.03495   0.02735  -0.0272  0.0217  1.0000
  11.500   1.3176   0.03953   0.03219  -0.0252  0.0246  1.0000
  12.000   1.3170   0.04493   0.03784  -0.0227  0.0276  1.0000