XFOIL         Version 6.94
  
 Calculated polar for: GOE 184 (MVA H.29) 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   0.5314   0.00960   0.00212  -0.1017  0.5844  0.0773
   0.500   0.5863   0.00973   0.00222  -0.1015  0.5561  0.1027
   1.000   0.6418   0.00982   0.00225  -0.1014  0.5357  0.1094
   1.500   0.6974   0.00992   0.00233  -0.1013  0.5174  0.1181
   2.000   0.7528   0.01003   0.00245  -0.1011  0.4974  0.1349
   2.500   0.8061   0.00896   0.00278  -0.1014  0.4763  0.7313
   4.000   0.9589   0.01019   0.00341  -0.0985  0.3094  1.0000
   4.500   1.0107   0.01086   0.00385  -0.0980  0.2651  1.0000
   5.500   1.0947   0.01477   0.00638  -0.0947  0.0210  1.0000
   6.000   1.1439   0.01562   0.00719  -0.0938  0.0043  1.0000
   6.500   1.1934   0.01637   0.00802  -0.0928  0.0044  1.0000
   7.000   1.2412   0.01726   0.00903  -0.0915  0.0047  1.0000
   7.500   1.2877   0.01821   0.01013  -0.0901  0.0053  1.0000
   8.000   1.3315   0.01936   0.01144  -0.0882  0.0061  1.0000
   8.500   1.3721   0.02070   0.01296  -0.0858  0.0070  1.0000
   9.000   1.4021   0.02279   0.01529  -0.0820  0.0081  1.0000
   9.500   1.4230   0.02502   0.01777  -0.0769  0.0094  1.0000
  10.000   1.4193   0.02888   0.02186  -0.0700  0.0101  1.0000
  10.500   1.4199   0.03319   0.02642  -0.0652  0.0111  1.0000
  11.000   1.4127   0.03909   0.03239  -0.0598  0.0126  1.0000