XFOIL         Version 6.94
  
 Calculated polar for: GOE 180 (MVA H.26) 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   0.4862   0.00967   0.00194  -0.0967  0.5786  0.0423
   0.500   0.5394   0.00890   0.00208  -0.0968  0.5529  0.3665
   1.500   0.6459   0.00793   0.00229  -0.0955  0.5134  1.0000
   2.000   0.7002   0.00823   0.00242  -0.0951  0.4972  1.0000
   2.500   0.7552   0.00847   0.00257  -0.0949  0.4830  1.0000
   3.000   0.8098   0.00874   0.00276  -0.0946  0.4693  1.0000
   3.500   0.8614   0.00907   0.00287  -0.0938  0.4168  1.0000
   4.000   0.9083   0.01000   0.00314  -0.0924  0.2928  1.0000
   4.500   0.9530   0.01141   0.00393  -0.0909  0.1928  1.0000
   5.000   0.9979   0.01278   0.00477  -0.0894  0.1056  1.0000
   6.000   1.0889   0.01523   0.00679  -0.0862  0.0047  1.0000
   6.500   1.1378   0.01598   0.00767  -0.0849  0.0051  1.0000
   7.000   1.1857   0.01679   0.00859  -0.0835  0.0058  1.0000
   7.500   1.2312   0.01781   0.00979  -0.0817  0.0066  1.0000
   8.000   1.2757   0.01887   0.01103  -0.0796  0.0078  1.0000
   8.500   1.3145   0.02038   0.01275  -0.0767  0.0092  1.0000
   9.000   1.3415   0.02267   0.01527  -0.0720  0.0103  1.0000
   9.500   1.3574   0.02513   0.01796  -0.0657  0.0113  1.0000
  10.000   1.3557   0.02882   0.02186  -0.0578  0.0127  1.0000