XFOIL         Version 6.94
  
 Calculated polar for: GOE 155 (SSW D.1) 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   0.5000   0.01337   0.00620  -0.0961  0.7516  0.0607
   0.500   0.5565   0.01235   0.00486  -0.0953  0.7326  0.0525
   1.000   0.6120   0.01130   0.00378  -0.0949  0.7145  0.0500
   1.500   0.6679   0.01080   0.00327  -0.0947  0.6939  0.0485
   2.000   0.7239   0.01059   0.00299  -0.0945  0.6705  0.0490
   2.500   0.7795   0.01048   0.00287  -0.0942  0.6418  0.0652
   3.000   0.8319   0.00887   0.00291  -0.0936  0.5939  1.0000
   3.500   0.8815   0.00962   0.00305  -0.0926  0.4920  1.0000
   4.000   0.9308   0.01060   0.00350  -0.0918  0.4099  1.0000
   4.500   0.9781   0.01183   0.00413  -0.0910  0.2975  1.0000
   5.000   1.0246   0.01315   0.00491  -0.0902  0.2167  1.0000
   5.500   1.0627   0.01559   0.00643  -0.0884  0.0480  1.0000
   6.000   1.1102   0.01660   0.00737  -0.0873  0.0231  1.0000
   6.500   1.1571   0.01762   0.00830  -0.0861  0.0081  1.0000
   7.000   1.2043   0.01852   0.00950  -0.0849  0.0071  1.0000
   7.500   1.2484   0.01971   0.01097  -0.0831  0.0070  1.0000
   8.000   1.2877   0.02127   0.01290  -0.0807  0.0071  1.0000
   8.500   1.3177   0.02350   0.01550  -0.0770  0.0074  1.0000
   9.000   1.3299   0.02650   0.01883  -0.0711  0.0077  1.0000
   9.500   1.3301   0.03048   0.02311  -0.0646  0.0081  1.0000
  10.000   1.3320   0.03524   0.02815  -0.0591  0.0087  1.0000
  10.500   1.3481   0.04074   0.03405  -0.0541  0.0096  1.0000
  11.000   1.3707   0.04704   0.04078  -0.0502  0.0098  1.0000
  11.500   1.3744   0.05246   0.04667  -0.0464  0.0091  1.0000
  12.000   1.3618   0.06105   0.05579  -0.0428  0.0096  1.0000
  12.500   1.3444   0.07296   0.06852  -0.0390  0.0144  1.0000