XFOIL         Version 6.94
  
 Calculated polar for: WBL FX 05-H-126 AIRFOIL                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.035 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
  -2.750   0.0935   0.04060   0.03180  -0.0302  0.4010  0.1919
  -2.500   0.1229   0.03908   0.02932  -0.0297  0.3812  0.1895
  -2.250   0.1550   0.03804   0.02759  -0.0299  0.3698  0.1885
  -2.000   0.1866   0.03728   0.02616  -0.0297  0.3623  0.1904
  -1.750   0.2182   0.03644   0.02491  -0.0293  0.3561  0.1969
  -1.500   0.2489   0.03585   0.02395  -0.0286  0.3512  0.2067
  -1.250   0.2789   0.03535   0.02308  -0.0275  0.3474  0.2169
  -1.000   0.3086   0.03474   0.02238  -0.0266  0.3446  0.2372
  -0.750   0.3379   0.03386   0.02182  -0.0259  0.3422  0.2812
  -0.500   0.3678   0.03073   0.02075  -0.0229  0.3403  1.0001
  -0.250   0.3980   0.03149   0.02067  -0.0220  0.3386  1.0001
   0.000   0.4274   0.03231   0.02105  -0.0217  0.3372  1.0001
   0.250   0.4569   0.03322   0.02163  -0.0215  0.3361  1.0001
   0.500   0.4865   0.03421   0.02237  -0.0216  0.3354  1.0001
   0.750   0.5160   0.03532   0.02328  -0.0219  0.3345  1.0001
   1.000   0.5454   0.03656   0.02434  -0.0223  0.3332  1.0001
   1.250   0.5743   0.03793   0.02555  -0.0227  0.3315  1.0001
   1.500   0.6028   0.03945   0.02692  -0.0232  0.3297  1.0001
   1.750   0.6310   0.04114   0.02848  -0.0238  0.3283  1.0001
   2.000   0.6591   0.04297   0.03028  -0.0248  0.3279  1.0001
   2.250   0.6869   0.04507   0.03259  -0.0268  0.3294  1.0001
   2.500   0.7115   0.04815   0.03603  -0.0301  0.3326  1.0001
   2.750   0.7294   0.05254   0.04078  -0.0340  0.3372  1.0001
   3.000   0.7415   0.05762   0.04609  -0.0377  0.3423  1.0001
   3.250   0.7513   0.06279   0.05139  -0.0409  0.3475  1.0001