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 =     0.040 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
  -3.000   0.0496   0.04202   0.03306  -0.0273  0.3572  0.1802
  -2.750   0.0867   0.04083   0.03092  -0.0288  0.3472  0.1743
  -2.500   0.1203   0.03967   0.02920  -0.0292  0.3404  0.1721
  -2.250   0.1520   0.03833   0.02747  -0.0292  0.3349  0.1707
  -2.000   0.1828   0.03737   0.02604  -0.0288  0.3304  0.1705
  -1.500   0.2428   0.03578   0.02379  -0.0270  0.3242  0.1791
  -1.250   0.2734   0.03526   0.02303  -0.0261  0.3219  0.1901
  -1.000   0.3030   0.03461   0.02233  -0.0250  0.3199  0.2031
  -0.750   0.3324   0.03405   0.02177  -0.0240  0.3183  0.2269
  -0.500   0.3614   0.03287   0.02128  -0.0234  0.3170  0.3085
  -0.250   0.3914   0.03087   0.02043  -0.0203  0.3159  1.0001
   0.000   0.4211   0.03167   0.02074  -0.0198  0.3152  1.0001
   0.250   0.4505   0.03255   0.02127  -0.0197  0.3142  1.0001
   0.500   0.4799   0.03354   0.02197  -0.0196  0.3128  1.0001
   0.750   0.5091   0.03465   0.02282  -0.0197  0.3111  1.0001
   1.000   0.5380   0.03592   0.02384  -0.0199  0.3092  1.0001
   1.250   0.5666   0.03739   0.02508  -0.0201  0.3076  1.0001
   1.500   0.5956   0.03894   0.02647  -0.0206  0.3068  1.0001
   1.750   0.6253   0.04029   0.02784  -0.0215  0.3073  1.0001
   2.000   0.6551   0.04170   0.02933  -0.0227  0.3085  1.0001
   2.250   0.6840   0.04346   0.03133  -0.0245  0.3106  1.0001
   2.500   0.7112   0.04578   0.03392  -0.0268  0.3135  1.0001
   2.750   0.7354   0.04871   0.03710  -0.0293  0.3171  1.0001
   3.000   0.7569   0.05206   0.04062  -0.0317  0.3210  1.0001
   3.250   0.7772   0.05548   0.04410  -0.0334  0.3247  1.0001