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 =     0.026 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
  -3.000   0.0937   0.04878   0.04140  -0.0482  0.6785  0.2407
  -2.250   0.1819   0.04262   0.03403  -0.0429  0.5353  0.2487
  -2.000   0.2171   0.04121   0.03210  -0.0428  0.4858  0.2538
  -1.750   0.2454   0.03926   0.02980  -0.0401  0.4604  0.2668
  -1.500   0.2729   0.03766   0.02773  -0.0374  0.4439  0.2868
  -1.250   0.3000   0.03632   0.02609  -0.0350  0.4324  0.3180
  -1.000   0.3265   0.03488   0.02495  -0.0332  0.4233  0.3910
  -0.750   0.3602   0.03221   0.02309  -0.0296  0.4163  1.0001
  -0.500   0.3900   0.03305   0.02290  -0.0288  0.4098  1.0001
  -0.250   0.4178   0.03389   0.02300  -0.0282  0.4046  1.0001
   0.000   0.4458   0.03486   0.02339  -0.0278  0.4006  1.0001
   0.250   0.4752   0.03623   0.02448  -0.0286  0.3973  1.0001
   0.500   0.5042   0.03778   0.02580  -0.0295  0.3948  1.0001
   0.750   0.5325   0.03956   0.02742  -0.0307  0.3930  1.0001
   1.000   0.5602   0.04173   0.02951  -0.0323  0.3919  1.0001
   1.250   0.5864   0.04461   0.03239  -0.0347  0.3917  1.0001
   1.500   0.6073   0.04905   0.03699  -0.0389  0.3932  1.0001
   1.750   0.6178   0.05534   0.04348  -0.0443  0.3963  1.0001
   2.000   0.6217   0.06184   0.05006  -0.0489  0.3997  1.0001
   2.750   0.5405   0.08999   0.07856  -0.0670  0.4342  1.0001
   3.000   0.5196   0.09643   0.08500  -0.0699  0.4499  1.0001
   3.250   0.4914   0.10307   0.09173  -0.0734  0.4774  1.0001
   4.250   0.4365   0.12304   0.11177  -0.0865  0.6703  1.0001
   4.500   0.4463   0.12599   0.11455  -0.0870  0.6705  1.0001
   4.750   0.4529   0.12843   0.11683  -0.0871  0.6695  1.0001
   5.000   0.4590   0.13069   0.11895  -0.0871  0.6674  1.0001