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 =     0.030 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
  -3.000   0.0741   0.04579   0.03835  -0.0380  0.6166  0.2176
  -2.750   0.1104   0.04341   0.03541  -0.0384  0.5637  0.2123
  -2.500   0.1515   0.04157   0.03340  -0.0415  0.4728  0.2125
  -2.250   0.1733   0.03900   0.03010  -0.0365  0.4378  0.2167
  -2.000   0.1980   0.03736   0.02749  -0.0330  0.4169  0.2231
  -1.750   0.2278   0.03648   0.02566  -0.0316  0.4027  0.2300
  -1.500   0.2585   0.03569   0.02439  -0.0308  0.3929  0.2404
  -1.250   0.2875   0.03504   0.02341  -0.0298  0.3857  0.2611
  -1.000   0.3165   0.03432   0.02269  -0.0289  0.3804  0.2971
  -0.750   0.3426   0.03255   0.02216  -0.0278  0.3763  0.4320
  -0.250   0.4068   0.03223   0.02113  -0.0243  0.3692  1.0001
   0.000   0.4352   0.03317   0.02153  -0.0238  0.3667  1.0001
   0.250   0.4644   0.03419   0.02217  -0.0237  0.3648  1.0001
   0.500   0.4938   0.03531   0.02308  -0.0241  0.3635  1.0001
   0.750   0.5234   0.03657   0.02419  -0.0246  0.3626  1.0001
   1.000   0.5530   0.03801   0.02552  -0.0255  0.3621  1.0001
   1.250   0.5822   0.03968   0.02713  -0.0266  0.3616  1.0001
   1.500   0.6106   0.04165   0.02910  -0.0280  0.3609  1.0001
   1.750   0.6375   0.04399   0.03150  -0.0298  0.3601  1.0001
   2.000   0.6621   0.04685   0.03444  -0.0321  0.3594  1.0001
   2.250   0.6829   0.05044   0.03818  -0.0349  0.3595  1.0001
   2.500   0.6984   0.05498   0.04288  -0.0383  0.3613  1.0001
   2.750   0.7110   0.05977   0.04775  -0.0412  0.3645  1.0001