Patent Publication Number: US-2007098562-A1

Title: Blade

Description:
The present invention relates to a blade, and in particular to a fan blade for a turbofan gas turbine engine.  
      Small tip chord turbofan clapper less fan blades may suffer from vibration where altitude aerodynamic forces lead to excitation of a fan blades natural modes of vibration, e.g. second flap mode, away from coincidence with the harmonics of a fan blades rotational speed, i.e. a non integral vibration. At high fan blade rotational speeds, forward propagating pressure waves normal to passage shock waves are formed in the passages defined circumferentially between the radially outer tips of adjacent fan blades and bounded by the fan casing which provides useful compression of the air flow. However, at altitudes greater than about 40000 ft, 12200 m, and over specific speed ranges, greater than about 1500 fts −1 , 457 ms −1  and fan blades having a tip chord length of less than 300 mm, excitation of natural modes of vibration of the fan blades due to unsteady motion of the shock waves has led to divergent fan blade vibration.  
      These unsteady pressure waves from the normal to the passage shock propagate in an upstream direction in the passages between the tips of the fan blades in the high Mach No. flow. These unsteady pressure waves are of concern where the pressure waves have short wavelengths approximating to 0.5, 1.5, 2.5 times the chord wise length of the passage between the tips of adjacent fan blades, the passage length extends from the leading edge to the trailing edge of the fan blades. These unsteady pressure waves may provide anti-phase excitation of leading edge motion of adjacent fan blades. If there is a coincidence of the mode shape, e.g. significant leading edge motion of the fan blades within the second flap vibration mode shape, divergent blade vibration is produced, which reduces the life of the fan blades and increases the incidence of mechanical failure, e.g. cracking.  
      Accordingly the present invention seeks to provide a novel blade, which at least reduces the above problem.  
      Accordingly the present invention provides a blade comprising a root portion and an aerofoil portion, the aerofoil portion has a leading edge, a trailing edge and a tip remote from the root portion, a concave pressure surface extends from the leading edge to the trailing edge and a convex suction surface extends from the leading edge to the trailing edge, the tip of the aerofoil portion between the leading edge and the trailing edge of the aerofoil portion has a groove extending radially inwardly from the remainder of tip of the aerofoil portion and extending from the convex suction surface to the concave pressure surface, the groove in the tip of the aerofoil portion is spaced from the leading edge and is spaced from the trailing edge.  
      Preferably the groove in the tip of the aerofoil portion extends radially inwardly from the remainder of the tip of aerofoil portion in the range of 0.5% to 1.5% of the chord length of the tip of the aerofoil portion.  
      Preferably the groove in the tip of the aerofoil portion extends radially inwardly from the remainder of the tip of aerofoil portion by 1% of the chord length of the tip of the aerofoil portion.  
      Preferably the groove in the tip of the aerofoil portion extends radially inwardly from the remainder of the tip of the aerofoil portion by 3 mm.  
      Preferably the groove in the tip of the aerofoil portion extends from a position at about 40% of the chord length from the leading edge to a position at about 60% of the chord length from the leading edge.  
      Preferably the groove in the tip of the aerofoil portion extends from a position at about 45% of the chord length from the leading edge to a position at about 55% of the chord length from the leading edge.  
      Preferably the groove in the tip of the aerofoil portion extends chordally of the tip of aerofoil portion by 35 mm of the chord length of the tip of the aerofoil portion.  
      Preferably the centre of the groove is arranged at a position at about 50% of the chord length from the leading edge.  
      Preferably the blade is a fan blade.  
      Preferably the blade has a tip chord length of less than 300 mm. 
    
    
      The present invention will be more fully described by way of example with reference to the accompanying drawings in which:  
       FIG. 1  shows a turbofan gas turbine engine having a fan blade according to the present invention.  
       FIG. 2  shows a fan blade according to the present invention.  
       FIG. 3  shows an enlarged view of a tip of the fan blade shown in  FIG. 2 . 
    
    
      A turbofan gas turbine engine  10 , as shown in  FIG. 1 , comprises in flow series an inlet  12 , a fan section  14 , a compressor section  16 , a combustion section  18 , a turbine section  20  and an exhaust  22 . The fan section  14  comprises a fan rotor  24  carrying a plurality of circumferentially spaced radially outwardly extending fan blades  26 . The fan blades  26  are arranged in a bypass duct  28  defined by a fan casing  30 , which surrounds the fan rotor  24  and fan blades  26 . The fan casing  30  is secured to a core engine casing  34  by a plurality of circumferentially spaced radially extending fan outlet guide vanes  32 . The fan rotor  24  and fan blades  26  are arranged to be driven by a turbine (not shown) in the turbine section  20  via a shaft (not shown). The compressor section  16  comprises one or more compressors (not shown) arranged to be driven by one or more turbines (not shown) in the turbine section  20  via respective shafts (not shown).  
      A fan blade  26  according to the present invention is shown more clearly in  FIGS. 2 and 3 . The fan blade  26  comprises a root portion  36  and an aerofoil portion  38 . The root portion  36  is arranged to locate in a slot  40  in the rim  42  of the fan rotor  24 , and for example the root portion  36  may be dovetail shape, or firtree shape, in cross-section and hence the corresponding slot  40  in the rim  42  of the fan rotor  24  is the same shape. The aerofoil portion  38  has a leading edge  44 , a trailing edge  46  and a tip  48  remote from the root portion  36  and the fan rotor  24 . A concave pressure surface  50  extends from the leading edge  44  to the trailing edge  46  and a convex suction surface  52  extends from the leading edge  44  to the trailing edge  46 .  
      A groove  54  is provided in the tip  48  of the aerofoil portion  38  between the leading edge  44  and the trailing edge  46 . The groove  54  in the tip  48  of the aerofoil portion  38  is spaced from the leading edge  44  and the trailing edge  46 . The groove  54  in the tip  48  of the aerofoil portion  38  extends radially inwardly by a radial depth D from the remainder of the tip  48  of aerofoil portion  38  and the radial depth D is in the range of 0.5% to 1.5% of the chord length C of the tip  48  of the aerofoil portion  38 .  
      In particular the groove  54  in the tip  48  of the aerofoil portion  38  extends radially inwardly from the remainder of the tip  48  of aerofoil portion  38  by a radial depth D of 1% of the chord length C of the tip  48  of the aerofoil portion  38 . The groove  54  in the tip  48  of the aerofoil portion  38  extends radially inwardly by a radial depth D from the remainder of the tip  48  of the aerofoil portion  38  of 3 mm.  
      The groove  54  in the tip  48  of the aerofoil portion  38  extends from a position F at about 40% of the chord length C from the leading edge  44  to a position G at about 60% of the chord length C from the leading edge  44 . In particular the groove  54  in the tip  48  of the aerofoil portion  38  extends from a position F at about 45% of the chord length C from the leading edge  44  to a position G at about  55 % of the chord length C from the leading edge  44 .  
      The groove  54  in the tip  48  of the aerofoil portion  38  extends chordally of the tip  48  of aerofoil portion  38  by 35 mm of the chord length C of the tip  48  of the aerofoil portion  38 .  
      Preferably the centre of the groove  54  is arranged a distance E, at a position at about 50% of the chord length C, from the leading edge  44 .  
      The fan blade  26  has a tip chord length C of less than 300 mm.  
      The groove  54  in the tip  48  of the aerofoil portion  38  of the fan blade  26 , provides a local over the tip  48  leakage path for working fluid, air, which disrupts the forward, upstream, propagating unsteady pressure wave. The groove  54  in the tip  48  of the aerofoil portion  38  of the fan blade  26  allows a natural flow of fluid, air, from the concave pressure surface  50  to the convex suction surface  52  of the aerofoil portion  38 , which attenuates and disrupts the unsteady forward, upstream, propagating unsteady pressure waves. The dimension of the groove  54  in a chordal direction is arranged to exceed the predicted wavelength of the unsteady pressure wave. The radial depth of the groove  54  is arranged to be a minimum, while achieving useful attenuation without compromising other aerodynamic performance factors. The groove  54  is arranged within the tip  48  of the aerofoil portion  38  to suit a predicted peak of unsteady amplitude of the forward, upstream, propagating pressure wave and may for example be at the mid-chord position, or at other suitable positions, in the tip  48  of the aerofoil portion  38 .  
      The groove  54  in the tip  48  of the aerofoil portion  38  of the fan blade  26  disrupts the unsteady pressure wave reinforcing the divergent non-integral fan blade  26  vibration at high speed and high altitude operation. This leads to increased life of the fan blade  26  and reduces the possibility of mechanical failure of the fan blade  26  under high altitude cruise conditions.  
      The present invention is applicable to clapperless fan blades which lead to excitation of other natural modes of vibration, e.g. first flap mode, third flap mode, first torsion mode, second torsion mode or combinations thereof or any of the first ten fundamental vibration modes. The present invention is applicable to metal fan blades and hybrid structured fan blades e.g. composite fan blades. In the case of some designs of hybrid structured fan blades there may be other natural modes of vibration that are not easy to describe using first flap mode, second flap mode, third flap mode, first torsion mode or second torsion mode because the complex structure of these hybrid structured fan blades may distort such mode shapes out of recognition.