Abstract:
A turbomachine blade comprising a blade tip and a metallic leading edge having a plurality of shear zones angled to the blade tip, wherein the shear strength of the shear zones is less than the shear strength of the remainder of the leading edge, such that in the event of an impact shear is initially initiated at the shear zones.

Description:
FIELD OF INVENTION 
       [0001]    The present invention relates to a fan blade for a gas turbine engine. 
       BACKGROUND 
       [0002]    Gas turbine engines are typically employed to power aircraft. Typically a gas turbine engine will comprise an axial fan driven by an engine core. The engine core is generally made up of one or more turbines which drive respective compressors via coaxial shafts. The fan is usually driven directly off an additional lower pressure turbine in the engine core. 
         [0003]    A fan of the gas turbine engine generally includes a plurality of blades mounted to a hub. A fan casing and liner circumscribe the fan blades. Fan blades may be metallic or have a composite construction. Generally a composite fan blade will have a composite non-metallic core, e.g. a core having fibres within a resin matrix. Typically a composite blade will have a metallic leading edge to prevent erosion and to protect the blade against impact damage from foreign objects. The metallic leading edge generally wraps around the leading edge of the composite core and covers a portion of the suction surface of the blade and a portion of the pressure surface of the blade. 
         [0004]    In the event of the leading edge becoming detached from the remainder of the fan blade (e.g. if a fan blade is released from the hub), the construction of the leading edge means that it can apply high impact forces to the fan casing. The fan casing and liner can be designed to absorb the impact energy imparted by a released blade and leading edge. However, this generally leads to a heavy system that has associated efficiency penalties. 
       SUMMARY OF INVENTION 
       [0005]    A first aspect of the invention provides a turbomachine blade comprising blade tip and a metallic leading edge having a plurality of shear zones angled to the blade tip, wherein the shear strength of the shear zones is less than the shear strength of other regions of the leading edge. 
         [0006]    In the present application, reference to an angled plane refers to an angle greater than 0°. 
         [0007]    The shear zones may be acutely or obtusely angled to the blade tip. 
         [0008]    The provision and angling of the shear planes means that in the event of an impact, shear can be initially initiated at the shear zones. The use of a plurality of shear zones promotes progressive collapse of the leading edge so as to increase energy absorption during a fan blade off event, which in turn reduces the loading requirements for the fan case. During progressive collapse of the leading edge, each shear zone is intended to shear so that the leading edge breaks up into multiple pieces so as to absorb impact energy. 
         [0009]    For example, the shear zones may be angled in a plane defined by a spanwise and a chordwise direction of the blade. Each shear zone may define a slip plane. The slip plane may be angled. 
         [0010]    The shear strength of the shear zones may be less than the shear strength of the remainder of the leading edge. 
         [0011]    The blade may comprise a core. The core may be a composite core or a metallic core. The metallic leading edge may be provided at a leading edge or end of the core. 
         [0012]    The blade may have a leading edge and a trailing edge; and a suction surface extending between the leading edge and the trailing edge and a pressure surface extending between the leading edge and the trailing edge. In the present application, a chordwise direction is a direction extending between the leading edge and the trailing edge; a spanwise direction is a direction extending between the tip of the blade and the root of the blade; and the thickness direction is a direction extending between the pressure surface and the suction surface of the blade. 
         [0013]    The plurality of shear zones may be distributed along the leading edge in a spanwise direction. 
         [0014]    The zones may be angled towards a tip of the blade in a direction from the leading edge towards a trailing edge of the blade. For example, the angle between the zones and the tip of the blade may be acute. That is, the angle from the zone to the tip in a clockwise direction may be acute. In use, the zones may form an acute angle with a longitudinal axis of the turbomachine (e.g. the angle from the shear zone to the longitudinal axis in a clockwise direction may be acute). 
         [0015]    The zones may be angled such that, in use, the angle between the zones and the longitudinal axis of the turbomachine is equal to or between 30° and 80°. For example, equal to or less than 70° or equal to or more than 40°, e.g. 45° or 60°. 
         [0016]    The metallic leading edge may comprise a plurality of sections arranged and adjacently attached in a spanwise direction. The surfaces of the sections intended to be bonded to another section may be angled to an adjacent surface. 
         [0017]    The shear zones may include the bondline between the adjacently attached sections. The sections may be attached using an adhesive, metal bonding process, or mechanical connection. 
         [0018]    The sections may be welded together. 
         [0019]    One or more cavities may be provided in the shear zones. 
         [0020]    The leading edge may be thinner in a region of the shear zones compared to regions directly adjacent said shear zones. 
         [0021]    The shear zones may be each angled in a direction defined by a thickness direction and a spanwise direction. 
         [0022]    The metallic leading edge may include two wings and a fore portion provided between the two wings. 
         [0023]    The leading edge may be formed in two portions, the two portions being connected together in the fore portion. 
         [0024]    Each of the shear zones in the region of the wings may be angled in a direction defined by a thickness direction and a spanwise direction. 
         [0025]    The shear zones may be angled on each wing such that an innermost position of the shear zone is nearer to the blade tip than an outermost position. 
         [0026]    The leading edge may be made by additive manufacture. Alternatively, by way of example only, the leading edge may be machined from solid, forged, cast or metal injection moulded. 
         [0027]    The metallic leading edge may comprise a plurality of sections arranged in a spanwise direction, the sections being connected together by a connection having a weaker shear strength than said sections, and the connection between the sections being angled to a tip of the blade. 
         [0028]    The surfaces of the sections intended to be bonded to an adjacent section may be angled to an adjacent side, such that the surface is angled to the blade tip. 
         [0029]    The sections may be welded together. 
         [0030]    A second aspect of the invention provides a gas turbine engine comprising a fan and a fan case that circumscribes the fan, wherein the fan comprises a plurality of blades according to the first aspect. 
     
    
     
       DESCRIPTION OF DRAWINGS 
         [0031]    The invention will now be described, by way of example only, with reference to the accompanying drawings in which: 
           [0032]      FIG. 1  illustrates a gas turbine engine; 
           [0033]      FIG. 2  illustrates a composite fan blade; 
           [0034]      FIG. 3  illustrates a partial cross section of the composite fan blade of  FIG. 2 ; 
           [0035]      FIG. 4  illustrates a section view of a leading edge of the fan blade of  FIG. 2 ; 
           [0036]      FIG. 5  illustrates a schematic of the fan blade of  FIG. 2  viewed from the pressure side and illustrates the position of a plurality of shear zones; 
           [0037]      FIGS. 6 and 7  illustrate a partial sectional view of the leading edge of  FIG. 4  along the line S-S; and 
           [0038]      FIG. 8  illustrates a partial view of the leading edge of  FIG. 7  in the direction of the arrow X. 
       
    
    
     DETAILED DESCRIPTION 
       [0039]    With reference to  FIG. 1  a bypass gas turbine engine is indicated at  10 . The engine  10  comprises, in axial flow series, an air intake duct  11 , fan  12 , a bypass duct  13 , an intermediate pressure compressor  14 , a high pressure compressor  16 , a combustor  18 , a high pressure turbine  20 , an intermediate pressure turbine  22 , a low pressure turbine  24  and an exhaust nozzle  25 . The fan  12 , compressors  14 ,  16  and turbines  20 ,  22 ,  24  all rotate about the major axis of the gas turbine engine  10  and so define the axial direction of the gas turbine engine. 
         [0040]    Air is drawn through the air intake duct  11  by the fan  12  where it is accelerated. A significant portion of the airflow is discharged through the bypass duct  13  generating a corresponding portion of the engine thrust. The remainder is drawn through the intermediate pressure compressor  14  into what is termed the core of the engine  10  where the air is compressed. A further stage of compression takes place in the high pressure compressor  16  before the air is mixed with fuel and burned in the combustor  18 . The resulting hot working fluid is discharged through the high pressure turbine  20 , the intermediate pressure turbine  22  and the low pressure turbine  24  in series where work is extracted from the working fluid. The work extracted drives the intake fan  12 , the intermediate pressure compressor  14  and the high pressure compressor  16  via shafts  26 ,  28 ,  30 . The working fluid, which has reduced in pressure and temperature, is then expelled through the exhaust nozzle  25  generating the remainder of the engine thrust. 
         [0041]    The intake fan  12  comprises an array of radially extending fan blades  40  that are mounted to the shaft  26 . The shaft  26  may be considered a hub at the position where the fan blades  40  are mounted. The fan blades are circumscribed by a fan casing  39 . The fan casing includes a liner proximal to the fan blades. 
         [0042]    In the present application a forward direction (indicated by arrow F in  FIG. 3 ) and a rearward direction (indicated by arrow R in  FIG. 3 ) are defined in terms of axial airflow through the engine  10 . 
         [0043]    Referring to  FIG. 2 , the fan blades  40  each comprise an aerofoil portion or core  42  having a leading edge  44 , a trailing edge  46 , a concave pressure surface  48  extending from the leading edge to the trailing edge and a convex suction surface (not shown in  FIG. 2  but indicated at  50  in  FIG. 3 ) extending from the leading edge to the trailing edge. The fan blade has a root  52  via which the blade can be connected to the hub. The fan blade has a tip  56  at an opposing end to the root. The fan blade may also have an integral platform  54  which may be hollow or ribbed for out of plane bending stiffness. The fan blade includes a metallic leading edge  44  covering the leading edge of the core and extending along a portion of the pressure surface and suction surface of the core. The fan blade also includes a metallic trailing edge covering the trailing edge of the core and extending along a portion of the pressure surface and the suction surface of the core. 
         [0044]    In the present application, a chordwise direction C is a direction extending between the leading edge and the trailing edge; a spanwise direction S is a direction extending between the tip of the blade and the root  52  of the blade  40 ; and the thickness direction T is a direction extending between the pressure surface  48  and the suction surface  50  of the blade  40 . 
         [0045]    Referring now to  FIGS. 3 and 4 , the metallic leading edge  44  includes a fore portion  58  provided between two wings  60 ,  62 . One of the wings  60  extends partially along the suction side of the core  42  and the other of the wings  62  extends partially along the pressure side of the core. 
         [0046]    Referring to  FIG. 5 , the leading edge  44  includes a plurality of shear zones defining slip planes  70 . Four slip planes are illustrated in  FIG. 4 , but the number of slip planes may be more or less than four, and the number of slip planes can be selected to promote the desired leading edge failure in the event of the fan blade being released from the fan during use. The slip planes extend the full chordal length of the leading edge, but in alternative embodiments the slip planes may be limited to the fore portion and a forward region of the wings. 
         [0047]    The slip planes  70  are angled to the root  52  or the tip  56  of the blade. In the present embodiment, the slip planes are angled towards the tip in a direction from the leading edge to the trailing edge, in this way, in use, the slip planes are acutely angled (indicated by angle α) to an axis parallel to the longitudinal axis of the gas turbine engine  10 . The angle of the slip planes can be selected to achieve the desired failure mode for a given blade and casing design. 
         [0048]    Referring to  FIGS. 6 to 8 , the slip planes  70  may be formed in a number of different ways. In one embodiment, the leading edge  44  may be formed from a plurality of sections  72  adjacently stacked in a spanwise direction and attached (e.g. welded) together. In such embodiments the slip planes  70  may be defined by the bondline between the stacked sections. The bondline may be treated and/or may have a pattern of bonded and non-bonded areas so as to adapt the bondline to shear at a desired predetermined shear load. 
         [0049]    Additionally or in alternative embodiments, the leading edge  44  may be thinned in the region of the slip planes  70 . For example, a groove  74  on the outer and/or inner surface of the leading edge may be provided. Alternatively, the weld relief (or weld preparation) may contribute to thinning the leading edge. 
         [0050]    Referring in particular to  FIG. 8 , as well as being angled in a leading edge to trailing edge direction (e.g. a chordwise direction C), the slip planes  70  may also be angled in a thickness direction. In such embodiments, local thinning of the leading edge  44  (e.g. by providing grooves  74  or cavities) may be used to angle the slip planes in the thickness direction. In the embodiment illustrated in  FIG. 8 , the slip planes are angled so that the slip plane on the inner side of each wing  60 ,  62  is angled towards the blade tip. 
         [0051]    Referring back to  FIG. 4 , in the present embodiment the metallic leading edge  44  is formed of a first portion  64  and a second portion  66  connected together in a region of the fore portion  58 . Each of the first portion and the second portion form one wing  60 ,  62  and part of the fore portion  58 . In the present embodiment, the first portion and the second portion connect in a central region of the fore portion. However, in alternative embodiments the leading edge may be formed as a single component. 
         [0052]    To manufacture a blade  40  of the described embodiment, the first portion and the second portion may be made using additive manufacture, machining from solid or any suitable metal forming method. The first portion may be bonded to the second portion using welding, e.g. electron beam welding, or diffusion bonding. 
         [0053]    In embodiments where the shear planes are defined by a weld between two sections, the heat treatment applied to the weld may be selected so as to achieve the desired shear strength in the region of the weld. In exemplary embodiments, heat treatment may be omitted. 
         [0054]    It will be appreciated by one skilled in the art that, where technical features have been described in association with one or more embodiments, this does not preclude the combination or replacement with features from other embodiments where this is appropriate. Furthermore, equivalent modifications and variations will be apparent to those skilled in the art from this disclosure. Accordingly, the exemplary embodiments of the invention set forth above are considered to be illustrative and not limiting. 
         [0055]    For example, in the described embodiments the leading edge is manufactured in two portions that are bonded together. However, in alternative embodiments the leading edge may be made as a single component. When the leading edge is manufactured as a single component the leading edge may be manufactured using additive layer manufacture and the shear zones may be provided by including voids, cavities and/or thinned regions in the leading edge. 
         [0056]    In the present application the leading edge has been shown as having two wings, but in alternative embodiments the leading edge may have a “bullet” shape, that is be shaped to exclude the wings. 
         [0057]    The fan blade described is a composite blade with a metallic leading edge. However, the leading edge may be a leading edge of a metallic blade (e.g. a solid or hollow metallic blade). In such embodiments, the leading edge may be integrally formed with the core of the blade.