GB2536707A

GB2536707A - Turbomachinery blade

Info

Publication number: GB2536707A
Application number: GB1505243.4A
Authority: GB
Inventors: Richard Webster John
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2015-03-27
Filing date: 2015-03-27
Publication date: 2016-09-28
Also published as: GB201505243D0

Abstract

A turbomachinery blade 40 comprising a blade 44 and a disk, where the blade comprises pressure and suction sides 46, 48 and a superelastic alloy damping member 50 located between them and in contact with both. Preferably the pressure and suction sides are unconstrained at least at a first end such that at least the first end of the pressure and suction sides are able to move relative to one another. The superelastic alloy may be a titanium alloy and may be a shape memory alloy. The superelastic titanium alloy may comprise any of titanium niobium zirconium tin alloy and titanium niobium tantalum zirconium alloy. Preferably the first and second layers are spaced by a third layer in an axial direction and preferably the damping member is bonded to one or both the pressure and suctions sides.

Description

Turbomachinerv Blade

Field of the Invention

The present invention relates to a turbomachinery blade, a gas turbine engine comprising a turbomachinery blade and a method of forming a turbomachinery blade.

Background to the Invention

Vibration of many engineering blades is a major issue leading to significant design compromises and limitations in longevity of the blade. This is particularly evident in turbomachinery blades for gas turbine engine such as blade and vane aerofoils, and has been made worse with the introduction of "blisks", in which the rotor disk and aerodynamic blade are integrally formed. Vibration is particularly problematic with blisks in view of the lack of damping that would otherwise be provided by friction between the disk and the blades.

Polymeric viscoelastic materials have been used successfully to reduce vibration, for example by applying external patches on the blade comprising a visco-elastic material.

Visco-elastic material has also been introduced to the interior of some blades, such as hollow fan blades of gas turbine engines in order to provide damping, and thereby reduce vibration, as described for example in US patent application US7311500. However, it has been found that the material is generally pushed to the tip by centrifugal forces during operation of the gas turbine engine, such that damping cannot be readily provided at the root of the blade. The low strength of the visco-elastic material also causes significant design limitations, since the outer metal layers of the blade have to provide all of the necessary strength of the blade. External patches of shape memory allow have also been used to provide damping for gas turbine blades such as turbine blades, as described for example in US2012183718.

Such blades may vibrate according to a normal mode", or two or more superposed normal modes. A normal mode is a pattern of motion of an oscillating system, in which -2 -all parts of the system move in a generally sinusoidal manner with the same frequency and with a fixed phase relation. In the art, several "mode numbers" may be identified, characterised by the number of half waves in the vibration. For example, if a vibrating blade with both ends physically constrained displays a mode shape of half a sine wave, it would be vibrating in normal mode 1 (or first order normal mode). If the blade has a full sine wave (i.e. one peak and one valley), it is said to be vibrating in normal mode 2 (or second order normal mode). Turbomachinery blades may also vibrate along several axes at once, such as along the longitudinal axis of the blade (known in the art as "flutter") and normal to the longitudinal axis (known in the art as "flap").

Where the blade experiences second or third order modes, it may be desirable to have more damping in some areas, since these areas experience larger physical movements away from the nominal position, or larger bending moments or stresses. However, such variable damping may be difficult to provide using existing, relatively fluid low strength viscoelastic materials, due to their tendency to move if the blade interior is not fully filled.

The present invention describes a turbomachinery blade and a gas turbine engine comprising a turbomachinery blade which seeks to overcome some or all of the above problems

Summary of the Invention

According to a first aspect of the present invention, there is provided a turbomachinery blade for a gas turbine engine, the blade comprising a pressure side and a suction side, wherein the blade comprises a damping member comprising a superelastic alloy, the damping member being located between and in contact with both the pressure and suction sides such that relative movement of the pressure and suction sides results in a stress in the damping member. 3 -

Advantageously, the provision of a damping member comprising a superelastic alloy between the suction and pressure sides of the turbomachinery blade has been found to provide damping of movement or vibration of the blade in use. These alloys have a high strength, and are capable of damping vibration due to their high level of hysteresis when a strain is applied. At the same time, their high strength allows them to be used as structural members in the blade, and to maintain their shape and location within the blade. By placing the damping member in contact with the pressure and suction sides, a force is applied by the first and second layers to the member, thereby resulting in increased damping, particularly for higher order vibrations. Consequently, the invention is particularly effective at reducing flutter, first order flap and other types of vibration in turbomachinery blades.

The damping member may consist essentially of a superelastic alloy. Advantageously, these materials alone may be able to provide the necessary strength, lightness and damping properties.

The superelastic alloy may comprise an alloy that deforms by a superelastic transformation, such as a titanium superelastic alloy. The term "superelastic" may be taken to mean that the transformation is a reversible phase transformation that occurs under stress between a parent phase (e.g. body-centred cubic titanium) and a product phase (e.g. hexagonal alpha titanium, orthorhombic alpha double prime titanium or hexagonal omega titanium).

The superelastic alloy may comprise a shape memory alloy (SMA) such as one of the group comprising nickel titanium alloy and copper aluminium zinc alloy.

Alternatively, the superelastic titanium alloy may comprise any of titanium niobium zirconium tin alloy or titanium niobium tantalum zirconium alloy.

The damping member may be bonded to one or both of the first and second layers, such as by diffusion bonding, activated diffusion bonding, brazing, welding or an adhesive. By bonding the damping member to the pressure and suction sides, the -4 -surfaces of the damping member in contact with the pressure and suction sides are substantially constrained such that they move with the corresponding surfaces of the pressure and suction sides, and any relative movement will result in a shear force in the damping member. Consequently, the damping member restrains movement between the pressure and suction sides, thereby forming a higher strength turbomachinery blade compared to prior visco elastic filled turbomachinery blades.

The turbomachinery blade may define a radially inner root end and a radially outer tip end. One or more of the pressure side, suction side and damping member may be radially constrained at one end, and may be substantially radially unconstrained at their other end. For example, one or more of the pressure side, suction side and damping member may be joined at their respective root to a disc.

Advantageously, the pressure side, suction side and damping member may be able to move relative to one another within the plane of the blade, with more movement being possible toward the radially outer end of the blade. This movement will be predominantly radially relative to the other sides for first modes, but can be in any direction in higher order modes, thereby producing a shear force between them. It has been found in particular that superelastic beta-titanium alloy is capable of providing a large degree of damping when a shear force is applied.

The blade may be configured such that the damping member operates within a non-linear stress strain region of the material in use. For example, the blade may be configured such that one or more of dynamic, aerodynamic or thermal forces in the operating environment produce stress between the damping member and the pressure and suction surfaces, such that the damping member operates within the non-linear stress stain region of the material in use (i.e. at the operating temperature, and other operating conditions at which vibration is expected).

The damping member may be pre-stressed, such that the damping member is subject to stress in use when substantially no load is applied to the blade. The damping member may be pre-stressed to the non-linear stress strain region of the material.

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Where the material comprises a titanium niobium zirconium tin alloy or titanium niobium tantalum zirconium alloy, the damping member may be pre-stressed to approximately MPa. The damping member may be pre-stressed by one or more of a heat treatment step and a mechanical stressing step. Advantageously, the damping member is stressed to a stress strain region in which damping is maximised in use.

The damping member may comprise a plurality of spaced elements of the material, which may be spaced from one another in an axial and / or radial direction. A first end of each element may be attached to one of the pressure side and the suction side at a first radial position, and a second end of each element may be attached to the other of the suction side and the pressure side at a different second radial position.

The or each element may extend between the pressure and suction sides and / or the leading and trailing edges in a direction substantially parallel to a direction of local peak strain for a selected vibrational mode. Advantageously, particular vibrational modes of the blade can be damped effectively. Examples of selected vibrational modes include first to fifth flap modes, and first to fifth flutter modes.

At least one of the suction and pressure sides may comprise a titanium alloy such as Ti6AI4V alloy, or an aluminium alloy. Alternatively, at least one of the suction and pressure sides may comprise a high temperature alloy such as a nickel super alloy.

The turbomachinery blade may comprise a rotor, or may comprise a stator, and may comprise one of a turbine, a compressor and a fan, and may comprise a wide chord fan blade.

The thickness of the damping member may be varied in one or more directions in order to optimise the local strain energy and hence vibration damping whilst maintaining acceptable working stresses in use. As an example, where the turbomachinery blade comprises a blade or vane having an axial direction and a radial direction, the axial thickness of the third layer may vary in a radial direction in order to maximise the local -6 -strain energy in a first flap vibrational mode. For example, the damping member may have a greater axial thickness at a tip portion relative to a root portion of the blade. Consequently, the shear stresses in use at different parts of the blade can be adjusted. In the above example, the shear stresses are greater at the root portion than at the tip portion for a given relative radial movement of the pressure and suction sides, due to the reduced thickness of the root portion relative to the tip portion. Consequently, the blade is damped to a substantially uniform degree along the radial extent of the blade where the vibrational movement comprises a first flap mode, since the relative radial movement of the blade will be greater at the tip than at the root when the blade is to vibrated in first flap mode. In higher order vibration modes, the changes in thickness may be more complex, but may follow the same general approach as described above, matching different amounts of relative movement between the pressure and suction sides with different thicknesses to produce generally even strains across the damping member. In a further example, where the target vibrational mode comprises a second flap mode, a thinner region in the damping element could be provided at the tip and root portion, relative to a relatively thicker region at a mid-portion, since the relative movement of the suction and pressure sides is greater at the tip and root relative to the middle portion when vibrated in second flap mode.

The stress may primarily comprise any of a shear stress, a compressive stress and an extensive stress.

According to a second aspect of the present invention, there is provided a method of forming a turbomachinery blade for a gas turbine engine, the method comprising: locating a damping member comprising a superelastic alloy between and in contact with a pressure side and a suction side of the turbomachinery blade.

The method may further comprise configuring the blade such that the blade has a total stress (i.e. the sum of pre-stresses and mechanical stresses in use) which lies within a non-linear stress strain region of the material. 7 -

The method may comprise pre-stressing the damping member such that, when the turbomachinery blade is in a relaxed state, the damping member is mechanically stressed such that the damping member is subject to a stress that lies within a nonlinear stress strain region of the material of the third blade, which may comprise approximately 200 MPa.

The pre-stressing step may comprise one or more of a heat treatment process and a mechanical stressing process.

Where the pre-stressing step comprises a mechanical stressing process, the pre-stressing step may comprise mechanically stretching the material prior to placing the damping member between the pressure and suction sides. The method may further comprise bonding the pressure and suction sides to the damping member subsequent to placing the damping member between the pressure and suction sides while the mechanical stretching is maintained.

Alternatively or in addition, where the pre-stressing step comprises a mechanical stressing process, the pre-stressing step may comprise bending the pressure side, suction side and damping member, subsequently bonding the pressure side, suction side and damping members together, and then subsequently releasing pressure side, suction side and damping members. Advantageously, by bonding the layers while they are bent, a pre-stress is introduced when the blade is released. In this manner, the prestress will vary along the length of the blade, thereby providing different amounts of damping at different regions of the blade.

The method may comprise bonding the damping member to the pressure and suction sides subsequent to the step of locating the damping member between the pressure and suction sides. The bonding step may comprise diffusion bonding. Subsequent to the diffusion bonding step, the pressure and suction sides may be superplastically formed to obtain a desired external profile. The superplastic forming step may comprise injecting a fluid into a space between the pressure and suction sides under superplastic conditions to inflate the blade. The damping member and / or internal surfaces of at 8 -least one of the pressure and suction surfaces may comprise a stop-off material covering part of a contact surface thereof, such that parts of the surface covered by the stop-off material do not bond to the contacting surface during the diffusion bonding step.

Advantageously, the superplastic material can form a "warren girder" arrangement between the first and third layers, thereby providing a strong yet lightweight blade that is internally damped. Such a construction is particularly suitable for a fan blade.

Features of the first and second aspects of the invention may be utilised in the other respective aspect of the invention According to a third aspect of the present invention there is provided a gas turbine engine comprising a turbomachinery blade in accordance with the first aspect of the invention.

Brief Description of the Drawings

Fig. 1 is a cross sectional view of a gas turbine engine; Fig. 2 is a perspective view of a first turbomachinery blade of the gas turbine engine of Fig. 1; Fig. 3 is a cross sectional view of the turbomachinery blade of Fig. 2 in a first position; Fig. 4 is a cross sectional view of the turbomachinery blade of Fig. 2 in a second position while vibrating in a first order flap vibrational mode; Fig. 5 is a cross sectional view of a second turbomachinery blade; Fig. 6 is a cross sectional view through the line A-A of the turbomachinery blade of Fig. 5; 9 -Fig. 7 is a cross sectional view of a third turbomachinery blade; Figs. 8a and 8b are cross sectional views of a fourth turbomachinery blade before and after an inflation step in a diffusion bonding process; Figs. 9a and 9b are a cross sectional views of a fifth turbomachinery blade in first and second position respectively of a first order flap vibrational mode; Fig. 10 is a diagram showing the relationship between stress and strain for repeated loading of a material forming part of the turbomachinery blade of Figs. 2 to 9; Figs. lla to lie show example stresses for vibration of a typical blade in first, second third, fourth and fifth vibrational modes respectively.

Detailed Description

Fig.1 shows a gas turbine engine 10. The engine 10 comprises an air intake 12 and a propulsive fan 14 that generates two airflows A and B. The gas turbine engine 10 comprises, in axial flow A, an intermediate pressure compressor 16, a high pressure compressor 18, a combustor 20, a high pressure turbine 22, an intermediate pressure turbine 24, a low pressure turbine 26 and an exhaust nozzle 28. A nacelle 30 surrounds the gas turbine engine 10 and defines, in axial flow B, a bypass duct 32.

The fan 14, compressors 16, 18 and turbines 22, 24, 26 each comprise turbomachinery blades in the form of rotors and stators. The rotors are configured to rotate about an axis and direct air in an axial direction, while the stators remove swirl from the airflow, and direct the air in an axial direction The rotors could be in the form of conventional rotors, comprising a disk having a plurality of aerodynamic blades attached at a radially outer edge thereof. Alternatively, the disk and blades could be integrally formed to form a "blisk" or "bling".

-10 -Fig. 2 shows part of a turbomachinery blade in the form of a first compressor blisk 40.

The blisk comprises a disk 42 (part of which is shown in Fig. 2) and a plurality of aerodynamically profiled blades 44 (one of which is shown in Fig. 2). The blade 44 comprises a root portion 52 at a radially inner end adjacent the disk 42, a tip portion 54 at a radially outer end proximal from the disk 42, as well as leading 56 and trailing edges 58 at axially forward and rearward ends respectively. The blade 44 defines a principle plane extending between the root and tip 52, 56 and leading and trailing edges 56, 58.

Each blade 44 comprises pressure and suction sides 46, 48 which each comprise a first titanium alloy material such as Ti6AI4V (i.e. titanium metal alloyed with 6% and 4% by weight aluminium and vanadium respectively). Alternatively, the sides 46, 48 could comprise other materials such as aluminium alloy. Where the blade is to be used in higher temperature regions of the engine 10 (such as in the turbine region), the blade could comprise higher temperature capable materials such as nickel super alloy or titanium-aluminium alloy. The invention has been found to be suitable for relatively high temperature applications up to around 400°C, and so could be used for high pressure compressor blades or stators, and low pressure turbine blades or stators. A damping member 50 comprising a a superelastic alloy is located between the sides layers 46, 48.

In this example, the damping member 50 comprises a superelastic alloy that deforms by superelastic transformation. A superelastic transformation is a substantially reversible phase transformation that occurs under stress between a parent phase (e.g. body-centred cubic titanium) and a product phase (e.g. hexagonal alpha titanium, orthorhombic alpha double prime titanium or hexagonal omega titanium). The damping member 50 has a low elastic modulus (i.e. Young's modulus), and also a low shear modulus. In one example, where the turbomachinery blade is to form part of a compressor rotor or stator of a gas turbine engine, a superelastic titanium alloy is chosen having a low modulus at temperatures up to 300° C, and in some cases up to 400°C. The low modulus and high failure stress mean that a high strain energy can be generated when loaded, whilst still retaining sufficient strength to maintain the integrity of the blade 44. The damping member 50 could comprise further materials, or could consist essentially of a superelastic material.

A first known suitable titanium superelastic alloy is "gum metal", also known as TNTZ.

TNTZ has a nominal composition comprising, in weight per cent, 23% niobium, 0.7% tantalum, 2% zirconium, and 1% oxygen, with the balance being titanium. TNTZ can exhibit superelastic properties over a range of compositions and also include vanadium and hafnium. Such alloys may be relatively strong, with an ultimate tensile strength of around 2 GPa. The Young's modulus of gum metal can range from 20 to 60 GPa. The Shear modulus also varies, and may be as low as 20 GPa.

A second known superelastic titanium alloy comprises, in weight per cent, approximately 24% niobium, 4% zirconium and 7.5 % tin, with the balance being titanium. Such a composition is sometimes referred to in the art as "Ti2448". This alloy is a 13 type titanium alloy, having a body centred cubic crystal parent phase structure. Nanostructured alloy having a grain size less than 50 nm has been produced by cold rolling, and has been found to be particularly suitable, having an ultimate tensile strength of approximately 500 to 1000 MPa. This superelastic titanium alloy has a Young's modulus below 80GPa, and preferably below 60 GPa. In one example, the Young's modulus is approximately 65 GPa, and the Poisson's ratio (i.e. the negative ratio of transverse to axial strain) of the material is approximately 0.28. A damping member 50 comprising this material may be capable of relatively large strains, up to 2% in some cases, and stresses of around 500 MPa without exceeding the elastic deformation limit. This material has been found to have a coefficient of thermal expansion of approximately 0.7 x 10-6 mm/mm/°C at 100°C, 1.0 x 10-6 mm/mm/°C at 200°C.

A third known superelastic titanium alloy comprises, in weight per cent, 28% niobium, 1% iron and 0.5% silicon, with the balance being titanium.

-12 -Alternatively, the damping member 50 could comprise a superelastic alloy in the form of a shape memory alloy (SMA) which displays superelastic properties at the operating temperatures of the blade 44. Examples of suitable SMA materials which display superelastic properties at the required temperatures for a compressor blade of a gas turbine engine include titanium nickel alloy, such as an alloy comprising substantially 50% by weight titanium and 50% by weight nickel, and copper aluminium zinc alloy.

Fig. 10 shows the stress / strain curve for the third layer 50 comprising the second above described superelastic titanium alloy comprising approximately 24% niobium, 4% zirconium and 7.5 % tin. Superelastic materials in general, and this material in particular, has been found to have a non-linear stress stain curve, having a relatively low Young's modulus (i.e. a low ratio of stress to strain when a load is applied) at a stress above a critical value. The critical value for this material is approximately 200 x106 Pascals (i.e. 200 MPa). Consequently, the third layer 50 will shear to a relatively large extent when a shear force greater than 200 MPa is applied by the first and second layers 46, 48.

As can be seen from Fig. 10, the superelastic material displays a hysteresis curve. That is to say that the material has a different stress / strain relationship when the material is loaded (as shown by arrow 70) compared to when the load is released (as shown by arrow 72). This might typically be at a stress of between 200 and 300 MPa during elastic deformation.

Fig. 3 shows the arrangement of the blade 44 in more detail, from the perspective of looking through the plane of the blade 44. Fig. 3 is a simplified drawing, which omits details such as the aerodynamic shape of the blade 44. The sides 46, 48 comprise continuous sheets of material which extend from the leading edge 56 to the trailing edge 58 of the blade 44, and from the root portion 52 to the tip portion 54 to form a continuous aerodynamic profile. The damping member 50 similarly comprises a continuous sheet of material which extends from the leading edge 56 to the trailing edge 58, and from the root 52 to the tip 54, and is in contact with an internal surface of the sides 46, 48 on either side, thereby spacing the sides 46, 48 in a generally axial -13 -direction. In other embodiments, the damping member 50 could comprise several layers of material.

The damping member 50 is bonded to the sides 46, 48, for example by any of diffusion bonding, activated diffusion bonding, brazing, welding or an adhesive. Alternatively, the sides 46, 48 could be placed close together and in contact with the damping member 50, such that movement between the pressure side 46 and the damping member 50, and between the suction side 48 and the damping member 50, is substantially prevented by friction.

Each of the components 46, 48, 50 of the blade 44 is also bonded to, attached to or integrally formed with the disk 42 at their root ends 52. Consequently, the sides 46, 48 are substantially constrained at the root 52, such that the roots of the sides 46, 48 and damping member 50 do not move to a significant degree relative to each other in use.

However, the tips of each side 46, 48 and damping member 50 are not bonded together, and the sides 46, 48 are spaced by the relatively elastic damping member 50. Consequently, the tips 54 of the sides 46, 48 are able to move relative to one another in use, but do not move to a substantial degree relative to the respective contact surfaces of the damping member 50, resulting in a shear force in the damping member 50 in a generally radial direction, as the damping member 50 is deformed by these movements.

During vibration of the blade 44 in use, the stress produced in the third layer 50 by relative movement of the first and second layers 46, 48 increases as the tip 54 moves away from the first position, as shown in Fig. 3, toward the second position, as shown in Fig. 4 when vibrated in first flap mode. The stress produced in the third layer 50 then decreases as the tip 54 moves away from the second position, as shown in Fig. 4, toward the first position, as shown in Fig. 3. However, due to the hysteresis of the material of the damping member 50 during elastic deformation, heat is dissipated during this process. Consequently, the vibrational energy is reduced, thereby damping the first vibrational mode.

-14 -It has been found that, in some cases, the stress induced by vibration is less than the critical value. For example, the stress induced in the damping member 50 by relative movement of the pressure and suction sides 46, 48 during vibration in the first mode may be approximately 100 MPa. Consequently, it may be necessary to pre-stress the damping member 50 such that the damping member 50 operates within the non-linear stress /strain region 70, 72 during use.

One way of obtaining a pre-stress in the damping member 50 may be to treat the damping member 50 prior to assembly. For example, the member 50 could be formed having a profile that does not correspond to the final desired profile to fit between the pressure and suction sides 46, 48. Subsequent to forming the member 50, the member 50 is then hot or cold formed to the final desired position, thereby providing residual stress in the member 50. For example, the damping member 50 could be bonded to the pressure and suction sides 46, 48 in a position similar to that shown in Fig. 4, with the blade 44 bent axially forward or backward. Once released, a stress would be provided in the damping member 50.

In a second example, the damping member 50 may be mechanically stressed, for example for tensioning the member 50 prior to and during bonding with the pressure and suction sides 46, 48 to provide tensile stress, or tensioning the pressure and suction sides 46, 48 prior to and during bonding to provide compressive stress in the damping member 50.

In a third example, the pre-stress could be applied to a different extent to different parts of the damping member 50. One method of achieving this would be to subject parts of the damping member 50 to heat (and / or cool parts of the damping member 50), thereby causing uneven expansion of the member 50, resulting in local stress. Once joined to the first and second layer 46, 48, this stress would be "locked in".

Advantageously, a non-uniform pattern of stresses can be provided in the member 50.

Consequently, the member 50 could provide more damping for movements in some directions compared to others, and different parts of the member 50 could provide -15 -different amounts of damping in use. This can be particularly useful where it is desired to damp particular vibrational modes, which impart stresses in particular directions in the blade 44. For example, it may be desirable to produce more damping at the root 52 of the blade than at the tip 54. In such cases, more pre-stress may be applied to the root 52 than the tip 54.

As a further alternative, or in addition, the blade 44 may be subject to non-vibrational stresses in use (i.e. after assembly), such as centrifugal and aerodynamic loads, particularly where the blade 44 is a rotor. These stresses may increase the stress in the member 50 when the blade 44 is in the first vibrational position, thereby ensuring that the third layer 50 operates within the non-linear region of the stress / strain curve when subject to vibration.

Figs. 3 and 4 illustrate a first vibrational mode known as "flapping", in which the tip 54 of the blade is unconstrained. In the first order unconstrained vibrational mode, the blade 44 moves between a first position as shown in Fig. 3, and a second position as shown in Fig. 4. In this case, the blade 44 is shown as vibrating in a generally axial direction, though vibration in a generally circumferential direction may also occur. As the sides 46, 48 are bonded to the damping member 50, relative movement between the tips of the sides 46, 48 causes shear stress in the damping member 50 as the blade 44 vibrates in the first vibrational mode. Other vibrational modes may also be present during operation, such as second and higher order unconstrained vibrational modes.

Figs. 11a to 11e show maximum displacements of the blade 44 during second, third, fourth and fifth order vibrational modes respectively. The maximum displacement in different areas of the blade 44 is shown by lines 1 to 9 on each figure. Lines 1 to 9 represent lines of equal displacement during vibrational bending of the respective order, with higher numbers representing higher maximum displacements. These vibrational modes may be observed at different vibrational frequencies for example. The direction of movement in flap vibration is generally through the plane of the blade 44 (i.e. in a generally circumferential direction). Similar modes can also be described for flutter (which essentially consists of twisting movement of the blade about a radial axis).

-16 -As can be seen in Fig. 11a, in first order vibration, maximum displacements are highest near the tip, and lowest near the root, with the lines of equal displacement extending from the leading edge 56 toward the trailing edge 58, but angled somewhat radially outwardly from the leading to trailing edge 56, 58.

In second order vibration (as shown in Fig. 11b), the lowest displacements occur at a mid-region of the blade 44, approximately half way between the root 52 and tip 54. Regions of higher displacement are again found towards the tip 54, and the maximum displacement also increases towards the root 52, although to a lesser degree. The negative sign on the constant lines of displacement toward the root 52 indicates that the direction of movement is opposite to that of the movement toward the tip 52. Constant line 0 at the mid-region of the blade 46 therefore represents an inflexion point. Again, the lines of constant displacement extend from the leading to the trailing edges 56, 58, but are angled radially outwardly to a lesser degree.

Fig. 11c shows vibration in the third order mode. Again, regions of higher displacement are located toward the tip 54, and a region of lower displacement is located at a mid-region of the blade 46. However, the lines of equal displacement show a more complex pattern, with higher displacement regions being located near the leading edge 56, and moving in an opposite direction to the tip 54.

Higher modes are shown in Figs. 11d and e. As can be seen, the mode shapes become more complex, with lines of equal displacement becoming more complex, and in some cases, multiple inflexion points are observed.

In some applications, one or more of these modes may be of more interest than others when preventing adverse effects of vibration. For example, at the operating temperatures, pressures and rotational speeds of the blade 46, one or more mode may dominate. Consequently, the regions of the blade 46 having the most displacement, and therefore requiring the most damping in order to reduce vibration, will depend on the dominant mode at the operating conditions.

-17 -Identification of the dominant mode or modes can be made using computer modelling, such as finite element analysis (FEA) or by component testing In general, the maximum shear stress in the damping member 50 will extend locally in a direction generally perpendicular to a line of equal maximum displacement for a given vibrational mode. Consequently, in mode 1 for instance, the direction of maximum shear stress is substantially radial, with a bias toward the tip 54 at the leading edge 46.

Figs. 5 and 6 show a second turbomachinery blade in the form of a second blisk 140.

The second blisk 140 is similar to the first blisk 40, having suction and pressure sides 146, 148 comprising titanium alloy, and a damping member 150 bonded to the sides 146, 148 and located therebetween.

However, the damping member 150 has a different profile to the member 50 of the first blisk 40. The damping member 150 has a varying axial thickness, having a relatively large radial thickness region 162 at a radially outer end near the tip 154, and a relatively small radial thickness region 164 at a radially inner end near the root 152.

As shown in Fig. 11a, it has been found that the stress, and therefore the differential shear movement between the first and second layers 146, 148 will be greater at the tip 154 than at the root 152 during vibration in the first mode. Consequently, by providing an axially a thinner region 164 of the third layer 150 at the root 152 relative to the thicker region 164 at the tip 154, the shear stress in the thinner region 164 is increased at the root 152 relative to a damping member having a relatively large axial thickness.

This arrangement can therefore be used to accommodate the relatively small stress at the root 152 of the blade 144 when the blade is vibrating in the first mode, whilst still achieving the desired amount of damping.

Similarly, where the higher vibrational modes are of greater importance, the dimensions of the damping member can be designed such that the greatest damping is provided in areas having the greatest stresses. In other words, thicker regions of material can be -18 -provided in areas having greater stress magnitudes. For example, where the second order vibrational mode is of greater importance, a thicker region of the damping member may be provided at both a root region and a tip region, with the mid-region having a thinner profile.

More generally, the dimensions and pre-stresses of the damping member and pressure and suction surfaces will be a consideration of the steady and oscillating external loads on the blade in use, the external geometry of the blade, all of the materials stressing and the local geometry of the damping member, pressure side and suction side.

Optimisation of the local geometry by change of thickness has been described above.

Further geometry optimisation to use the full benefit of the damping member can achieved by the use of voids to control the local stressing, in particular the shear. An isolated strip of damping material can be optimised to give an optimised stiffness, damping and component life. For example, an increase in thickness (i.e. distance between contact surfaces of the two sides of the damping element) or reduction in contact area between the damping member and pressure and suction surfaces will give a lower stiffness against relative movement between the pressure and suction surfaces, and also a lower stress in the damping member. The local energy dissipation will be dependent upon the local strain energy and the material damping of the damping member under the local stressing conditions. The blade would be designed by optimisation of these parameters to provide adequate damping whilst maintaining sufficiently low stresses to give a good component fatigue life. This local optimisation will need to take place at positions throughout the blade and for all of the vibration modes of interest. Avoidance of stress concentrations will be appropriate in many applications, by the avoidance of sharp edges for example. In practice, the optimisation of this structure is likely to require finite element analysis.

Fig. 6 shows a cross sectional view of the damping member 150 through the line A-A of Fig. 5. The damping member 150 comprises optional cut-outs 160. Consequently, the amount of damping material is thereby reduced, thereby increasing the stress for a given displacement further. Since the displacement in the regions of the cut-outs 160 is relatively small in the first vibrational mode, sufficiently high stresses are imposed on -19 -these regions to ensure that the regions of the cut-outs 160 operate within the nonlinear stress / strain region of the material, thereby leading to improved damping in these areas compared to where no cut-outs are provided.

Optionally, the cut-outs 160 could be utilised to provide further pre-stress by pressurising or evacuating the volume within the cut-outs 160. The cut-outs 160 have rounded edges to prevent stress concentration.

Fig. 7 shows part of a third turbomachinery blade in the form of a blade 240 which is optimised to damp vibrational energy from first order vibration. The pressure and suction sides (not shown) of the blade 240 are similar to those of the blades 40, 140 of the first and second embodiments. However, the damping member 250 comprises a plurality of sheets of superelastic material 266. Each of the sheets extends from a leading edge 256 to a trailing edge 258, and defines gaps in between. As can be seen, the sheets 266 are elongate, and extend in a direction generally perpendicular to the lines of equal maximum displacement for the first order vibration (see Fig. 11a), and therefore extend in the direction of maximum shear stress for the vibrational mode of interest.

A blade can therefore be designed by identifying one or more vibrational modes that require damping, identifying the corresponding pattern of maximum displacement of the corresponding one or more vibrational modes, and thereby determining the directions of maximum shear stress in the damping member extending between the pressure and suction surface of the blade. The damping member is designed to maximise damping along these orientations by for instance providing elongate strips extending in the local direction of maximum stress in the blade, to thereby damp vibration of the blade in use by inducing stress by deformation of the damping member as the blade distorts.

This technique can be used where multiple vibrational modes are of concern. For example, where the second and third modes are of concern, damping elements extending in a generally radial direction can be provided near the tips, where the primary stress is from the second vibrational mode, while further damping elements -20 -extending in a generally chordwise (i.e. extending from the leading edge to the trailing edge) can be provided approximately mid-way between the root and tip, where the primary stress arises from the third vibrational mode.

Figs. 8a and 8b show a third turbomachinery blade in the form of a blade 340. Again, the blade 340 comprises suction and pressure sides 346, 348 comprising a conventional metallic material such as titanium alloy, while a damping member 350 located between the sides 346, 348 comprises a superelastic material.

A stop-off material such as yttria is provided at spaced locations 374 on both sides of the damping member 350. The damping member 350 is placed between the sides 346, 348 and the blade 340 is then subjected to heat and pressure to diffusion bond parts of the damping member 350 to the suction and pressure sides 346, 348. The stop-off material at locations 374 prevents the damping member 350 in these areas from bonding to the sides 346, 348.

The blade 340 is then "inflated" by injecting a fluid into the cavity between the pressure and suction sides 346, 348 at a superplastic deformation temperature, thereby forming a "warren girder" structure from the damping member 350, as shown in Fig. 8b. The blade 340 is consequently strong yet lightweight, while being damped by the superelastic properties of the third layer 350. Again, the third layer may have varying thickness, varying pre-stress, and optional cut-outs in order to vary the amount of damping provided at different locations.

Figs. 9a and 9b show part of a fourth turbomachinery blade in the form of a blade 240 in first and second positions respectively.

The blade 340 is similar to the blades 40, 140, 240, 340 having first and second layers 446, 448 comprising titanium alloy, and a third layer 450 bonded to the first and second layers 446, 448 and located therebetween. However, the blade 440 has a third layer 450 of a different configuration. -21 -

The third layer 450 comprises a plurality of radially spaced elongate elements 464 comprising a superelastic material. Each element 264 is attached to the first layer 246 at a respective first radial position and to the second layer 248 at a respective second radial position. The respective first and second radial positions are radially spaced, such that the elements 264 are angled toward the blade axis in either an upstream or downstream direction.

Fig. 9a shows the blade 440 in a first, relaxed position, when it is not subject to a vibrational load in the first unconstrained vibrational mode. Each of the elongate elements 464 extends a first distance X1.

Fig. 9b shows the blade 440 in a second position, when the blade 440 is displaced by a maximum extent when vibrated in the first vibrational mode. The extent of distortion of the blade 440 has been exaggerated to illustrate the embodiment. The suction and pressure sides 446, 448 are moved in a radial direction relative to one another, such that the radial position of the attachment point of each elongate member 464 is different to that when the blade 440 is in the first position. In this second position, each of the elongate elements 464 is stretched, such that it extends a distance X2. Each of the elongate elements 464 is therefore subject to a tensile load, in addition to a shear load.

Consequently, as the blade 440 moves between the first and second positions during vibration in the first vibrational mode, the elongate elements 464 stretch and relax, and therefore undergo a tensile hysteresis cycle similar to that shown in Fig. 10. The elements 464 therefore damp vibrational movement to a greater extent compared to blade 40, since this tensile stress is in addition to any shear stresses in the elements 464.

Accordingly, the invention provides a turbomachinery blade and a method of forming a turbomachinery blade having a number of advantages. The turbomachinery blade is strong and lightweight, whilst having significant damping properties, thereby reducing the effects of vibration in use. Where the blade comprises an aerodynamic blade, the blade is less susceptible to flutter and flap, due to the improved damping properties.

-22 -The invention is particular suitable for fan blades for gas turbine engines, and in particular, wide chord fan blades. Wide chord fan blades are susceptible to damage due to vibration, which may be caused by foreign object strikes such as bird strikes, or flutter caused by some operational conditions. Consequently, the invention provides a fan blade which is less susceptible to adverse effects of bird strikes and flutter. This reduced sensitivity to flutter and bird strikes may enable a designer to design a fan blade having reduced weight or improved aerodynamic properties compared to prior designs.

113 While the invention has been described in conjunction with the exemplary embodiments described above, many equivalent modifications and variations will be apparent to those skilled in the art when given this disclosure. Accordingly, the exemplary embodiments of the invention set forth above are considered to be illustrative and not limiting. Various changes to the described embodiments may be made without departing from the spirit and scope of the invention.

For example, the turbomachinery blade could comprise a stator, and could comprise a turbine. The first and second layers could comprise different materials from one another, and could comprise non-metallic materials such as reinforce carbon fibre. The third layer could comprise a different superelastic material. Further layers could be provided. For example, the third layer could comprise a pair of sheets of superelastic material, with a further layer of non-superelastic material located between the sheets.

Claims

-23 -CLAIMS1 A turbomachinery blade for a gas turbine engine, the blade comprising a pressure side and a suction side, wherein the blade comprises a damping member comprising a superelastic alloy, the damping member being located between and in contact with both the pressure and suction sides such that relative movement of the pressure and suction sides results in a stress in the damping member.

A turbomachinery blade according to claim 1, wherein the damping member consists essentially of one a superelastic alloy.

A turbomachinery blade according to claim 1 or claim 2, wherein the pressure and suction sides are unconstrained at least at a first end such that at least the first end of the pressure and suction sides are able to move relative to one another in use.

A turbomachinery blade according to any of the preceding claims, wherein the superelastic titanium alloy comprises a shape memory alloy.

A turbomachinery blade according to any of claims 1 to 4, wherein the superelastic titanium alloy comprises any of titanium niobium zirconium tin alloy and titanium niobium tantalum zirconium alloy.

A turbomachinery blade according to any of the preceding claims, wherein the first and second layers are spaced by the third layer in an axial direction.

A turbomachinery blade according to any of the preceding claims, wherein the damping member is bonded to one or both of the pressure and suction sides. 2. 4. 5. 6. 7.

-24 - 8. A turbomachinery blade according to any of the preceding claims, wherein the blade is configured such that the damping member operates in a non-linear stress strain region of the material of the damping member.

9. A turbomachinery blade according to claim 8, wherein the damping member is pre-stressed, such that the damping member is subject to stress when substantially no load is applied.

10.A turbomachinery blade according to claim 9, wherein the damping member is pre-stressed to approximately 200 MPa.

11.A turbomachinery blade according to any of the preceding claims, wherein at least one of the pressure and suction sides comprise a titanium alloy.

12.A turbomachinery blade according to any of the preceding claims, wherein the blade comprises one of a rotor and a stator.

13.A turbomachinery blade according to claim 12, wherein the blade comprises one of a turbine and a compressor.

14.A turbomachinery blade according to any of the preceding claims, wherein an axial thickness of the damping member varies in a radial direction.

15.A turbomachinery blade according to claim 14, wherein the damping member has a greater axial thickness at a tip portion relative to a root portion.

16.A turbomachinery blade according to any of the preceding claims, wherein the damping member comprises a plurality of spaced elements.

17.A turbomachinery blade according to claim 16, wherein one or more spaced element extends in a direction substantially parallel to a direction of local peak strain for a selected vibrational mode.

-25 - 18.A turbomachinery blade according to claim 16 or claim 17, wherein one or more spaced element is attached to the pressure side at a first radial position and is attached to the suction side at a second radial position, and wherein the first and second radial positions are radially spaced from one another.

19.A method of forming a turbomachinery blade for a gas turbine engine, the method comprising: locating a damping member comprising a superelastic alloy between and in contact with a pressure side and a suction side of the turbomachinery blade.

20.A method according to claim 19, wherein the method further comprises configuring the blade to have a total stress which lies within a non-linear stress region of the material of the damping member.

21.A method according to claim 20, wherein the non-linear stress region of the material of the third blade comprises approximately 200 MPa.

22.A method according to claim 20 or 21, wherein the method comprising applying a pre-stress to the damping member.

23.A method according to claim 21, wherein the pre-stressing step may comprise one or more of a heat treatment process and a mechanical stressing process.

24.A method according to claim 23, wherein the pre-stressing step comprises mechanically stretching the material prior to placing the third layer between the first and second layer.

25.A method according to claim 24, wherein the method further comprises bonding the first and second layers to the third layer subsequent to placing the third layer between the first and second layers while the mechanical stretching is maintained.-26 - 26.A method according to any of claims 23 to 25, wherein the pre-stressing step comprises bending the first, second and third layers, subsequently bonding the first, second and third layers together, and then subsequently releasing the first second and third layers.