GB2551527A

GB2551527A - Method of producing a gas turbine engine component with an abrasive coating

Info

Publication number: GB2551527A
Application number: GB1610767.4A
Authority: GB
Inventors: Pattinson Glen; Pallett Lloyd
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2016-06-21
Filing date: 2016-06-21
Publication date: 2017-12-27
Also published as: GB201610767D0

Abstract

A method of producing a gas turbine engine component 40 with an abrasive coating includes the steps of providing a preform 28 including a backing plate 30 coated with a composite layer 32 of abrasive particles 34 embedded in a retaining matrix 36, and attaching the backing plate to a gas turbine engine component such that the composite layer forms an abrasive coating on the component. The abrasive coating can reduce or prevent propagation of cracks into the component. The backing pate may be attached by brazing or gluing, and the composite layer may be produced by providing a powder mix of abrasive particles and matrix precursor on the backing plate, and melting the precursor. The composite layer may be produced by electroplating the matrix. The preforms may be cut from an oversized plate, and the preform may be heat treated prior to attaching it to the component. The abrasive coating may be heat treated after attachment. Also claimed is a gas turbine engine component produced by such a method, and a gas turbine engine having that component.

Description

METHOD OF PRODUCING A GAS TURBINE ENGINE COMPONENT WITH AN

ABRASIVE COATING

Field of the Invention

The present invention relates to a method of producing a gas turbine engine component with an abrasive coating.

Background

Gas turbine engines have numerous sealing elements to limit or control air movement within the engine.

For example, such elements can control leakage of hot working gas from the working gas annulus to disc cavities in the engine that contain cooling fluid. One method of forming a rotating seal is to reduce the clearance between adjacent components at a seal fin. This approach is used, for example, in labyrinth seals, which create resistance to airflow by forcing the air to traverse a series of such fins.

Another example is the use of such elements to limit gas flow leakage in the gaps between the tips of rotor blades and surrounding casings.

To obtain a small clearance between components, an abrasive coating may be applied to the seal fins or rotor blade tips. For example, the coating may include abrasive particles of cubic boron nitride, embedded in a retaining matrix. The component can then be installed such that the abrasive coating is in contact with a runner surface of the adjacent facing component. When the component having the abrasive coating rotates, the boron nitride particles abrade the softer material of the runner surface such that one or more grooves are formed in the runner surface, providing a tight clearance between the components.

Several methods may be used to apply such abrasive coatings. These include: electroplating the abrasive particles to the component; laser melting the component surface to incorporate the abrasive particles into the material of the component; and brazing a braze material mixed with abrasive particles onto the component.

However, all of these methods may place abrasive particles in direct contact with or close to the component. When the component is exposed to cyclic loading, the abrasive particles may act as stress raisers and initiate cracks which can propagate into the component. Such cracks may eventually lead to fatigue failure of the component. Testing has shown that such particles may result in a 60% reduction in the component’s lifespan.

Summary

In general terms, the present invention provides a method of producing a gas turbine engine component with an abrasive coating which can reduce or prevent propagation of cracks into the component.

Accordingly, in a first aspect, the present invention provides a method of producing a gas turbine engine component with an abrasive coating, the method including steps in the order of: providing a preform including a backing plate coated with a composite layer of abrasive particles embedded in a retaining matrix; and attaching the backing plate to a gas turbine engine component such that the composite layer forms an abrasive coating on the component.

In this way, cracks nucleated within the composite layer, and particularly cracks emanating from a stress-concentrating feature or particle at least partially contained within the composite layer, are prevented from propagating into the gas turbine engine component. Advantageously, the backing plate may act to provide a crack-arresting layer between the composite layer and the gas turbine engine component. The backing plate may be sufficiently ductile to blunt cracks initiated within the composite layer. In a similar manner, the incorporation of the backing plate between the preform and the gas turbine engine component may prevent diffusion of elements between the composite layer and the gas turbine engine component.

Additionally, the blade tip may be manufactured in separation to the gas turbine engine component. Two or more such tips may be manufactured concurrently in a separate process. Thus, the gas turbine engine component and blade tip may be assembled as part of original component manufacture, or the blade tip replaced according to requirements.

In a second aspect, the present invention provides a gas turbine engine component produced by a method according to the first aspect. For example, such a component may have a backing plate attached thereto, the backing plate being coated with a composite layer of abrasive particles embedded in a retaining matrix to form an abrasive coating on the component.

Another aspect of the present invention provides a gas turbine engine having a component according to the second aspect.

Optional features of the invention will now be set out. These are applicable singly or in any combination with any aspect of the invention.

The composite layer may be of a thickness greater than the longest dimension of the abrasive particles.

In this way, abrasive particles may extend through, or be partially contained within the retaining matrix so as to project from the surface of the retaining matrix in order to provide an abrasive effect. Alternatively, the abrasive particles may be dispersed and/ or contained within the matrix so that abrasion of the composite layer reveals further abrasive particles. The revealing of further abrasive particles can provide an added abrasive effect and prevent rapid wear of the composite layer should abrasive particles become worn or dislodged from the blade tip.

Additionally or alternatively, the number of abrasive particles protruding from the matrix may be controlled and/or varied according to requirements.

The backing plate may be attached to the component by brazing or gluing.

The providing step may further include the sub-steps of: covering a backing plate with a powder mixture of abrasive particles and precursor particles of the retaining matrix; and melting the precursor particles to form the retaining matrix of the composite layer coating on the backing plate. The precursor particles may be laser melted to form the retaining matrix.

Alternatively, the providing step may include the sub-steps of: forming the composite layer by electroplating the retaining matrix. For example, an entrapment layer may be formed (e.g. electroplated) immediately beneath the composite layer, undersides of the abrasive particles of the composite layer being entrapped in the entrapment layer to hold the composite particles in position on the backing plate before encapsulation of the abrasive particles in the retaining matrix. The entrapment layer may be formed of nickel.

The backing plate may be a region of an oversized plate which is coated with the composite layer, and the providing step may include a sub-step of cutting the preform from the oversized plate. In this way, plural preforms may be cut from the oversized plate.

The method may further include the step of heat-treating the preform before the step of attaching the backing plate to the component. The heat-treatment can relieve stresses in the preform and/or improve the microstructure of the preform.

The method may further include the step of heat-treating the abrasive coating after the step of attaching the backing plate to the component. This heat-treatment can relieve stresses in the abrasive coating and/or improve the microstructure of the component and/or the coating

The retaining matrix may be nickel, cobalt, iron or an alloy of any one or more thereof.

The abrasive particles may be cubic boron nitride, titanium carbide, diamond, alumina, sapphire and/or zirconia particles. Other suitable abrasives particles are also known to the skilled person.

The gas turbine engine component may be made of a nickel-based superalloy, such as In718, Nimonic 75 or Nimonic 102, steel, or a titanium alloy, such as Ti-6AI-4V.

The backing plate may be a nickel-based superalloy, steel, or a titanium alloy. For example, the backing plate and the component can be made of the same material.

When the backing plate is attached to the component by brazing, the braze material may be any material that is suitable for joining the respective materials of the component and the backing plate in a gas turbine engine environment. Such materials are known to the skilled person. In general the braze material does not contain copper, as this is known to be detrimental in gas turbines.

However, the gas turbine engine component may be made of may be a plastic matrix composite material, such as a carbon or glass-fibre reinforced plastic. In this case the backing plate can be made of a nickel-based superalloy, steel, or a titanium alloy.

The gas turbine engine component may have one or more seal fins, a respective backing plate being attached to the, or each, seal fin. The one or more seal fins may form part of a labyrinth seal.

The gas turbine engine component may be a rotor blade or a blisk, the backing plate being attached to the tip of the rotor blade or respective backing plates being attached to the tips of the rotor blades of the blisk.

Brief Description of the Drawings

Embodiments of the invention will now be described by way of example with reference to the accompanying drawings in which:

Figure 1 shows a longitudinal cross-section through a ducted fan gas turbine engine;

Figure 2 shows schematically a cross-section through an abrasive coating on a tip of a rotor blade; and

Figure 3 shows schematically an oversized plate from which preforms are cut.

Detailed Description and Further Optional Features

With reference to Figure 1, a ducted fan gas turbine engine incorporating the invention is generally indicated at 10 and has a principal and rotational axis X-X. The engine comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, an intermediate pressure turbine 17, a low-pressure turbine 18 and a core engine exhaust nozzle 19. A nacelle 21 generally surrounds the engine 10 and defines the intake 11, a bypass duct 22 and a bypass exhaust nozzle 23.

During operation, air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the intermediate-pressure compressor 13 and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediate-pressure compressor 13 compresses the air flow A directed into it before delivering that air to the high-pressure compressor 14 where further compression takes place.

The compressed air exhausted from the high-pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines 16, 17,18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines respectively drive the high and intermediate-pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.

The engine 10 contains rotor blades, and the tips of these blades may be coated in an abrasive coating which forms part of a rotating seal that reduces or controls airflow between the blade tips and the surrounding casing.

Figure 2 shows schematically a cross-section through a tip of a rotor blade according to the present invention. The rotor blade can be made of a nickel-based superalloy such as In718, Nimonic 75 or Nimonic 102, steel, a titanium alloy such as Ti-6AI-4V, or even polymer-based composite material. Superalloys are typically used in the turbine section and later stages of the compressor section, while titanium alloy or steel may be used in the compressor section. Fan rotor blades may be made of titanium alloy, or carbon and/or glass fibre reinforced plastic. The tip of the blade has an abrasive coating applied by means of a preform 28.

The preform 28 has a backing plate 30 coated with a composite layer 32. Typically the backing plate is made from a similar material to that of the rotor blade, as this helps to match physical properties, such as thermal expansion coefficient, across the interface between the backing plate and the blade. However, it may be formed from a different material. The material choice can depend on factors such as machinability and ease of attachment of the composite layer to the backing plate.

The composite layer 32 can include, for example, abrasive cubic boron nitride particles 34 embedded in a retaining matrix 36. In some examples, the composite layer 32 is of a thickness which is less than the longest dimension of the abrasive particles 34 so that one or more abrasive particles 34 protrude from the retaining matrix 36 in order to provide an abrasive effect. In further examples, the composite layer 32 is of a thickness equal to or greater than the longest dimension of the abrasive particles 34 so that one or more abrasive particles 34 are completely contained within the retaining matrix 36. Where the composite layer 32 is of a thickness equal to or greater than the longest dimension of the abrasive particles 34, it is required that one or more abrasive particles 34 protrude from the retaining matrix 36 in order to provide an abrasive effect.

In this way, the thickness of the composite layer 32 and/ or particle size distribution can vary so as to allow only a particular percentage (by number) of abrasive particles 34 to protrude from the retaining matrix 36. In some examples, the composite layer 32 is of a thickness which is greater than the longest dimension of the abrasive particles 34 so that greater than about 50% (by number) of the abrasive particles 34 are completely contained within the retaining matrix 36. Alternatively, the composite layer 32 is of a thickness such that at least about 20% (by number) of the abrasive particles 34 are completely contained within the retaining matrix 36. Alternatively, the composite layer 32 is of a thickness such that at least about 5% (by number) of the abrasive particles 34 are completely contained within the retaining matrix 36. In a final example, the composite layer 32 is of a thickness such that at least about 1% (by number) of the abrasive particles 34 are completely contained within the retaining matrix 36.

Such a composite layer 32 is compatible for use in a high temperature environment with a superalloy blade. However, different types and/ or sizes of abrasive particles may be used.

In this way, the percentage (by number) of abrasive particles 34 completely contained within the composite layer 32 can vary according to thickness of the composite layer 32 and/ or the particle size distribution. During use, the abrasive particles enable the blade tip to abrade a groove in the casing surrounding the blade as it rotates, forming a rotating seal that reduces unwanted airflow between the blade tip and casing. Such abrasive particles 34 can be embedded within the matrix 36 so that abrasion of the composite layer 32 reveals further abrasive particles 34. The revealing of further abrasive particles 34 can provide added abrasive effect and prevent rapid wear of the composite layer 32 should abrasive particles 34 become worn or dislodged from the matrix 36.

The preform 28 can be produced by covering the backing plate 30 with a powder mixture of abrasive (e.g. cubic boron nitride) particles 34 and matrix precursor (e.g. Co-CrC) particles. The precursor particles are then melted using a laser to form the retaining matrix 36. Advantageously, producing the preform independently of the component allows the preform to be stockpiled. It also allows the production of a universal tip that can be cut to any required shape and size. The preforms can then be used to both replace worn abrasive coatings or to form abrasive coatings on new components. Further, using laser melting techniques can improve the cost effectiveness, efficiency and accuracy of preform production.

To attach the preform 28 to the component, the backing plate 30 is brazed to the tip of the rotor blade 40. This forms a braze layer 38 between the tip of the blade and the backing plate. Advantageously, as the backing plate is thus incorporated in the component beneath the composite layer 32, it can help to reduce or prevent the propagation of cracks from the composite layer 32 into the body 40 of the rotor blade. The lifespan of the blade may therefore be increased. Particularly in polymer composite systems, alternative methods of attaching the backing plate to the rotor blade may be used, such as gluing.

To further improve the performance of the abrasively coated blade, the preform 28 may be heat treated before attachment to the blade. Advantageously, heat treatment may improve the microstructure of and/or relieve stress in the preform. Similarly, the abrasive coating, formed after the backing plate has been brazed to the blade, may be heat treated. This may relieve stress in the abrasive coating, and/or improve the microstructure of the abrasive coating and/or component.

Advantageously, multiple preforms can be simultaneously produced to increase the efficiency and cost effectiveness of production. For example, Figure 3 shows schematically an oversized plate 42 for use in forming plural of the preforms. The oversized plate 42 is covered in a layer 44 of a powder mixture of abrasive particles (e.g. cubic boron nitride particles) and matrix precursor particles (e.g. Co-CrC particles), the precursor particles then being melted to form the retaining matrix for the abrasive particles. Individual preforms are then cut from respective regions 46 of the oversized plate using e.g. a laser or water jet.

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. Thus, the invention is not limited to rotor blade applications but may be used for other applications. For example, in a gas turbine engine context, the abrasive coating can be usefully applied to the tips of seal fins located on a gas turbine engine component. The abrasive coating thereby enhancing the ability of the fins to abrade a facing runner surface.

In the case of seal fins, the fins may form part of a labyrinth seal, wherein resistance to airflow is created by forcing the air to traverse through a series of fins. As another example, the abrasive coating can be formed by electroplation, e.g. by electroplating an entrapment layer which entraps undersides of the abrasive particles of the composite layer to hold them in position on the backing plate, followed by electroplating of the retaining matrix to encapsulate the abrasive particles and complete the composite layer. For example, Praxair Surface Technologies TBT406™ electroplating process or Abrasive Technologies ATA3C™ electroplating process may be used.

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.

Claims

1. A method of producing a gas turbine engine component (40) with an abrasive coating, the method including steps in the order of: providing a preform (28) including a backing plate (30) coated with a composite layer (32) of abrasive particles (34) embedded in a retaining matrix (36); and attaching the backing plate to a gas turbine engine component such that the composite layer forms an abrasive coating on the component.

2. A method according to claim 1, the composite layer being of a thickness which is greater than the longest dimension of the abrasive particles.

3. A method according to claims 1 or 2, wherein the backing plate is attached to the component by brazing or gluing.

4. A method according to any one of the previous claims, wherein the providing step includes the sub-steps of: covering a backing plate with a powder mixture of abrasive particles and precursor particles of the retaining matrix; and melting the precursor particles to form the retaining matrix of the composite layer coating on the backing plate.

5. A method according to claim 4, wherein the precursor particles are laser melted to form the retaining matrix.

6. A method according to claims 1 to 3, wherein the providing step includes the substep of: forming the composite layer by electroplating the retaining matrix.

7. A method according to any one of the previous claims, wherein the backing plate is a region of an oversized plate (42) which is coated with the composite layer, and the providing step includes a sub-step of cutting the preform from the oversized plate.

8. A method according to any one of the previous claims, further including the step of heat-treating the preform before the step of attaching the backing plate to the component.

9. A method according to any one of the previous claims, further including the step of heat-treating the abrasive coating after the step of attaching the backing plate to the component.

10. A method according to any of the previous claims, wherein the retaining matrix is nickel, cobalt, iron or an alloy of any one or more thereof.

11. A method according to any of the previous claims, wherein the abrasive particles are cubic boron nitride, titanium carbide, diamond, alumina, sapphire and/or zirconia particles.

12. A method according to any of the previous claims, wherein the gas turbine engine component is made of a nickel-based superalloy, steel or a titanium alloy.

13. A method according to any of the previous claims, wherein the backing plate is made of a nickel-based superalloy, steel or a titanium alloy.

14. A method according to any of the previous claims, wherein the gas turbine engine component is a rotor blade or a blisk, the backing plate being attached to the tip of the blade or respective backing plates being attached to the tips of the rotor blades of the blisk.

15. A method according to any of claims 1 to 13, wherein the gas turbine engine component has one or more seal fins, a respective backing plate being attached to the, or each, seal fin.

16. A gas turbine engine component produced by the method of any one of the previous claims.

17. A gas turbine engine having the component of claim 16.