US20140119883A1

US20140119883A1 - Bleed flow passage

Info

Publication number: US20140119883A1
Application number: US14/050,838
Authority: US
Inventors: Daniel KEMPF
Original assignee: Rolls Royce Deutschland Ltd and Co KG
Current assignee: Rolls Royce Deutschland Ltd and Co KG
Priority date: 2012-11-01
Filing date: 2013-10-10
Publication date: 2014-05-01
Also published as: GB201219617D0; EP2728196A2

Abstract

A liner wall insert is provided for a compressor rotor stage of a gas turbine engine. Several liner wall inserts are provided radially outboard of the tips of the rotor blades. The liner wall inserts have bleed flow channels formed therein. The bleed flow channels are arranged to remove flow from a trailing edge region of the stage and re-inject the bleed flow at an upstream region. The re-injected bleed flow alters the flow field around the tips of the rotor blades, for example the tip leakage flow. Thus, the bleed flow is used to improve the efficiency of the compressor rotor stage, and thus of the gas turbine engine.

Description

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is based upon and claims the benefit of priority from British Patent Application Number GB1219617.6 filed 1 Nov. 2012, the entire contents of which are incorporated by reference.

BACKGROUND OF THE INVENTION

1. Field of the Invention
The present invention is concerned with bleed flow passages. The invention is concerned with, for example, bleed flow passages arranged to improve the operation of a compressor rotor stage of a gas turbine engine.
2. Description of the Related Art
A conventional gas turbine engine comprises a compressor, a combustor, and a turbine. Each of the compressor and turbine generally comprises a plurality of alternating rotor and stator stages. A rotor stage has a plurality of rotor blades which, in use, rotate about an engine axis within a casing. The casing is generally provided with a liner wall which comprises the gas washed wall that forms the radially outer flow boundary for the flow through the rotor stage.
In order to allow the blades to rotate within the liner wall, a small gap must be provided between the tips of the blades and the surrounding wall, thereby providing “tip clearance”. As the rotor blades rotate flow tends to leak over the tips of the blades, generally from the pressure surface to the suction surface, through this small gap. This leakage flow may itself reduce the efficiency of the rotor stage. Furthermore, secondary flow structures, such as vortices, may be generated in and/or around the tip clearance gap as a result of the leakage flow. Such secondary flow structures may have a further detrimental impact on the efficiency of the rotor stage, especially if left uncontrolled.

OBJECTS AND SUMMARY OF THE INVENTION

It is therefore desirable to reduce the negative impact of the tip clearance leakage flow in rotor stages of a gas turbine engine. For example, it is desirable to reduce the negative impact of the tip clearance leakage flow (and associated flow structures) in a compressor rotor stage of a gas turbine engine.
According to an aspect, there is provided a liner wall insert (which may be referred to as a liner wall segment) for a compressor rotor stage of a gas turbine engine, the rotor stage having rotor blades rotatable about an engine axis. The liner wall insert comprises a gas washed surface that forms a part of an outer flow boundary inside which the rotor blades rotate during use. The liner insert also comprises a bleed flow passage extending from a bleed flow inlet to a bleed flow outlet. The bleed flow inlet and the bleed flow outlet are both formed in the gas washed surface. The bleed flow inlet and bleed flow outlet are arranged such that, when the liner wall insert is assembled in the compressor rotor stage, the bleed flow inlet is axially downstream of the bleed flow outlet. The liner wall insert may be said to have the bleed flow passage formed therein and/or the bleed flow passage may be said to be defined by the liner wall insert.
According to an aspect, there is provided a compressor rotor stage comprising a plurality of rotor blades rotatable about an engine axis. The compressor rotor stage also comprises a plurality of liner wall inserts as described and/or claimed herein. The plurality of liner wall inserts are attached together so as to define an outer flow boundary inside which the tips of the rotor blades pass during use.
Thus, the gas washed surfaces of the liner wall inserts may form the outer flow boundary. Each liner wall insert may have one or at least one bleed flow passage. Alternatively, liner wall inserts having bleed flow passage may be combined with linear wall inserts that do not have such bleed flow passages in order to define the outer flow boundary. The number of bleed flow passages may be equal to, more than, or less than the number of blades in the rotor blade stage.
According to an aspect, there is provided a method of improving the operation of a compressor rotor stage of a gas turbine engine, the compressor rotor stage comprising rotor blades extending from a root to a tip and being rotatable about an engine axis within an outer flow boundary. The method comprises bleeding flow from the compressor flow stream into a bleed flow inlet, through a bleed flow passage, and back out into the flow stream from a bleed flow outlet that is axially upstream of the bleed flow inlet. The outer flow boundary is formed by a plurality of liner wall inserts joined together to circumferentially surround the rotor blades. Both the bleed flow inlet and the bleed flow outlet are formed in the outer flow boundary. A plurality of bleed flow passages may be provided. A bleed flow passage may be formed in a plurality of the liner wall inserts.
The compressor rotor stage and/or the liner wall insert used in methods according to the invention may be as described and/or claimed herein.
Methods, liner wall inserts and compressor rotor stages as described and/or claimed herein may help to improve the flow characteristics of the flow through the compressor rotor stage and/or through other parts of the engine. For example, the overtip leakage flow may be controlled and/or reduced. This may help to improve the operational stability and/or surge margin and/or efficiency of the rotor blade stage and/or an engine comprising the liner wall insert or rotor stage.
Bleed flow may flow through the bleed flow passage in an axially upstream direction from the bleed flow inlet to the bleed flow outlet due to the pressure difference in the main flow through the compressor between the bleed flow inlet and the bleed flow outlet. In other words, the pressure in the main flow through the compressor is higher at the axially downstream bleed flow inlet than at the (relatively) axially upstream bleed flow outlet.
The flow entering the bleed flow inlet may be, or may comprise, boundary layer flow from the outer wall of the flow passage through the compressor. This flow may be re-injected into the boundary layer flow at the bleed flow outlet. This may help to re-energize the boundary layer flow at the bleed flow outlet and/or to modify the overtip leakage flow, including any secondary flow structures such as vortices. In turn, this may help to improve the performance, as described herein for example.
Aspects of the invention may comprise any one or more of the following features.
The cross sectional area of the bleed flow passage may smaller at the bleed flow outlet than at the bleed flow inlet. Thus, the cross sectional area of the bleed flow outlet may be smaller than the cross sectional area of the bleed flow inlet. The cross sectional area of the bleed flow passage may decrease from the bleed flow inlet to the bleed flow outlet, for example in a downstream direction.
Having a smaller cross sectional area at the bleed flow outlet may mean that the bleed flow speed is greater at the bleed flow outlet than the bleed flow inlet. This may further assist in re-energizing the boundary layer flow and/or modifying the flow field at the bleed flow outlet.
The gas washed surface of the liner wall insert may be a segment of a cylinder or a segment of a frusto-conical shape. The liner wall insert itself may be said to be a segment of a cylinder or a segment of a frusta-conical shape. In a compressor rotor stage according to the invention, the outer flow boundary may be cylindrical or may be frusto-conical. Where the terms cylindrical and frusto-conical are used herein, it will be appreciated that these include shapes that are substantially cylindrical and frusto-conical respectively.
The bleed flow passage of a liner wall insert may follow a path that is substantially parallel to its gas washed surface. When assembled in a compressor rotor stage having a rotational axis, the bleed flow passage may be radially outboard of the gas washed surface. The bleed flow passage may be substantially parallel to the axial-circumferential surface of a compressor rotor stage. Such arrangements may help to reduce the thickness (for example in a radial direction) of the liner wall insert.
The bleed flow passage may have a significant and/or major component in a direction that corresponds to a circumferential direction of the compressor rotor stage at the bleed flow inlet and/or the bleed flow outlet. Such a direction may generally correspond to a rotational direction of the blades of the compressor stage. This may, for example, allow the bleed flow inlet to be orientated so as to receive bleed flow most efficiently and/or the bleed flow outlet to be orientated so as to exit the bleed flow most effectively.
The bleed flow passage in a liner wall insert may be arranged such that, when the liner wall insert is inserted into a compressor rotor stage, the axial location of the bleed flow inlet corresponds to a trailing edge region of the rotor blades; and/or the axial location of the bleed flow outlet corresponds to a leading edge region of the rotor blades.
Leading edge region may mean, by way of example only, a location that is upstream of the mid-chord axial location of the rotor blades (for example at their tips). By way of further example, leading edge region may mean, for example, a location that is within 50% chord length of the leading edge (for example at the blade tips), for example within 25%, for example within 10%.
Trailing edge region may mean, by way of example only, a location that is upstream of the mid-chord axial location of the rotor blades (for example at their tips). By way of further example, trailing edge region may mean, for example, a location that is within 50% chord length of the trailing edge (for example at the blade tips), for example within 25%, for example within 10%.
The axial location of each bleed flow outlet may be (or correspond to a position that, when installed, is) upstream of the axial location of the leading edge of the rotor is blades, for example the axial location of the leading edge of the blades at the tips. This may allow the flow structure to be advantageously modified upstream of the rotor blades.
A plurality of the liner wall inserts may be joined together, for example brazed or welded together, to form a ring. This may be a particularly convenient and efficient manner of producing a liner wall for a rotor stage of a gas turbine engine that has at least one bleed passage formed therein.
A compressor rotor stage according to the invention may further comprise a casing. The casing may be radially outboard of the liner wall inserts in the compressor rotor stage. In some cases, the liner wall inserts may be considered to be a part of what is commonly referred to as a casing.
The casing and the liner wall inserts may have cooperating location features. In use, the cooperating location features are engaged so as to hold the liner wall inserts in position. Such an arrangement may be particularly convenient if the casing is a split casing, that is a casing that is formed from at least two parts that are joined together at circumferential joining locations.
Alternatively, the liner wall inserts may be joined to the casing so as to be held in position. The joining could be, for example, welding or brazing. Such an arrangement may be particularly suitable for a ring casing, that is a casing formed as a continuous ring.
According to an aspect of the invention, there is provided a gas turbine engine comprising at least one liner wall insert, for example in at least one compressor rotor stage, as described and/or claimed herein. Such a gas turbine engine may comprise a plurality of compressor stages, each having an array of rotor blades (which may be referred to as a compressor rotor stage) axially upstream of an array of stator vanes (which may be referred to as a compressor stator stage). In such an arrangement, the bleed flow inlet for one compressor rotor stage may be upstream of the leading edge of the stator vane stage of the compressor stage.
According to an aspect, there is provided a method of manufacturing a liner wall insert as described and/or claimed herein. The method comprises manufacturing a lower portion of the liner wall insert; manufacturing an upper portion of the liner wall insert; and joining the lower portion and the upper portion together. According to such a method, the lower portion comprises first surfaces (or a first surface) that form part of the bleed flow passage. The upper portion comprises second surfaces (or a second surface) that form part of the bleed flow passage. When the lower portion and upper portion are joined together, the first surfaces and the second surfaces may come together to form the bleed flow passage.
A liner wall insert may be manufactured at least in part using metal injection moulding (MIM). For example, where the liner wall insert may be manufactured in two parts (for example an upper (or radially outer) part and a lower (or radially inner) part), one or both of the parts may be manufactured using MIM. Using MIM may result in a high level of reproducibility, the ability to manufacture complex geometry (for example complex bleed flow passages within the liner wall inserts) and/or a product that requires little or no finishing before use. However, it will be appreciated that any alternative suitable method and/or technique and/or material could be used to manufacture a liner wall insert.

BRIEF DESCRIPTION OF THE DRAWINGS

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

FIG. 1 is a cross section through a gas turbine engine having a liner wall insert and compressor blade stage according to an embodiment of the present invention;

FIG. 2 is a schematic cross section in an axial-radial plane through a tip portion of a compressor blade, a casing and a liner wall insert in accordance with the present invention;

FIG. 3 is a schematic cross section in an axial-radial plane through a tip portion of a compressor blade, a casing and a liner wall insert in accordance with the present invention;

FIG. 4 is a schematic cross section perpendicular to an axial direction through a casing and liner wall inserts in accordance with the present invention;

FIG. 5 is a perspective view of an example of a liner wall insert in accordance with the invention;

FIG. 6 is a perspective view of an alternative liner wall insert in accordance with the invention;

FIG. 7 is a perspective view of a bleed flow channel;

FIG. 8 is a perspective view of a number of liner wall inserts joined together, forming a part of a ring; and

FIG. 9 is a perspective view of an upper and lower liner wall portions prior to being joined together.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

With reference to FIG. 1, a ducted fan gas turbine engine generally indicated at 10 has a principal and rotational axis X-X. The engine 10 comprises, in axial flow series, an air intake (which may be referred to as a nacelle) 11, a propulsive fan 12, an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, and intermediate pressure turbine 17, a low-pressure turbine 18 and a core engine exhaust nozzle 19. The ducted fan gas turbine engine 10 has a bypass duct 22 and a bypass exhaust nozzle 23.
The gas turbine engine 10 works in a conventional manner so that air entering through 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 16, 17, 18 respectively drive the high and intermediate pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.
The gas turbine engine 10 shown in FIG. 1 has a compressor rotor stage having a liner wall insert 100 in accordance with an aspect of the invention. Indeed, in the FIG. 3 example, four of the compressor rotor stages are shown as having liner wall inserts 100, but it will be appreciated that any number of rotor blade stages may be provided with liner wall inserts 100.
A more detailed cross-sectional view of a liner wall insert 100 is shown schematically in FIG. 2. In FIG. 2, the liner wall insert 100 is provided as part of a compressor rotor stage 200. The liner wall insert 100 has a bleed flow passage 110 provided and/or formed therein. The bleed flow passage 110 may be said to be entirely formed in the liner wall insert 100. The radially inner surface of the liner wall insert 100 is, or comprises, a gas washed surface 140. The gas washed surface 140 forms the radially outer flow boundary of the main flow through the compressor rotor stage 200. In the example show in FIG. 2, the main flow through the compressor rotor stage 200 is the core engine flow A, but it will be appreciated that the main flow could be other flow streams, such as the fan flow stream. Thus, a liner wall insert 100 may be provided to any compressor rotor stage, including a fan stage 12.
The bleed flow passage 110 has a bleed flow inlet 120 and a bleed flow outlet 130. The bleed flow inlet 120 and the bleed flow outlet 130 may be formed in the gas washed surface 140, as in the FIG. 2 example. In operation, bleed flow C is extracted from the main compressor flow A into the bleed flow inlet 120, through the bleed passage 110 and then re-injected back into the main compressor flow A through the bleed flow outlet 130. As shown in FIG. 2, the bleed flow outlet 130 is upstream (relative to the main flow A through the compressor stage 200) of the bleed flow inlet 120. In the FIG. 2 example, the axial location of the bleed flow inlet 120 corresponds to a trailing edge region of the rotor blade 210, and the axial location of the bleed flow outlet corresponds to a leading edge region of the rotor blade 210.
The compressor rotor stage 200 also comprises a plurality of rotor blades 210 which, in operation, rotate around the engine axis X-X. A tip clearance gap 190 is provided between the tip 212 of the rotor blade 210 and the gas washed surface 140 of the liner wall insert 100. In operation, flow has a tendency to leak through the tip clearance gap 190, for example from a pressure surface of the blade 210 to a suction surface. This overtip leakage flow, if left unaddressed, may create complex flow structures at and around the tips 212 of the blades 210, such as tip vortices. Such uncontrolled overtip leakage flow may thus adversely affect the efficiency of the rotor stage 200, and thus of the gas turbine engine 10.
Re-injecting the bleed flow C into the main flow at an upstream position relative to where it is removed from the main flow helps to control the overtip leakage flow and/or the flow structures resulting therefrom, thereby reducing the adverse impact of the overtip leakage flow.
The bleed flow outlet 130 may, by way of example, be upstream of the leading edge of the rotor blade 210, as in the FIG. 2 example. This may be particularly effective in controlling the overtip leakage flow and the resulting flow structures, although it will be appreciated that in other examples, the bleed flow outlet 130 need not necessarily be upstream of the leading edge of the blade 210. For example the bleed flow outlet 130 could be at, or downstream of, the leading edge of the blade 210.
The compressor rotor stage 200 shown in the FIG. 2 also comprises a casing 300. The casing 300 could be any sort of casing. The liner wall insert 100 may be attached, or connected, to the casing 300 in any suitable manner. In FIG. 2, for example, the liner wall insert 100 may be brazed to the casing 300 and/or the casing 300 may be a ring casing. A ring casing may be a casing that extends circumferentially around the entire compressor rotor stage, for example manufactured as a single part without circumferential joints.
FIG. 3 shows a further example of a rotor stage 200 having a liner wall insert 150 according to an example of the invention. The example shown in FIG. 3 has many of the features of the example shown in and described in relation to FIG. 2. Like features are given the same reference numerals in FIGS. 2 and 3 and will not be described again in relation to FIG. 3.
In FIG. 3, the liner wall insert 150 has a bleed flow passage 110, a gas washed surface 140, a bleed flow inlet 120 and a bleed flow outlet 130, any one or more of which may be as described by way of example in relation to FIG. 2. The rotor stage 200 also has rotor blades 210, each having a tip 212, with a tip clearance gap 190 formed between the gas washed surface 140 of the liner wall insert 150 and the tip 212.
The liner insert 150 in the FIG. 3 example is provided with a location feature 160 which cooperates with a corresponding location feature 410 in the casing 400. The location feature 160 is thereby used to locate and/or secure the liner wall insert 150 in position in the compressor rotor stage 200. In this way, it may not be necessary to join (for example braze or weld) the liner wall insert 150 to the casing 400 in order to hold it in position. Thus, a stop feature may be used to hold the insert 150 in position.
The casing 400 in the FIG. 3 example may be any sort of casing, for example a split casing 400. Such a split casing 400 may comprise at least two circumferential portions joined together at a circumferential joining location.
As shown schematically in the examples shown in FIGS. 2 and 3, the cross-sectional area of the bleed flow passage 110 may reduce along the flow path, from inlet 120 to outlet 130. In this way, the flow at the bleed flow exit 130 may be quicker (i.e. higher velocity/speed) than the flow at the bleed flow inlet 120. This may help to further control the overtip leakage flow and/or any flow structures resulting therefrom.
FIG. 4 shows a schematic cross section perpendicular to an axial direction through a partially assembled compressor rotor stage 200, as shown by way of example in FIGS. 2 and 3. The compressor rotor stage 200 is partially assembled in that not all of the liner wall inserts 100/150 are shown in position. In a complete rotor blade stage 200, liner wall inserts 100/150 would be provided around the entire circumference of the casing 300/400.
It will be appreciated that the liner wall inserts 100/150 shown in and described in relation to FIGS. 2 to 4 are by way of schematic example only. For example, the shape and/or arrangement of the bleed flow passage 110 is shown schematically only in FIGS. 2 and 3. For example, the bleed flow passage 110 may extend from the bleed flow inlet 120 and/or bleed flow outlet 130 in a direction that has a major component in the circumferential direction of the engine, which would correspond to the in-out of page direction in FIGS. 2 and 3.
By way of example, FIGS. 5 and 6 show schematic perspective views of liner wall inserts 100/150 that may correspond to those shown in FIGS. 2 and 3 respectively. FIGS. 5 and 6 both have a direction X-X shown which corresponds to the axial direction of the engine when the liner wall insert 100/150 is inserted therein.
As shown in both FIG. 5 and FIG. 6, the bleed flow passage 110 may follow a path that lies in a surface that is substantially parallel to, or has a major component parallel to, the gas washed surface 140. Purely by way of example, the path of the bleed flow passage 110 may start, from the bleed flow inlet 120, in a direction that has a major component in a direction that corresponds to a circumferential direction of the engine. The bleed flow passage 110 may then curve generally towards the axial (and upstream) direction of the engine, before curving back towards a generally circumferential direction at the bleed flow outlet 120. Thus, the bleed flow C in the bleed flow passage 110 may curve through a path that turns from generally circumferential, to generally axial, to generally circumferential (but opposite to the initial circumferential direction). The bleed flow passage 110 may be generally “C-shaped”.
An example of the bleed flow passage 110 in isolation is shown in FIG. 7. It will be understood that the bleed flow passage is shown in FIG. 7 for illustrative purposes only. Once again, FIG. 7 shows an example in which the bleed flow C enters (at the bleed flow inlet 120) and exits (at the bleed flow outlet 130) the bleed flow channel 110 in a direction that has a major component in the circumferential direction, or at least in the local circumferential-axial plane. The circumferential direction may correspond to the local rotational direction of the blades 210, labelled P in FIG. 7.
Of course, the configuration and/or direction of the bleed flow passage 110 described herein are merely an examples of various arrangements of bleed flow passages 110 within the scope of the invention.
FIG. 8 shows a circumferential portion of a liner for a circumferential rotor stage of a gas turbine engine. The liner portion shown in FIG. 8 comprises multiple liner wall inserts 100/150 connected together so as to form ring (only a part of the ring is shown in FIG. 8). Thus, each liner wall portion 110/150 may take the form of a ring segment, such as an annular segment or a segment of a frusto-cone. For example, the liner wall inserts 100/150 may be connected together, for example by brazing. The liner wall inserts 100/150 may be joined along substantially axially extending edge regions. The liner wall inserts 100/150 may be connected together at any suitable stage of manufacture, for example before or during installation in the casing 300/400 (not shown in FIG. 8 for clarity). Once again, in FIG. 8 the main flow through the compressor stage is labelled A and the bleed flow is labelled C.
Each liner wall insert 100/150 may comprise one bleed flow passage 110, as in the examples shown and described herein. However, a liner wall insert in accordance with the invention may comprise more than one bleed flow passage 110, for example 2, 3, 4 5, or more than 5 bleed flow passages. Any number of bleed flow passages may be provided in a compressor rotor stage 200. Purely by way of example, the number of bleed flow passages 110 may be the same as the number of rotor blades 210 in the stage 200.
In order to form a liner wall, or outer flow boundary, for the compressor stage 200, liner wall inserts 100/150 comprising one or more bleed flow passages 110 may be joined together. All of the liner wall inserts 100/150 may have bleed flow passages 110, as described herein by way of example. Alternatively, liner wall inserts 100/150 having at least on bleed flow passage 110 may be joined with one or more liner wall inserts 100/150 that do not have bleed flow passages in order to form the liner wall.
A liner wall insert 100/150 could be manufactured using any suitable method and/or technique. For example, a liner wall insert 100/150 (including parts thereof) could be manufactured using metal injection moulding. Furthermore, a liner wall insert 100/150 could be manufactured in any number of separate parts which may be assembled and/or joined together to form the final liner wall insert 100/150.
FIG. 9 shows an example of a liner wall insert 100 being manufactured from two parts 102, 104, prior to joining the two parts 102, 104 together. It will be appreciated that any liner wall insert, for example the liner wall inserts 100/150 described by way of example herein, could be manufactured from two parts 102, 104.
The two parts 102, 104 may be an upper (or radially outer) part 104 and a lower (or radially inner) part 102. Each of the two parts 102, 104 may comprise a part (for example one or more surfaces) of the bleed flow channel 110. In FIG. 9, for example, the lower part 102 comprises lower (or radially inner) surfaces 110 a of the bleed flow channel 110, and the upper part 104 comprises upper (or radially outer) surfaces 110 b of the bleed flow channel 110.
It will be appreciated that many alternative configurations and/or arrangements of liner wall insert 100/150, compressor rotor stage 200 and/or gas turbine engine 10 and components/parts thereof other than those described herein may fall within the scope of the invention. For example, alternative arrangements of bleed flow passage 110, such as shape and/or path, and/or components/parts thereof (such as the bleed flow inlet 120 and/or the bleed flow outlet 130) may fall within the scope of the invention and may be readily apparent to the skilled person from the disclosure provided herein. Liner wall inserts may be used in any type of gas turbine engine, for example any type of axial flow gas turbine engine, such as a turbofan (for example a two-shaft or a three-shaft turbofan engine), turboprop or turbojet gas turbine engine, for any use, such as for use in aircraft, marine applications or industrial power generation. Furthermore, any feature described and/or claimed herein may be combined with any other compatible feature described in relation to the same or another embodiment.

Claims

1. A liner wall insert for a compressor rotor stage of a gas turbine engine, the compressor rotor stage having rotor blades rotatable about an engine axis, and the liner wall insert comprising:

a gas washed surface that forms a part of an outer flow boundary inside which the rotor blades rotate during use; and

a bleed flow passage extending from a bleed flow inlet to a bleed flow outlet, the bleed flow inlet and the bleed flow outlet both being formed in the gas washed surface, wherein:

the bleed flow inlet and bleed flow outlet are arranged such that, when the liner wall insert is assembled in the compressor rotor stage, the bleed flow inlet is axially downstream of the bleed flow outlet.

2. A liner wall insert according to claim 1, wherein the cross sectional area of the bleed flow passage is smaller at the bleed flow outlet than at the bleed flow inlet.

3. A liner wall insert according to claim 1, wherein the gas washed surface of the liner wall insert is a segment of a cylinder or a segment of a frusto-conical shape.

4. A liner wall insert according to claim 1, wherein the bleed flow passage follows a path that is substantially parallel to the gas washed surface.

5. A method of manufacturing a liner wall insert according to claim 1, comprising:

manufacturing a radially inner portion of the liner wall insert;

manufacturing a radially outer portion of the liner wall insert; and

joining the radially inner portion and the radially outer portion together, wherein:

the radially inner portion comprises first surfaces that form part of the bleed flow passage;

the radially outer portion comprises second surfaces that form part of the bleed flow passage; and

the first surfaces and the second surfaces come together to form the bleed flow passage when the radially inner portion and radially outer portion are joined together.

6. A method of manufacturing a liner wall insert according to claim 1, wherein:

the method comprises metal injection moulding to manufacture at least a part of the insert.

7. A method of manufacturing a liner wall insert according to claim 5, wherein:

8. A compressor rotor stage comprising:

a plurality of rotor blades rotatable about an engine axis, the rotor blades extending to a radially outer tip; and

a plurality of liner wall inserts according to claim 1, wherein

the plurality of liner wall inserts are attached together so as to define an outer flow boundary inside which the tips of the rotor blades pass during use.

9. A compressor rotor stage according to claim 8, wherein:

the axial location of the bleed flow inlet of each liner wall insert corresponds to a trailing edge region of the rotor blades; and

the axial location of the bleed flow outlet of each liner wall insert corresponds to a leading edge region of the rotor blades.

10. A compressor rotor stage according to claim 8, wherein the axial location of each bleed flow outlet is upstream of the axial location of the leading edge of the rotor blades.

11. A compressor rotor stage according to claim 8, wherein the plurality of liner wall inserts are brazed together to form a ring.

12. A compressor rotor stage according to claim 8, further comprising a casing radially outboard of the liner wall inserts, wherein:

the casing and the liner wall inserts have cooperating location features that are engaged so as to hold the liner wall inserts in position.

13. A compressor rotor stage according to claim 8, further comprising a casing radially outboard of the liner wall inserts, wherein:

the liner wall inserts are joined to the casing so as to be held in position.

14. A gas turbine engine comprising at least one compressor rotor stage according to claim 8.

15. A method of improving the operation of a compressor rotor stage of a gas turbine engine, the compressor rotor stage comprising rotor blades (210) extending from a root to a tip and being rotatable about an engine axis within an outer flow boundary, the method comprising:

bleeding flow from the compressor flow stream into a bleed flow inlet, through a bleed flow passage, and back out into the flow stream from a bleed flow outlet that is axially upstream of the bleed flow inlet, wherein:

the outer flow boundary is formed by a plurality of liner wall inserts joined together to circumferentially surround the rotor blades; and

both the bleed flow inlet and the bleed flow outlet are formed in the outer flow boundary.

16. A method according to claim 15, wherein a plurality of the liner wall inserts have a bleed flow passage formed therein.