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 
       [0001]    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 
       [0002]    1. Field of the Invention 
         [0003]    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. 
         [0004]    2. Description of the Related Art 
         [0005]    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. 
         [0006]    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 
       [0007]    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. 
         [0008]    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. 
         [0009]    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. 
         [0010]    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. 
         [0011]    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. 
         [0012]    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. 
         [0013]    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. 
         [0014]    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. 
         [0015]    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. 
         [0016]    Aspects of the invention may comprise any one or more of the following features. 
         [0017]    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. 
         [0018]    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. 
         [0019]    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. 
         [0020]    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. 
         [0021]    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. 
         [0022]    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. 
         [0023]    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%. 
         [0024]    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%. 
         [0025]    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. 
         [0026]    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. 
         [0027]    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. 
         [0028]    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. 
         [0029]    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. 
         [0030]    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. 
         [0031]    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. 
         [0032]    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 
         [0033]    Embodiments of the invention will now be described by way of example only, with reference to the accompanying diagrammatic drawings, in which: 
           [0034]      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; 
           [0035]      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; 
           [0036]      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; 
           [0037]      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; 
           [0038]      FIG. 5  is a perspective view of an example of a liner wall insert in accordance with the invention; 
           [0039]      FIG. 6  is a perspective view of an alternative liner wall insert in accordance with the invention; 
           [0040]      FIG. 7  is a perspective view of a bleed flow channel; 
           [0041]      FIG. 8  is a perspective view of a number of liner wall inserts joined together, forming a part of a ring; and 
           [0042]      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 
       [0043]    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 . 
         [0044]    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. 
         [0045]    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. 
         [0046]    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 . 
         [0047]    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 . 
         [0048]    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 . 
         [0049]    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 . 
         [0050]    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. 
         [0051]    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 . 
         [0052]    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. 
         [0053]      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 . 
         [0054]    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 . 
         [0055]    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. 
         [0056]    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. 
         [0057]    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. 
         [0058]      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 . 
         [0059]    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 . 
         [0060]    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. 
         [0061]    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”. 
         [0062]    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 . 
         [0063]    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. 
         [0064]      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. 
         [0065]    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 . 
         [0066]    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. 
         [0067]    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 . 
         [0068]      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 . 
         [0069]    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 . 
         [0070]    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.