GB2382380A

GB2382380A - A removable abradable lining for the casing assembly of a gas turbine engine

Info

Publication number: GB2382380A
Application number: GB0128258A
Authority: GB
Inventors: John Michael Scott
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2001-11-24
Filing date: 2001-11-24
Publication date: 2003-05-28
Also published as: GB0128258D0

Abstract

A casing assembly for a gas turbine engine includes an abradable lining arrangement 46 including a mounting member 49 provided with an abradable material 60. The mounting member 49 is removably attached to the inside surface of the generally cylindrical turbine casing for easy replacement when worn. The casing may have a plurality of channel regions in which the mounting member may be situated, the abradable lining comprising semi-circular segments. The lining may also include a plurality of parts which may be detached from one another. The casing may be of unitary construction or a plurality of adjacent axially spaced casing elements.

Description

CASING ASSEMBLY FOR A GAS TURBINE ENGINE This invention relates to a casing assembly for a gas turbine engine, and to an abradable lining arrangement for a casing assembly for a gas turbine engine. The invention is particularly applicable to the casings for compressors of gas turbine engines.

Gas turbine engines comprise, in series, a fan (or low pressure compressor), one or more higher pressure compressors, a combustion chamber for burning the compressed air from the compressor with fuel and one or more turbines driven by the exhaust gases from the combustor. These components of the gas turbine engine are contained within an annular casing.

The compressors of gas turbine engines are commonly axial flow compressors which comprise alternate rows of rotating (rotor) blades and stationary (stator) vanes to accelerate and diffuse the air passing therethrough until a required pressure rise is obtained. The casing surrounding the compressor normally comprises a number of cylindrical casing elements bolted together. It is of importance in compressor design that the clearance between the tips of the compressor rotor blades and the inner surface of the cylindrical casing is accurately maintained to a predetermined clearance level, to ensure the highest possible efficiency and minimum fluid flow leakage over the tips of the blades.

However, during operation of the engine, the rotor blades thermally expand, and vibration or movement of the rotating shaft upon which the rotor blades are mounted may cause corresponding radial movement of the blades. As

such, the rotor blades may rub the inner surface of the casing which abrades the tips of the rotors thus increasing the gap between the blade tips and the casing.

It is known to provide the inner surface of the compressor casing with an abradable lining or"rub strip" which sacrificially wears away when rubbed by the tips of the rotor blades. Once worn away, this abradable lining may be replaced by removing the worn liner, blasting clean the inner surface of the annular compressor casing and respraying further abradable material onto the inner surface of the compressor casing. However, this necessitates dismantling the engine to gain access to the inner surface of the casing, and taking the compressor casing to a blasting and spraying area. Such a refurbishment operation is costly and time consuming, particularly as the spraying equipment is very expensive.

According to the invention, there is provided a gas turbine engine casing assembly for housing a rotor assembly including a plurality of radially extending rotor blades, the casing assembly including a casing and an abradable lining arrangement which is removably mountable on the casing, for surrounding the tips of the rotor blades in use.

Preferably the casing is substantially cylindrical, and includes a substantially annular channel formed in its inner surface, the abradable lining arrangement being removably mountable within the channel.

Preferably the casing includes a plurality of axially spaced substantially annular channels formed in its inner surface, each channel mounting an abradable lining arrangement, the respective lining arrangements surrounding adjacent axially spaced rotor assemblies in use.

The casing may comprise a substantially unitary member including a plurality of axially spaced channels formed in its inner surface.

Alternatively, the casing may comprise a plurality of adjacent axially spaced casing elements. In this embodiment, each channel may be formed substantially in one casing element but a side wall of the channel may be defined by an adjacent casing element.

Preferably the abradable lining arrangement comprises a substantially annular mounting member and an abradable material provided on the mounting member. Preferably the mounting member includes a generally cylindrical base which contacts a base of the channel in the casing and spaced side walls, the base and the side walls together defining a generally annular recess for receiving the abradable material.

The lining arrangement may comprise a plurality of parts which may be detached from one another. The lining arrangement may comprise two semi-circular segments.

According to the invention there is further provided an abradable lining arrangement for a casing of a gas turbine engine, the lining arrangement comprising a mounting member and an abradable material provided on the mounting member, the mounting member being removably mountable on an inner surface of the casing of the gas turbine engine.

An embodiment of the invention will now be described by way of example only, with reference to the accompanying drawings, in which; Fig. 1 is a cross-sectional view of half of a gas turbine engine; Fig. 2 is a diagrammatic perspective view of a

compressor casing assembly according to a first embodiment of the present invention, in-situ; Fig. 3 is a diagrammatic cross-sectional side view of a compressor of a gas turbine engine incorporating the casing assembly of Fig. 2; Fig. 4 is an enlarged and more detailed view of part of the assembly shown in Fig. 3; Fig. 5 is an enlarged diagrammatic view of an abradable lining arrangement of the casing assembly shown in Figs. 2 to 4; Fig. 6 is a diagrammatic end view of the casing assembly of Fig. 2 to 4; and Fig. 7 is a diagrammatic sectional side view of a part of a casing assembly according to a second embodiment of the invention.

Referring to Fig. 1, a gas turbine engine generally indicated at 10 comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high pressure compressor 14, a combustor 15, a turbine arrangement comprising a high pressure turbine 16, an intermediate pressure turbine 17 and a low pressure turbine 18, and an exhaust nozzle 19.

The gas turbine engine 10 operates in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 which produces two air flows, a first air flow into the intermediate pressure compressor 13 and a second air flow which provides propulsive thrust. The intermediate pressure compressor compresses the air flow directed into it before delivering that air to the high pressure compressor 14 where further compression take place.

The compressed air exhausted from the high pressure

compressor 14 is directed into the combustor 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 and 18, before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low pressure turbines 16,17 and 18 respectively drive the high and intermediate pressure compressor 14 and 13 and the fan 12 by suitable interconnecting shafts.

The intermediate pressure compressor 13 is connected to the intermediate pressure turbine 17 via an interconnecting shaft 26. Similarly the high pressure compressor 14 is connected to the high pressure turbine 16 by an interconnecting shaft 28 and the fan 12 is connected to the low pressure turbine 18 via an interconnecting shaft 30. These drive shafts 26,28 and 30 are concentrically mounted around a central axis X-X this also being the central axis of the generally annular gas turbine engine.

The high and intermediate pressure compressors 14 and 13 comprise alternate axially spaced arrays of rotor blades 20 and stator vanes 22 (the respective arrays being referred to as stages of the compressor). These can be seen more clearly by reference to Fig. 3.

The rotor blades 20 of each stage of the compressors 13 and 14 are arranged in a circumferential array with each blade extending radially outwardly in relation to the axis X-X. The compressors 13 and 14 (together with the combustor 15 and the turbines 16,17 and 18) are encased within an annular casing structure 24. The tips 38 of the rotor blades 20 are positioned close to an inner surface 32 of the casing 24, such that there is a minimum clearance

therebetween.

Referring to Figs. 2 to 6, in a first embodiment of the invention, a unitary casing structure 24A surrounds a number of stages of rotor blades 20 and stator vanes 22. It may be seen that the casing structure 24A is provided with a plurality of annular channels 44 recessed into its inner surface 32. Each channel includes a generally cylindrical base 39 and spaced side walls 40 and 42. Each channel is aligned with one stage of rotor blades, the channel surrounding the stage of rotor blades and extending for at least its full axial length.

Each channel 44 houses an abradable lining arrangement 46. Referring in particular to Figs. 2,3 and 5, the abradable lining arrangement 46 includes a mounting member 49 and an abradable lining material 60 housed within the mounting member. In Fig. 2, the mounting member 49 is cut away to show the material 60 more clearly. The abradable material 60 may be any of a range of commercially available, thermally sprayable abradable materials, such as aluminium and silicon based materials or nickel and bentonite based materials.

The abradable lining arrangement 46 is generally annular in overall shape, and is of complementary shape to the channel 44. Thus the abradable lining arrangement surrounds the rotor blades 20 of the respective compressor stage in use. However, in the embodiment of Figs. 2 to 6, the abradable lining arrangement comprises two half sections which together form the annular whole (see Fig.

6).

The mounting member 49 includes a generally cylindrical base 48, of complementary shape to the base 39 of the respective channel 44. The mounting member further

includes spaced apart side walls 50 and 52 (see Fig. 3).

The base 48 and walls 50 and 52 together define a recess which is filled with the abradable lining material 60. In the embodiment of Figs 2 to 6, the mounting member 49, together with the abradable lining material, is split into two halves.

The abradable lining arrangement 46 may be located within its respective channel, the division of the arrangement into two separate halves making such location possible. Once in place, the lining arrangement tends not to move, but it may be held in place temporarily during assembly by spot welding for example.

An abradable lining arrangement 46 is positioned radially outwardly of the tips 38 of each stage of rotor blade 20, in a respective channel 44. A small radial clearance 62 is provided between the tips 38 of the rotor blades 20 and the lining material 60. It is this radial clearance 62 which is minimized to promote the efficiency of the engine.

Referring to Fig. 7, an alternative casing structure 24 comprises a series of axially spaced, generally cylindrical casing elements 24a to 24d, each casing element surrounding a respective stage of the compressor. In this embodiment, the casing structure 24 again includes a plurality of annular channels 44. Each channel 44 is recessed into the casing structure 24, as in the previous embodiment. However, in this embodiment, the base 39 and one side wall 40 of the recess are formed within one casing element and the other side wall 42 is defined by an edge of an adjacent casing element. In this embodiment the abradable lining arrangements each comprise a single 3600 annular member. As in the previous embodiment, the lining

arrangement includes a mounting member 49 having a cylindrical base 48 and side walls 50 and 52. In this embodiment, the mounting member is unitary. The mounting member 49 houses an abradable lining material 60 as in the previous embodiment. In this embodiment, the casing assembly may be formed by sliding each abradable lining arrangement axially into its channel 44, before bolting the adjacent casing element in place.

In use, the rotor blades 20 of the gas turbine engine rotate and at least some of the blade tips 38 contact the abradable lining material 60 during the running of the engine due to, for example, slight radial movement of the compressor drive shaft on which the blades 20 are mounted or radial expansion of the rotor blades 20 due to temperature changes.

When parts of the abradable lining 60 are required to be replaced, the appropriate abradable lining arrangement 46 (including the mounting 49 and abradable material 60) is removed and replaced with a new lining arrangement. This enables the abradable lining material to be replaced quickly and efficiently without the need to strip down the engine and send the casing to be blasted clean.

There is thus provided a casing assembly in which worn out abradable lining material may be easily replaced.

Various modifications may be made to the above described embodiments without departing from the scope of the invention. For example the abradable lining arrangement could be split into more than two parts for ease of assembly. The shape of the arrangement may be modified to fit alternative casing structures.

Whilst endeavouring in the foregoing specification to draw attention to those features of the invention believed

to be of particular importance it should be understood that the Applicant claims protection in respect of any patentable feature or combination of features hereinbefore referred to and/or shown in the drawings whether or not particular emphasis has been placed thereon.

Claims

1. A gas turbine engine casing assembly for housing a rotor assembly including a plurality of radially extending rotor blades, the casing assembly including a casing and an abradable lining arrangement which is removably mountable on the casing, for surrounding the tips of the rotor blades in use.

2. A casing assembly according to Claim 1, wherein the casing is substantially cylindrical, and includes a substantially annular channel formed in its inner surface, the abradable lining arrangement being removably mountable within the channel.

3. A casing assembly according to Claim 2, wherein the casing includes a plurality of axially spaced substantially annular channels formed in its inner surface, each channel mounting an abradable lining arrangement, the respective lining arrangements surrounding adjacent axially spaced rotor assemblies in use.

4. A casing assembly according to claim 3, the casing comprising a substantially unitary member including a plurality of axially spaced channels formed in its inner surface.

5. A casing assembly according to claim 3, the casing comprising a plurality of adjacent axially spaced casing elements.

6. A casing assembly according to claim 5, wherein each channel is formed substantially in one casing element but wherein a side wall of the channel is defined by an adjacent casing element.

7. A casing assembly according to any preceding claim,

wherein the abradable lining arrangement comprises a substantially annular mounting member and an abradable material provided on the mounting member.

8. A casing assembly according to claim 7, wherein the mounting member includes a generally cylindrical base which contacts a base of the channel in the casing and spaced side walls, the base and the side walls together defining a generally annular recess for receiving the abradable material.

9. A casing assembly according to claim 7 or claim 8, the lining arrangement comprising a plurality of parts which may be detached from one another.

10. A casing assembly according to claim 9, wherein the lining arrangement comprises two semi-circular segments.

11. An abradable lining arrangement for a casing of a gas turbine engine, the lining arrangement comprising a mounting member and an abradable material provided on the mounting member, the mounting member being removably mountable on an inner surface of the casing of the gas turbine engine.

12. A casing assembly substantially as described herein with reference to any of Figs. 2 to 7 of the accompanying drawings.

13. An abradable lining arrangement substantially as described herein with reference to any of Figs. 2 to 7 of the accompany drawings.

14. Any novel subject matter or combination including novel subject matter disclosed herein, whether or not within the scope of or relating to the same invention as any of the preceding claims.