GB2419638A

GB2419638A - Compressor casing with an abradable lining and surge control grooves

Info

Publication number: GB2419638A
Application number: GB0423714A
Authority: GB
Inventors: Stuart Ellis
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2004-10-26
Filing date: 2004-10-26
Publication date: 2006-05-03
Also published as: GB0423714D0

Abstract

A compressor casing 10 for a gas turbine engine includes an abradable lining 16 which extends into a recess 20 in the surface of the casing 10 to aid retention. Circumferential surge control grooves 18 may be located either side of the recess 20, and arrangements including various numbers of grooves 18 and recesses 20 may be provided (see figures 8 and 9). The recess 20 may have a variety of cross-sectional shapes (see figures 3-7), and the abradable lining may be made from material which includes metallic or ceramic powders. The casing 10 may be produced by firstly forming a recess 20, then applying a coating of abrading material 16 and then forming grooves 18 either side of the recess 20 through the abrading material 16.

Description

Compressor Casing for a Gas Turbine Engine This invention relates to a

compressor casing for a gas turbine engine, a gas turbine engine including such a compressor casing, and a method of forming a compressor casing for a gas turbine engine.

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 a 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) blades to accelerate and diffuse the air passing therethrough until a required pressure rise is obtained. The blades and stators are contained within a cylindrical casing. 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 minimal fluid flow leakage over the tips of the blades.

However, during operation of the engine, the rotor blades thermally expand, and vibrational 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 compressor casing which abrades the tips of the rotors, thus increasing the gaps between the blade tips and the casing.

It is known to provide an abradable liner on the casing which in use will abrade if impacted by the rotor, without causing significant damage to the blade tips. It is further known to use abrasive liners which abrade a rotor, if the rotor impacts the lining. Such liners can be used with nickel blades and in high temperature applications it is also known to provide circumferential grooves to improve low speed stability and surge margin of compressors.

It is however difficult to combine such grooves with abradable or abrasive liners. If the liner is applied before the grooves are cut, the liner is prone to damage.

If the liner is applied after the groove has been cut it will tend to block the grooves. Also, the liner will tend not to adhere securely because it is applied in relatively narrow strips between the grooves.

When used in this specification the term "abrading" liner or material is to be understood as designating either an abradable or abrasive liner or material, as outlined above.

According to the present invention there is provided a compressor casing for a gas turbine engine, the casing including an abrading lining on the casing face, characterised in that a circumferential recess is provided in the casing face, and the abrading lining extends into the recess to aid retention of the lining on the casing.

At least two circumferential grooves preferably extend into the face of the casing, with a circumferential recess in the casing provided between the grooves. The grooves preferably extend into the casing for a distance beyond the abrading lining.

More than two circumferential grooves may be provided, with a recess in the casing provided in each space between adjacent grooves. A plurality of recesses may be provided in each space between adjacent grooves.

The circumferential recess or recesses may have a rectangular, triangular, semicircular or rhomboidal cross section.

The circumferential recess or recesses may extend at least as deep as the circumferential grooves.

The abrading material may include a metallic powder, which may be hardened and sintered. The abrading material may include a ceramic powder.

The invention also provides a gas turbine engine, characterjsed in that the engine includes a compressor casing according to any of the preceding six paragraphs.

The invention further provides a method of forming a compressor casing for a gas turbine engine, the method including forming a circumferential recess in the casing and applying a coating of abrading material onto the casing and to locate in the recess, forming two grooves in the casing through the abrading material, one on each side of the circumferential recess.

A plurality of recesses and a plurality of grooves may be formed in the casing, with one or more recesses between adjacent grooves.

Embodiments of the present invention will now be described by way of example only, and with reference to the accompanying drawings, in which:Fig. 1 is a sectional side view of the upper half of a gas turbine engine; Fig. 2 is a diagrammatic cross sectional view through part of a compressor with a casing according to the invention; Figs. 3 - 7 are diagrammatic cross sectional views through part of further compressor casings according to the invention; Fig. 8 is a diagrammatic sectional view through part of a further compressor casing according to the invention; and Fig 9 is a similar view to Fig 8 of a still further compressor casing according to the invention.

Referring to Fig. 1, a gas turbine engine is generally indicated at 101 and comprises, in axial flow series, an air intake 111, a propulsive fan 121, an intermediate pressure compressor 131, a high pressure compressor 141, a combustor 151, a turbine arrangement comprising a high pressure turbine 161, an intermediate pressure turbine 171 and a low pressure turbine 181, and an exhaust nozzle 191.

The gas turbine engine 101 operates in a conventional manner so that air entering the intake 111 is accelerated by the fan 121 which produce two air flows; a first air flow into the intermediate pressure compressor 131 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 141 where further compression takes place.

The compressed air exhausted from the high pressure compressor 141 is directed into the combustor 151 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 161, 171 and 181 before being exhausted through the nozzle 191 to provide additional propulsive trust. The high, intermediate and low pressure turbines 161, 171 and 181 respectively drive the high and intermediate pressure compressors 141 and 131 and the fan 121 by suitable interconnecting shafts.

Fig. 2 shows part of a compressor of a gas turbine engine with a casing 10 and a rotatable rotor 12. The casing 10 could be provided on the intermediate compressor 131 or the high pressure compressor 141 as shown respectively at 132 and 142 in Fig. 1. A small gap 14 is provided between the rotor 12 and casing 10. The size of the gap 14 will vary with time for various reasons including variations in temperature of the rotor 12 and other components and also vibration or movement of the components. An abradable lining 16 is provided on the casing 10. The lining 16 may include a metallic powder which may be hardened and sintered. Alternatively or in addition the abradable material may include other materials such as ceramic powders. If the rotor 12 contacts the casing 10 this will generally cause the lining 16 to abrade, without substantially damaging the tip of the rotor 12.

Two circumferential grooves 18 are provided in the casing 10 facing the rotor 12. Between the grooves 18 a recess 20 is provided in the rotor lining of a similar depth to the grooves 18 but which is thicker thereto. The recess 20 is filled with the abrasive lining material.

The casing 10 will be formed by forming the recess 20 and then applying the abrasive lining thereto such that the abrasive material fills the recess 20. The grooves 18 will then be cut into the casing 10 through the lining 16.

This arrangement permits the combination of circumferential grooves which provide improved low speed stability and surge margin with an abradable lining. The recess between the grooves provides for a secure bonding of the lining on the casing and prevents only a small area of abrasive material being mounted thereto.

Figs. 3 - 5 show alternative shape recesses 22, 24, 26. The recess 22 of Fig. 3 is of the same depth and width as the abradable lining 28. Fig. 4 shows a recess 24 again of the same depth as the abradable lining 30, but which is significantly wider, and of a similar width to the recess 20. Fig. 5 shows a recess 26 of similar dimensions to the recess 24 except that the recess 26 is rhomboidal, which may make for easier filling thereof with abradable material.

Figs 6 and 7 show further alternative shape recesses 32, 34. Fig 6 shows a pair of triangular recesses 32, whilst Fig 7 shows a semicircular recess 34. Again the shapes of the recesses 32, 34 may facilitate filling thereof with abradable material.

Fig. 8 shows a compressor casing in which four circumferential grooves 36 have been formed. A recess 38 has been provided between each of the grooves 36 and on the outside of the endmost ones thereof. The recesses 38 are filled with abradable material 40 which also provides a lining across the whole of the inside of the casing aside from the grooves 36.

Fig 9 shows a further compressor casing with a pair of circumferential grooves 42. On the outer side of each groove 42 a recess 44 is provided, whilst a pair of spaced recesses 46 are provided between the grooves 42. In this instance the recesses 44, 46 are filled with an abrasive material which will abrade a rotor, if the latter impacts thereagainst.

It is to be realised that any combination of the above features by way of numbers, shapes and the location of recesses and grooves, can be included as is appropriate for particular circumstances. As is indicated the invention is appropriate for abradable and also abrasive liners.

There are thus described advantageous casings which can be readily and robustly manufactured. Casings with a recess for the abradable or abrasive material can readily be formed to provide inexpensive and advantageous performance. Recesses may be provided in plain casings where there are no grooves, and the recesses as shown in any of the drawings could be used in plain casings. It is also to be realised that at overhaul or rework, it may be required to reapply the respective lining material. It may be necessary to locate blanking rings or something else in the grooves to prevent lining material entering thereinto.

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 compressor casing (10) for a gas turbine engine, the casing (10) including an abrading lining (16,28,30,40,48) on the casing face, characterjsed in that a circumferential recess (20,22,24,26,38,44,46) is provided in the casing face, and the abrading lining (16,28,30,40,48) extends into the recess (20,22,24,26,38,44,46) to aid retention of the lining on the casing (10)

2. A casing according to claim 1, characterised in that at least two circumferential grooves (18,36,42) extend into the face of the casing (10) , with a circumferential recess (20,22,24,26,38,44,46) in the casing (10) provided between the grooves.

3. A casing according to claim 2, characterised in that the grooves (18, 36,42) extend into the casing (10) for a distance beyond the abrading lining (16,28,30,40,48).

4. A casing according to claims 2 or 3, characterised in that more than two circumferential grooves (18,36,42) are provided, with a recess (20,22, 24,26,38,44,46) in the casing provided in each space between adjacent grooves (18,36, 42)

5. A casing according to claim 4, characterised in that a plurality of recesses (20,22,24,26,38,44,46) are provided in each space between adjacent grooves (18,36,42)

6. A casing according to any of claims 2 to 5, characterjsed in that the circumferential recess or recesses (20,22,24,26,38,44,46) extend at least as deep as the circumferential grooves (18,36,42).

7. A casing according to any of the preceding claims, characterised in that circumferential recess or recesses (20,22,24,26,38,44,46) have a rectangular, triangular, semicircular or rhomboidal cross section.

8. A casing according to any of the preceding claims, characterised in that the abrading material includes a metallic powder, which may be hardened and sintered.

9. A casing according to any of the preceding claims, characterised in that the abrading material includes a ceramic powder.

10. A gas turbine engine, characterised in that the engine (101) includes a compressor casing (10) according to any of the preceding claims.

11. A method of forming a compressor casing (10) for a gas turbine engine, the method including forming a circumferential recess (20,22,24,26,38,44, 46) in the casing (10) and applying a coating (16,28,30,38) of abrading material onto the casing (10) and to locate in the recess (20,22,24,26,38,44,46), forming two grooves (18,36,42) in the casing (10) through the abrading material (16,28,30,40,48), one on each side of the circumferential recess (20,22, 24,26,38,44,46).

12. A method according to claim 11, characterised in that a plurality of recesses (20,22,24,26,38,44,46) and a plurality of grooves (18,36,42) are formed in the casing (10), with one or more recesses (20,22,24,26,38,44,46) within adjacent grooves (18,36,42) being formed on each side of each recess (20,22,24,26, 38,44,46)

13. A compressor casing substantially as hereinbefore described and with reference to the accompanying drawings.

14. A gas turbine engine substantially as hereinbefore described and with reference to the accompanying drawings.

15. A method of forming a compressor casing for a gas turbine engine, the method being substantially as hereinbefore described and with reference to the accompanying drawings.