GB2262140A - Gas turbine engine sealing assembly - Google Patents

Abstract

Gas turbine engine sealing assembly comprises seal segments 64, 66 attached to support discs 72, 74 by serrated roots on the segments engaging serrations on the support discs 72, 74. The segments 64, 66 each have labyrinth seals 56, 58 which cooperate with platforms 30A, 32A on stator blades 30, 32. The support discs 72, 74 may be welded to the adjacent turbine rotors 38, 42, 46. The sealing assembly is particularly applicable to heavy-weight industrial gas turbine gas generators, and avoids the need for the labyrinth seals to be supported from the turbine rotor blade platforms. <IMAGE>

Description

GAS TURBINE ENGINE SEALING ASSEMBLY This invention relates a gas turbine engine sealing assembly for the sealing the gas flow path in a turbine of a gas turbine engine. In particular the invention relates to sealing the gas flow path in an industrial gas turbine engine.

Industrial gas turbine engines generally comprise a gas generator consisting of a compressor, a combustor in which fuel and air are mixed and burnt, a turbine which is driven by the products of combustion and which drives the compressor, and a power turbine driven by the high temperature, high velocity gases from the gas generator. The power turbine is arranged to drive a load, such as an electricity generator, or a pump for pumping oil or gas.

Heavyweight industrial gas generators are bulky and there can be large distances between the bearings of a shaft on which the compressor and turbine are mounted. The turbine of the gas generator will comprise one or more stages of blades, each stage comprising an array of rotor blades mounted on the gas generator rotor, and an array of stator blades mounted from a casing of the gas generator. The high temperature, high velocity gases flow through an annular passage in which the rotor and stator blades are disposed. The radially inner boundary of the passage is partially defined by platforms on the inner ends of the stator blades. To restrict leakage of the turbine gases around the ends of the stators between their inner platforms and the rotor, the platforms are usually sealingly engaged by sealing elements secured to the rotor.

The relatively large distances between the rotor bearings, for example up to nine metres, result in large relative axial movements between the rotor and the gas generator casing due to differential thermal expansion between the rotor and the casing. Thus, any seal components providing a seal between the rotating and static components of the gas generator turbine must be able to cope with such movements.

In the case of relatively low power engines, rotating sealing elements, against which the stator inner platforms run, can be achieved by casting axially extending projections or 'wings' onto the inner platforms of the rotor blades. These projections are provided on the rotor blades of adjacent stages and extend towards each other over the intevening gap so that their confronting edges abut one another.

On larger engines, these wings become so long that the bending stresses on them are excessive. Also, when the rotor blades are cast by directional solidification techniques, the material properties of the wings are not appropriate.

The present invention seeks to provide a form of annular sealing construction which avoids the need for wings on the rotor blade platforms to provide a seal, whilst maintaining a seal along the inner boundary of the gas flow annulus.

Accordingly, the present invention provides a rotating seal assembly for an axial flow gas turbine engine, the gas turbine engine comprising rotating structure and static structure surrounding the rotating structure, the static structure forming a static circumferential sealing surface, the seal assembly comprising a plurality of sealing segments having radially inner ends secured to a seal support rotor disc and radially outer sealing surfaces which form in combination a rotating circumferential sealing surface arranged to co-operate with the static circumferential sealing surface.

The seal support disc can be located between adjacent rotor blade support discs supporting corresponding rotor blade stages of the gas turbine engine, and the rotating sealing surface of the seal assembly can cooperate with a static sealing surface comprising a plurality of stator blade inner platforms. The seal support disc can be welded to the adjacent turbine rotors.

The seal segments can be attached to the seal support disc by means of serrated roots engaging with correspondingly shaped serrations in the seal support disc.

Where two or more seal assemblies are present in axial succession, with a blade support disc interposed between adjacent seal support discs, the serrations in the respective support discs can all be cut on a common diameter to facilitate assembly.

The seal segments can be machined from a typical rotor blade superalloy casting and include labyrinth seals to sealingly engage the platforms on the inner ends of the stator blades.

Arcuate seal strips can be used to seal between circumferentially adjacent edges of the seal segments, and between the seal segments and adjacent platforms provided on the rotor blades.

Exemplary embodiments of the present invention will now be more particularly described with reference to the accompanying drawings in which; Fig. 1 shows diagrammatically an industrial gas turbine engine; Fig. 2 is a more detailed cut-away view of area II in Fig.l, showing part of a gas generator turbine incorporating a known type of gas flow path sealing construction; and Fig. 3 shows a similar view of part of a larger gas generator turbine incorporating one form of gas flow path sealing construction according to the present invention.

Referring to the drawings, in Fig. 1 there is shown an industrial gas turbine power plant 10 comprising a gas generator 12 and a power turbine 14 arranged to drive a load 16, which can be, for example, an electricity generator or a pump. The gas generator 12 comprises, in axial flow series, a compressor 18, a combustor 20, and a turbine 22 mounted on a common shaft with the compressor.

High temperature, high velocity gas produced in the gas generator 12 by the compressor 18 and the combustor 20 drives the turbine 22, which drives the compressor 18 through the common shaft. The excess power in the turbine gases after passage through the turbine 22 is used to drive the power turbine 14.

Referring to Fig. 2, there is a detail of part of a known turbine 22 of a gas generator.

The static structure of the turbine 22 comprises an outer casing 24 to which are attached, via a support ring 24A, stator vanes stages 26 and 28 comprising stator vanes 30 and 32. An array of nozzle guide vanes 34 is secured between a further support ring 24B, also attached to casing 24, and a radially inner static support structure 36. The stator vanes 30 and 32, and the nozzle guide vanes 34, all have inner and outer platforms 30A, 30B, 32A, 32B and 34A, 34B respectively.

The rotating structure of the turbine 22 includes a first stage rotor disc 38, having rotor blades 40 located axially between the nozzle guide vanes 34 and the stator vanes 30, a second stage rotor disc 42 having rotor blades 44 located between the stator vanes 30 and 32, and a third stage rotor disc 46 having rotor blades 48 located downstream of the stator vanes 32. The rotor blades 40,44,48 all have inner platforms 40A,44A,48A, respectively. The outer tips of the first stage rotor blades 40 cooperate with a static sealing ring 50 held in support ring 24B, but the outer ends of the rotor blades 44 and 48 have shrouds 44B and 48B with projections which sealingly cooperate with abradeable surfaces 52 and 54 on circumferential lands of the support ring 24A.

The products of combustion flow through the gas flow path annulus 60 from the combustor 20 and between the nozzle guide vanes 34 in the direction of arrow A. The radially inner boundary of the gas flow path annulus 60 is defined by the inner platforms of the stator and rotor blades and also by wing seals 59B,61A,61B,62A and 62B.

Adequate sealing of the gas flow path is acheived on its inner boundary by labyrinth seals comprising circumferentially extending sealing fins 56A,56B,58A,58B and 59 on wing seals 61A,61B,62A,62B and 53S, respectively, which cooperate with the stator platforms 30A,32A and 34A. Wing seals 61A,62A extend rearwardly from the roots 40C,44C of rotor blades 40 and 44 respectively, while wing seals 59B,61B,62B extend forwardly from the roots 40C,44C and 48C of rotor blades 40, 44 and 48 respectively. Wing seals 61A,61B and 62A,62B therefore extend towards each other and their confronting edges define small axial gaps 61C,62C, to allow for thermal expansion. Circumferentially spaced webs 61D/E, 62D/E provide support to the longer wing seals 61A/B,62A/B against the effect of centrifugal forces. However, such support is not needed for the shorter, less massive wing seal 59B. The wing seals and their support webs are cast integrally with the blade roots 40C, 44C and 48C, and of course have the same circumferential extent as the blade platforms of which they form axially extending continuations.

It will be appreciated that as the engine size increases, the spacing between turbine rotors will increase, and so will the diameter of the rotors. Thus the wings 61,62 will tend to increase in length and be located at larger radii, while their support webs must increase in number and thickness to cope with the centrifugal working loads, which increase as the product of mass, radius and the square of angular velocity. Eventually, having regard to the working loads experienced by the wing seals and imposed by the wing seals on the blade roots 40C,44C,48C and on the rotor discs 38,42,46, the strength of available materials and the manufacturing methods available will limit the length of the wings and their diameters to those which will maintain adequate sealing and/or impose acceptable stresses on the blade roots and discs.

A further problem arises even for small size engines, in that while blades cast by directional solidification techniques are to be preferred for use because of their superior strength and temperature resistance, such casting techniques cannot be used for blades with integral wing seals because the extent of the wing seal lies in the a different direction from the desired radial metallurgical orientation in the body of the blade.

A solution to these problems, as embodied in an engine of larger size than in Fig. 2, is illustrated in Fig. 3, in which larger versions of similar smaller components already described with reference to Fig. 2 have been allotted similar references.

In order to seal the inner boundary of the gas path annulus 60, sealing fins 56 and 58, which co-operate with abradeable sealing lands 30C and 32C on stator vane platforms 30A and 32A, are provided on seal segment assemblies 64 and 66 respectively. The assemblies 64 and 66 are made up from a number of arcuate segments 68 and 70, having platforms 68A, 70A whose axially extending edges confront each other or abut to form a complete circumferential sealing surface. The platforms are supported from the bodies of the segments 68,70 by webs 68C/D and 70C/D.

With respect to manufacture and assembly, the seal segments 68, 70 can be cast to near net shape and finish machined similarly to a typical rotor blade superalloy casting. Each segment 68, 70 is machined with serrated root portions 68B, 70B, indicated by broken lines. The root portions are assembled onto support discs 72, 74 formed with corresponding serrations on a common diameter D. The seal support discs 72,74 can be machined from forgings like those from which the blade support discs 38, 42, and 46 are machined and the axial serrations of both types of disc can be machined on the same diameters. After the support discs 72,74 have been welded into the drive line between the main rotor discs 38,42,46, as shown by the cross-hatched lines W, the seal segments can be assembled onto the support discs by passing them through the serrations in the main discs.

In this construction, contrary to that described for Fig, 2, there is no need for the circumferential extent of each seal segment platform to be equal to that of the blade platfoms it adjoins, though of course their root portions must be of the same form as those of the turbine blades to the extent necessary to enable them to pass through the serrations in the main discs.

Seal strips 80 (see inset to Fig.3) can be housed in grooves in the confronting edges of adjacent platforms so as to bridge the small gaps between the circumferentially adjacent seal segments, and between the seal segments and the rotor blades.

An annulus inner boundary sealing construction according to the present invention, as shown in Fig. 3, provides a seal which will be maintained throughout the operating range of the engine regardless of engine component excursions due to differential thermal expansion and contraction between the static and rotating components of the engine.

As the sealing segments 68,70 are carried on separate support discs 72, 74, instead of the rotor blades, they do not affect the stress levels in the roots of the rotor blades or in the blade support discs.

Claims (16)

1. A rotating seal assembly for a gas turbine engine, the gas turbine engine comprising rotating structure and static structure surrounding the rotating structure, the static structure forming a static circumferential sealing surface, the seal assembly comprising a plurality of sealing segments having radially inner ends secured to a seal support rotor disc and radially outer sealing surfaces which form in combination a rotating circumferential sealing surface arranged to co-operate with the static circumferential sealing surface.

2. A rotating seal assembly according to claim 1, in which the seal segments are attached to the seal support disc by means of serrated roots engaging with correspondingly shaped serrations in the seal support disc.

3. A rotating seal assembly according to claim 1 or claim 2, in which the rotating sealing surface of the assembly includes labyrinth seal means for sealingly engaging the static sealing surface.

4. A gas turbine engine including a rotating seal assembly according to any preceding claim.

5. An axial flow gas turbine engine comprising rotating structure and static structure surrounding the rotating structure, the static structure forming a static circumferential sealing surface, the rotating structure being a seal assembly comprising a plurality of sealing segments having radially inner ends secured to a seal support rotor disc and radially outer platforms which form in combination a rotating circumferential sealing surface arranged to co-operate with the static circumferential sealing surface.

6. An axial flow gas turbine engine according to claim 5, in which the seal support disc is located between adjacent rotor blade support discs which support corresponding rotor blade stages of the gas turbine engine.

7. An axial flow gas turbine engine according to claim 6, in which the seal support disc is welded to the adjacent rotor blade support discs.

8. An axial flow gas turbine engine according to any one of claims 5 to 7, in which the static sealing surface comprises a plurality of stator blade inner platforms.

9. An axial flow gas turbine engine according to any one of claims 5 to 8, in which the seal segments are attached to the seal support disc by means of serrated roots engaging with correspondingly shaped serrations in the seal support disc.

10. An axial flow gas turbine engine according to claim 6, having at least two of the rotating seal assemblies in axial succession, with a blade support disc interposed between successive seal support discs, the seal segments and the rotor blades being attached to their respective support discs by means of root portions engaging with complementarily shaped grooves in the support discs.

11. An axial flow gas turbine engine according to claim 10, the grooves in the respective support discs being at similar radii to facilitate assembly of the seal segments and the blades into their respective support discs.

12. An axial f-low gas turbine engine according tp claim 10 or claim ill, in which the root portions and the grooves are of serrated form.

13. An axial flow gas turbine engine according to any one of claims 5 to 12, in which gaps between circumferentially adjacent edges of the seal segment platforms are provided with sealing strips therebetween to bridge the gaps.

14. An axial flow gas turbine engine according to claim 6, in which edges of the seal segment platforms are closely spaced from adjacent rotor blade platforms and sealing strips are provided between the edges of the seal segment platforms and the rotor blade platforms.

15. A rotating seal assembly substantially as described herein with reference to Fig. 3 of the accompanying drawings.

16. An axial flow gas turbine engine substantially as described herein with reference to Fig. 3 of the accompanying drawings.