GB2272946A - Gas turbine engine interstage seal. - Google Patents

Abstract

A gas turbine engine sealing apparatus to seal the inner boundary of the turbine gas flow annulus comprises one or more sealing rings 66, 68 located between rotors 38 and 42, and 42 and 46. The sealing ring 66 is provided with upstream and downstream dogs 80, 82 which engage underneath axial projections 76 and 74 on the rotors 38, 42. The ring 68 has a continuous ring (96, fig. 6) which engages under the projections 74 on the rotor 42, and dogs (100, fig. 6) which engage under the projection 78 on the rotor 46. Metallic sealing strips, seated in slots 90, may be used to seal between sealing rings and rotor blade platforms. Use of this apparatus avoids the manufacturing and stress problems inherent in the use of seal wings cast into rotor blades. <IMAGE>

Description

GAS TURBINE ENGINE SEALING APPARATUS This invention relates to a gas turbine engine sealing apparatus for the sealing of 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 combustion apparatus 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 are 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 shaft, 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 passage is partially defined by platforms on the inner ends of the stator blades. These platforms are usually sealingly engaged by sealing elements secured to rotors on which the rotor blades are located.

The relatively large distances between the shaft bearings, for example up to nine metres, results in excessive rotor blade movement relative to the gas generator casing due to differential thermal expansion between the shaft and the casing. Thus the types of seals between the rotating and static components of the gas generator turbine which are typical of lightweight turbines derived from aero-engines are not practical.

Also in the case of a known type of heavyweight gas generator turbine, the turbine rotors comprise drums which are welded together. Such a form of construction limits the options available for providing rotating sealing elements to co-operate with the platforms on the stator blades. This limitation arises because a welded construction does not allow the insertion of extra components between the rotors to carry the sealing elements.

In the case of relatively low power engines, a seal can be achieved by casting projections or 'wings' onto the platforms at the inner ends of the rotor blades. These projections on the rotors of adjacent stages abut one another to form a seal.

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

This present invention seeks to provide a form of 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 gas turbine engine sealing apparatus including at least one unitary sealing ring located between an adjacent pair of turbine rotors, the unitary sealing ring including a circumferential sealing surface arranged to co-operate with a static sealing surface, the unitary sealing ring and co-operating turbine rotors including ring location and retention means, the location and retention means on the unitary sealing ring being sized and spaced apart to allow the location of retention means to pass through rotor blade retention slots through the rim of at least one of the turbine rotors.

The unitary sealing ring can have at least at one end a plurality of spaced apart radially inwardly projecting legs, each leg terminating in an axially extending protection, the spacing between the legs being equal to the spacing of the blade retention slots on one of the turbine rotors.

The unitary sealing ring can include a plurality of the said projecting legs at each end of the sealing ring.

The unitary sealing ring can include at its opposite end a continuous axially projecting rim having an inner diameter greater than the outer diameter of one of the turbine rotors.

At least two of the turbine rotors can include a plurality of spaced apart projections engageable by the location of retention means of at least one unitary sealing ring.

A unitary sealing ring can be located and retained between adjacent ones of two pairs of turbine rotors, the two pairs of turbine rotors having a common rotor.

Circumferential slots can be provided at the edges of the or each sealing ring, and sealing strips can be located in each circumferential slot and engaging in corresponding slots in rotor blade platforms adjacent the edges of the or each unitary sealing ring.

The circumferential sealing surface on the or each unitary sealing ring can comprise a labyrinth seal.

Ribs can be provided on the or each unitary sealing ring extending between the legs on opposite sides of the ring.

The ribs can be inclined at an angle equal to the stagger angle of the rotor blade retention slots on one of the turbine rotors.

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.1, showing part of a gas generator turbine incorporating a known type of gas flow path sealing construction, Fig. 3 shows a gas generator turbine incorporating one form of sealing apparatus according to the present invention, Fig. 4 shows a perspective view of part of one of the turbine rotors shown in Fig. 3, Fig. 5 shows part of one of the unitary sealing rings shown in Fig. 3, and Fig. 6 shows part of another one of the unitary sealing rings shown in Fig. 3.

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 shown 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 59B, 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.

These problems, as presented in an engine of larger size than in Fig. 2, can be addressed as illustrated in Fig. 3, in which similar components described with reference to Fig. 2 have been allotted the same references.

In order to seal the annulus inner boundary between adjacent turbine rotors, unitary sealing rings 66 and 68 are provided, the rings 66 being located between the rotors 38 and 42 and the rings 68 being located between the rotors 42 and 46.

Referring to Figs. 4 and 5 the turbine rotor 42 is provided at its circumference with a plurality of equispaced slots 70 having serrations 72 in order to retain the turbine rotor blades 44 which are formed with roots to correspond with the serrations. The rim of the rotor is wider than the body of the rotor so that spaced apart portions of the rim 74 extend on each side of the body of the rotor.

The rim of the turbine rotor 38 is formed with slots and serrations in a similar manner to the rim of the rotor 42 in order that the rotor blades 40 can be retained upon the rotor 38. The rim of the rotor 38 only has projections 76 and the slots 70 in the rim of the rotor 42 are axially aligned with the corresponding slots in the rim of the rotor 38. In both cases the slots in the rim of the rotors 38 and 42 are formed parallel with the axis of rotation of the rotors 38, 42 and 46. That is to say that there is no stagger angle in the circumferential slots in the rims of the rotors 38 and 42.

The rim of the rotor 46 is also provided with retention slots and serrations in order that the rotor blades 48 can be retained upon the rotor disc 46. The rim of the rotor 46 only has projections 78 which extend in an upstream direction, and the slots in the rim of the rotor 46 are provided with a stagger angle.

Referring more particularly to Fig. 5 the unitary sealing ring 66 is provided with upstream and downstream axial location dogs 80 and 82 which are connected together by webs 84. The axial location dogs 80 and 82 are spaced apart by an amount equal to the spacing between the blade retention slots in the turbine rotors 38 and 42.

The unitary sealing ring 66 includes an integral ring 86 having on its outer surface a pair of spaced apart labyrinth seals which co-operate with the static sealing surface 30A attached to the stator vanes 30.

The edges of the ring 86 are provided with circumferential slots 90 in which metallic sealing strips (not shown, but well known in the industry) are located during assembly.

The sealing ring 66 is located in position by passing the sealing ring over the rim of the rotor 46 so that the axial location dogs 80 and 82 pass through the serration 70 in the rim of the rotor 42. The radius of the dogs is the same as the radius at the bottom of the slots. When the sealing ring 66 is located between the rotors 38 and 42 it is indexed by half of the pitch between adjacent serrations 70 so that the axial location dogs 80 and 82 pass beneath the projections 76 and 74 respectively on the rotors 38 and 42. The sealing ring 66 can then be secured in position by any appropriate anti-rotation mechanism.

Referring to Fig. 6, the unitary sealing ring 68 is similar in construction to the ring 66 in that it has an integral ring 92 having on its outer surface a pair of labyrinth seals 94 which co-operate with a static sealing surface 32A attached to the row of stator vanes 32. At the upstream end of the ring 68 a continuous ring 96 is attached to the outer ring 92 by a web 98.

At the downstream end of the ring 68 a plurality of equi-spaced location dogs 100 extend from legs 102, and the dogs are set at the stagger angle of the blade retention slots in the turbine rotor 46. The axial location dogs 100 are spaced apart at the same spacing as that of the blade retention slots in the turbine rotor 46. The legs 102 are connected to the web 98 by webs 104 which are inclined also at the stagger angle of the blade retention slots in the turbine rotor 46.

The unitary sealing ring 60 is also fitted from the rear of the turbine, the ring 68 passing over the rim of the turbine rotor 46 and the legs 102 pass through the blade retention slots in the rim of the turbine rotor 46.

When the ring 68 is located between the rotors 42 and 46, the ring 68 is indexed by half of the pitch between adjacent serrations in the rim of the disc so that the location dogs 100 pass beneath the projections 78, the ring 96 therefore being located beneath the projections 74 on the turbine rotor 42.

The ring 68 can be secured in position by any appro priate anti-rotation mechanism.

Claims (12)

CLAIMS:

1. Gas turbine engine sealing apparatus including at least one unitary sealing ring located between an adjacent pair of turbine rotors, the unitary sealing ring including a circumferential sealing surface arranged to co-operate with a static sealing surface, the unitary sealing ring and co-operating turbine rotors including ring location and retention means, the location and retention means on the unitary sealing ring being sized and spaced apart to allow the location and retention means to pass through rotor blade retention slots at the rim of at least one of the turbine rotors.

2. An apparatus as claimed in Claim 1 in which the unitary sealing ring has at least at one end a plurality of spaced apart radially inwardly projecting legs, each leg terminating in an axially extending projection, the spacing between the legs being equal to the spacing of the blade retention slots on one of the turbine rotors.

3. An apparatus as claimed in Claim 2 in which the unitary sealing ring includes a plurality of the said legs of each end.

4. An apparatus as claimed in Claim 2 in which the unitary sealing ring includes at its other end a continuous axially projecting rim having an inner diameter greater than the outer diameter of one of the turbine rotors.

5. An apparatus as claimed in any one of the preceding claims in which at least two of the turbine rotors include a plurality of spaced apart projections engageable by the location and retention means on at least one unitary sealing ring.

6. An apparatus as claimed in Claim 5 in which a unitary sealing ring is located and retained between adjacent ones of two pairs of turbine rotors, the said two pairs of turbine rotors having a common turbine rotor.

7. An apparatus as claimed in any one of the preceding claims in which circumferential slots are provided at the edges of the or each unitary sealing ring, sealing strips being located in each said slot and engaging in corresponding slots in rotor blade platforms adjacent the edges of the or each unitary sealing ring.

8. An apparatus as claimed in any one of the preceding claims in which the circumferential sealing surface on the or each unitary sealing ring is a labyrinth seal.

9. An apparatus as claimed in any one of the preceding Claims 2 to 8 in which ribs are provided on the or each unitary sealing ring extending between the legs on opposite sides of the ring.

10. An apparatus as claimed in Claim 9 in which the ribs are inclined at an angle equal to the stagger angle of rotor blade retention slots on one of the turbine rotors.

11. A gas turbine engine sealing apparatus constructed and arranged for use in operation substantially as herein described, and with reference to Figs. 3 to 6 of the accompanying drawings.

12. A gas turbine engine including a sealing apparatus as claimed in any one of the preceding claims.