GB2119027A

GB2119027A - Turbine assembly for a gas turbine engine

Info

Publication number: GB2119027A
Application number: GB08211928A
Authority: GB
Inventors: Thomas Douglas Cribbes
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 1982-04-24
Filing date: 1982-04-24
Publication date: 1983-11-09

Abstract

The turbine assembly has features designed to alleviate the problems caused by cooling air flowing past the seal between rotating and static members 28, 18 and into the main gas flow annulus. The features comprise passages 33 in the platforms 28 of the rotor blades 17 which interconnect the seal area with the outer surface of platforms in a manner and a region where the flow of air is less detrimental. <IMAGE>

Description

SPECIFICATION Turbineassemblyfora gas turbine engine This invention relates to aturbine assemblyfora gas turbine engine. In such assemblies it is normal to have a turbine rotor disc carrying at its periphery a stage ofturbine rotor blades. Upstream ofthe stage of rotor blades a stage of static nozzle guide vanes direct the motive gases upon the rotor blades at a predetermined angle and velocity, while downstream of the stage of rotor blades a second stage of static nozzle guide vanes performs the same duty for a subsequent stage of rotor blades.

It is usual to provide a flow of cooling airto the region ofthe turbine radially inside the gas flow annulus, this air being used to cool the disc and blades. In orderto avoid the unrestricted flow ofthis cooling air into the gas flowannulus, various seals are provided. In particular, seals are provided between the inner platform of the turbine blade or other adjacent structure and the static structure adjacent to orforming part of the nozzle guide vanes.

It is not possible to make these seals perfect, and there is inevitablysome leakagethrough them. This leakage can be detrimental to the aerodynamic efficiency ofthe various components, and in particular the leakagefromthe seal on the upstream face ofthe rotor disc can have a detrimental effect on the performance of the rotor blades.

The present invention provides a turbine assembly in which this leakage air is dealtwith in a less detrimental manner.

According to the present invention a turbine assemblyfora gasturbine engine comprises a rotor carrying rotor blades each having a platform, an aerofoil, at least one seal feature which cooperates with a corresponding seal feature on the static structure to form a seal which limitstheflow of air from inside the seal into the gas flow of airfrom insidethe seal into the gas flow annu lus of the engine, and passages formed in the platform from adjacentto said seal projection to the outer surface of said platform, the passages serving to allow leakageflowto be directed from the seal to the platform surface.

The passages preferably open on to the upper surface ofthe platform in the gas passage between the blade aerofoil, and may be directed to cause the leakage flow to follow the local gas flow.

The seal projections may comprise a plurality of part annular fins which coactwith those on adjacent blades to form a plurality of concentric annular seal fins. the passages may then extend into the space between the fin portions.

The invention will now be particularly described, merely byway of example, with reference to the accompanying drawings in which: Fig. lisa partly broken-away side elevation of a gas turbine engine having a turbine assembly in accordance with the invention, Fig. 2 is an enlarged section through the turbine assembly ofthe engine of fig. 1,and Fig. 3 is a perspective view of one ofthe turbine blades of fig. 2.

Infig. 1 is shown a gas turbine engine 10 having its external casing 11 broken-away at 12 to expose to viewthe turbine assembly ofthe engine.

Overall, the construction and operation of the engine is conventional in that a number of compressors compress intake air, which is mixed with fuel and burnt in the combustion chamber 13, the hot gases thus produced driving turbines 14 and 15 and exhausting from the engine to give propulsive thrust. The turbine 14 and 15 drive respective compressors through drive shafts.

Figu re 2 shows the turbine 14 in enlarged cross section. The downstream end ofthe combustion chamber 13 is visible, and a stage of nozzle guide vanes 16 direct the hot gases from the chamber 13 on to a stage ofturbine rotor blades 17. Each of the nozzle guide vanes 16 comprises an aerofoil and inner and outer platforms 18 and 19 respectively which serve to confine the gas flow from the chamber 13 into an annular flow path. The platforms 18 and 19 additionally serve a mechanical purpose, and are provided with various mounting flanges bywhichthevanesare mounted from fixed structure and some of which may be used to carry further seals etc.

One such flange 20 extends inwardly from the inner platform l8andcarriesanannularseal member 21.

The working face of this member has two annular seal fins 22 and 23 which,togetherwith the projecting edge 24 ofthe platform 18, made up the three staticfins of a labyrinth seal whose rotating parts are carried from the turbine rotor described below.

The turbine rotor comprises a turbine rotor disc 25 which carries at its peripherythe rotor blades 17. Each blade is mounted in the disc rim by a firtree root 26 (see Fig 3) which engages with a correspondingly shaped slot in the disc rim. The root 26 then carries, via a shank27, the platform 28 and aerofoil 29 of the blade. The aerofoil 29 performs the main purpose of the bladesthatistoextractpowerfromthe hot gas steam to drive the disc25 and its respective com prnss- or, whilethe platform 28 serves, like the vane platforms 18 and 19, to define a boundary of the gas flow annulus.

In orderto meterthe ingressofcooling airtothegas flow annulus, it is necessary to form a seal between the rotor blade platforms 28 and the nozzle guide vane platforms 18.The static seal member 21 carried from the platform 18 has been described above, and to complete the seal the forward edge of the platform 28 is provided with part-annular projections 30,31 and 32. These projections cooperate with those from adjacent blades to form annular seal fins, which are interdigitated with the static fins made up from the projections 22,23 and 24to form a labyrinth type of seal.

The main purpose ofthe seal in the present instance isto prevent the high pressure cooling and sealing air inboard of the seal flowing in an unrestricted manner with the gas flow annulus. However, there is always a degree of leakage air allowed for cooling purposes which, with the structure so far described, will flow through the annular gap between the platforms 18 and 28. This has been found to affect the performance ofthe aerofoils 29 deleteriously by building up the boundarylayeroverthe platform and spoiling the flow over the inner parts of the aerofoil.

In the present invention, therefore, passages 33,34 and35areformed in the platform,the passages extending from between the projections 31 and 32 to the outer surface ofthe platform in the gas passage between ,. ia 1; oi-oil 29. Since the pressure at the outletofthe passages is lowerthan that atthe gap between the platforms 18 and 28, the cooling airwill be encouraged to flow along the passages rather than directlythrough the gap and into the gas stream.

By arranging thatthe leakage flows through these passages, various advantages may be obtained. Thus the initial part ofthe aerofoils are not as badly effected bythethickened boundary layer as they would have been. Again, because the passages breakthrough the surface at a shallow angle the leakage air will be encouraged to flow with the main gasflow, and indeed the direction ofthe passages at outlet may be arranged to be the same as that ofthe main gas flow so asto reduce the discrepancy still further.

Itwill be appreciated thatthistechique could be applied to other rotors oftheturbinethan the high pressure rotor described herein, and that it may be desirable to use more complex passages than those illustrated. Again, more or less than the three passages illustrated could be used, and the details ofthe seal could well be changed.

Claims

1. Aturbine assemblyfor a gasturbine engine comprising a rotor carrying rotor blades each having a platform, an aerofoil, at least one seal feature which cooperates with a corresponding seal feature on the static structure to form a seal which limits the flow of airfrom inside the seal into the gas flow annulus of the engine, and passages formed in the platform from adjacent to said seal projections to the outer surface of said platforms, the passages serving to allow leakage flowto be directed from the seal to the platform surface.

2. Aturbine assembly as claimed in claim 1 and in which said seal features on the blade comprise a plurality of part-annularfins on the blade which cooperate with the fins on adjacent blades to form a plurality of concentri 3nnular seal fins.

3. Aturbine assembly as claimed in claim 2 and in which said seal features on the static structure comprise annularseal fins interdigitated with the fins on the blades to form a labyrinth seal.

4. Aturbine assembly as claimed 2 or3 and in which said passages open into the space between two of said partannularfins.

5. Aturbine assembly as claimed in any one of the preceding claims and in which the passages open on totheuppersurface ofthe blade platform in the gas passage between the aerofoils.

6. Aturbine assembly as claimed in claim 5 and in which at leastthe end portions ofthe passages are directedtocausetheleakageflowtofollowthe local gas flow.

7. Aturbine assembly substantially as hereinbefore particularly described with referencetothe accompanying drawings.

8. A gas turbine engine having a turbine assembly as claimed in any one ofthe preceding claims.