Rolls-royce Corporation

Mark A Halliwell

Steven D Lawer

Alec George Dodd

Adrian Lewis Harding

Henry Tubbs

Anthony B Phipps

Harald Schiebold

Richard Ab McCall

Marcus McBride

Steven B Morris

Philip J Cooke

Paul H Edwards

Within gas turbine engines is necessary to provide nozzle guide vanes between stages of the engine. These vanes are presented in vane segments and it is desirable to prevent leakage to retain engine operation efficiency as well as to avoid hot gas impingement on inappropriate parts of the engine. By use of anti-rotation blocks twisting between the segments can be prevented and therefore the segments retained in alignment. However, thermal distortion may open a chordal seal provided to inhibit gas flow leakage. By provision of chordal bumps it is possible to prevent forward rocking which will inhibit gaps between the chordal seal and an engaging support ring surface. Furthermore the anti-rotation blocks will generally incorporate appropriate mating surfaces to engage the chordal bumps across two or more vane segments to facilitate retention of vane segment alignment while achieving adjustment for thermal distortion.

Gas turbine engine vane arrangement

A casing arrangement of a gas turbine includes a casing member, which is formed to extend at least partially around the component. The casing member defines a fluid flow path for the flow of a cooling fluid therethrough.

Casing arrangement

A turbine blade tip clearance control system has a rigid two part outer casing () which sandwiches a control ring () therebetween, and an air pressurised flexible inner casing () which carries shroud segments () within it. Struts () span the annular space between the casings () and prevent flexing of casing () until blade tip clearance needs adjusting, whereupon, ring () is heated, along with the adjacent portion of outer casing () and expands, allowing casing () to flex outwards, thus lifting the shroud segments () away from the blade tips (). Closure of the tip clearance is achieved by cooling ring (), the resulting contraction thereof, via the struts (), flexing the inner casing () and shroud segments () inwards, against the air pressure.

Gas turbine blade tip clearance control structure

A seal segment ( ) as described for a seal segment ring ( ) of a turbine ( ) in a gas turbine engine ( ). The seal segment ( ) has an inner surface ( ) adapted to face the turbine blades ( ) in use. Path means ( ) is defined in the seal segment ( ). The path means ( ) is adapted to extend, in use, generally parallel to the principal axis of the turbine and has downstream inlet means ( ) through which a cooling fluid to cool the seal segment can enter the path means ( ) and upstream outlet means ( ) from which the cooling fluid can be exhausted from the path means ( ). The cooling fluid can flow along the path means ( ) in a generally upstream direction opposite to the flow of gas through the turbine.

Seal segment for a turbine

A sealing element ( ) for a turbine of a gas turbine engine includes a radially inner surface region provided with an integrally formed seal structure ( ) comprising a plurality of radially inwardly projecting walls ( ). The walls may be abradable and may define cells ( ) for receiving an abradable sealing material.

Abradable seals

A fluid seal comprises a plurality of seal segments (24) which form a peripheral ring at the tip of turbine rotor blades (20). The upstream end (25) of each seal segment (24) is mounted on the static structure (27). The downstream end (26) of each seal segment (24) is supported by an inner flange (34) on a vane (32). The vane (32) is also provided with a radially outer flange (36) which locates in a radially inclined slot (31) in the casing (30). During engine transients, the vane (32) expands axially forward relative to the casing (30). The outer flange (36) rides up the inclined slot (31) which causes the seal segment (24) to move radially outward and increases the seal clearance. However, during stabilized engine running, the relative axial movement between the casing (30) and the vane (32) is reduced. The outer flange (36) contracts down the inclined slot (31) to move the seal segment (24) radially inwards; the combination reduces the tip clearance and prevents excessive gas leakage.

Fluid seal

An annular seal is provided between a turbine rotor disc and the adjacent static structure. The seal comprises rotor fins extending from the turbine disc for rotation therewith and an abradable honeycomb layer radially outward of the rotor fins. The abradable honeycomb layer is located on the inner diameter of a sealing ring attached to the static structure. The sealing ring is moveable in an axial direction so that in operation the honeycomb has one axial position for transient conditions and an alternative axial position for stabilized running. The seal clearance can thus be optimised at the two axial positions reducing the leakage flow through the seal.

A bladed rotor and surround assembly comprising an annular casing, a bladed rotor element that is rotatable about an axis concentrically within the casing, and an annular shroud liner. The shroud liner, typically made up of an annular array of circumferentially abutting shroud liner segments, is disposed within the casing in an annular radial space defined between the casing and an outer circumference of the bladed rotor. The shroud liner segments have location members to locate each segment within the casing. The location members and the annular radial space are configured to enable axial insertion of the shroud liner segment between the bladed rotor and the casing. In addition the location members and the annular radial space allow a limited amount of radial translation of the shroud segment during insertion. The location members also provide a positive radial location to prevent radial translation of the shroud segment once each shroud segment is in a final assembled position.

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Derby, United Kingdom

Mark A Halliwell