Rolls-royce Corporation

Desmond Close

Anthony Pidcock

Michael Paul Spooner

Geoffrey M Dailey

William C Kwan

Lance P Bell

Simon Cross

Neil Gater

A gas turbine engine pre-diffuseris generally annular, including radially inner and radially outer wallsandand a generally cylindrical midlinedefined between the walls. The pre-diffuserincludes a central memberwhich forces air flowing through the pre-diffuserto separate, initially to be directed away from the midlinebefore subsequently being allowed to diffuse back towards the midline. The majority of the diffusion takes place on the walls of the central member and is thus in an inner region of the annulus of air ejected from the pre-diffuser to pass to the combustor. Any boundary losses therefore do not significantly effect air at the extremities of this annulus, this air being destined for the annuli of the combustor and requiring relatively high energy levels.

Diffuser for gas turbine engine

A wall element () for a wall structure () of a gas turbine engine combustor (). The wall element () comprises a main member () with an upstream edge region () and a downstream edge region (). A plurality of heat removal members () are provided on the main member (). The downstream edge () of the wall element and/or the downstream facing surface of the heat removal members closest to the downstream edge () are provided with a thermally resistant coating.

Wall elements for gas turbine engine combustors

A gas diffusion arrangement () for a gas turbine engine () is disclosed. The gas diffusion arrangement () has an inlet () and an outlet () for the gas, and comprises diffusion means () having an upstream region () and a downstream region (). Distribution means () is arranged between the downstream region () and the outlet () to distribute the gas into a desired flow pattern. The area ratio of the cross-sectional area of the downstream region () to the cross-sectional area of the upstream region () is greater than 1.5 to 1.

Gas diffusion arrangement

A wall element for use as part of an inner wall of a gas turbine engine combustor wall structure is of cast construction and includes a plurality of cooling apertures provided therethrough and formed during the casting process. The cooling apertures may be located in positions where they could not be conventionally formed by laser drilling.

Combustion apparatus

A wall element ( ) for a wall structure ( ) of a gas turbine engine combustor ( ). The wall element ( ) comprises a main member ( ) with an upstream edge region ( ) and a downstream edge region ( ). A plurality of heat removal members ( ) are provided on the main member ( ). The downstream edge ( ) of the wall element and/or the downstream facing surface of the heat removal members closest to the downstream edge ( ) are provided with a thermally resistant coating.

A wall structure for a gas turbine combustor arranged to have a general direction of fluid flow therethrough includes inner and outer walls defining a space therebetween. The inner wall is made up of a plurality of tiles having axial edges aligned generally with the direction of fluid flow, a gap being defined between axial edges of adjacent tiles. Orifices are provided within the axial edges to direct leakage air passing through the gaps to give the leakage air a flow component in the general direction of fluid flow through the combustor.

A wall element ( A, B) for a combustor ( ) of a gas turbine engine ( ). The wall element ( A, B) defines an axis. In use, the axis is arranged generally parallel to the principal axis of the engine ( ). In one aspect, the length of the wall element ( B) along the axis is at least substantially 20% of the length of the wall element ( B) transverse to the axis. In another aspect, the wall element ( A, B) has a first pair of opposite edges extending transverse to the axis and a second pair of opposite edges ( ) extending transverse to the first pair, at least one of the second pair of edges ( ) being angled relative to the axis of the wall element ( A, B).

A combustor arrangement ( ) for a gas turbine engine ( ) comprising a combustion chamber, fuel nozzles( ), and a bled diffuser ( ) located upstream of said combustion chamber to, in use, direct an airflow from an upstream compressor ( ) into the combustor ( ). The fuel nozzles ( ) arranged in use to supply fuel into the combustion chamber where it is mixed and combusted with the airflow from the compressor ( ). The bled diffuser ( ) adapted to bleed off a portion of said airflow from a main airflow into the combustion chamber. At least one bleed duct ( ) is connected to the bled diffuser ( ) to return and direct the air bled from the diffuser ( ) to a main gas flow through the engine ( ) at a location downstream of the fuel supply means ( ). The bled air from the diffuser ( ) preferably providing cooling of a part of the gas turbine engine ( ), for example part of the turbine ( ) or combustor outlet vane ( ), downstream of the combustor ( ).

Combuster arrangement

A fuel injector assembly designed to withstand a predetermined impact load resulting from ingestion of a bird or other foreign object into the combustion stage of a gas turbine engine comprises a body member adapted to be mounted adjacent to and to extend into the combustion chamber and an abutment member disposed downstream of the injector assembly in the direction of air flow through the combustion stage, the arrangement being such that the body member moves into contact with the abutment member when subjected to an impact load in excess of the impact load it is designed to withstand.

Gas turbine engines

A double wall structure for a gas turbine engine has an inner wall comprising a number of tiles. The outer wall is provided with a number of apertures through which air is directed into the space between the two walls. Inclined apertures are provided in the tiles so that cooling air can pass into the combustion chamber and form a cooling film underneath the tile. Each tile is provided with a number of pedestals. The orientation of the inclined apertures is such that the axis of each aperture lies upon an unobstructed channel between the pedestals.

A plurality of heat resistant tiles are provided which combine to form an internal liner around the walls of an annular combustor. In a first embodiment a plurality of tiles are arranged in an overlapping relationship to form a fully annular shield around an upstream bulkhead wall. In a second embodiment a plurality of tiles are arranged row by row in an overlapping relationship to form a protective shield around the combustors radially spaced sidewalls. In both embodiments the tiles are attached to the combustor in such a manner that at least one edge of the tiles is clamped against the combustor by an overlapping portion of a neighboring tile. The underside of the tile is exposed to a high pressure flow of cooling air which issues as a film over the exposed surface of the tile. The underside of each tile is sealed by virtue of the clamping effect holding the overlapping edges of adjacent tiles in sealing engagement. The arrangement reduces cooling air leakage on the underside of the tiles, minimizes the number of attachment means required per tile and maximizes the tile surface area available for effusion cooling purposes.

Arrangement of heat resistant tiles for a gas turbine engine combustor

The airflow into a gas turbine engine combustion chamber e.g. through the primary air inlet and the dilution air inlets is directionally controlled by airflow directing inserts located in the air inlets. The inserts preferably comprise an open-ended cellular honeycomb structure.

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Derby, United Kingdom

Desmond Close