Rolls-royce Corporation

Guy David Snowsill

Colin Young

Geoffrey M Dailey

Crispin David Bolgar

Timothy John Scanlon

Jonathan M King

Chloe J Palmer

Paul William Ferra

Leo Vivian Lewis

Jonathan E Holt

Sean Alan Walters

Michael J Sheath

Robert John Irving

Jeffrey Alison Dixon

Benjamin Thomas Littley

Jessica Kennedy

Kali Charan Nayak

Ivan Lars Brunton

Nomesh P Kandaswamy

John Irving

Clive Peter Gravett

Benjamin Littley

Nomesh Kandaswamy

Robert Irving

Timothy John Scanlan

Jonathan Michael King

Geoffrey Matthew Dailey

A fuel delivery system () is shown for delivering the hydrogen fuel from a cryogenic storage system to a fuel injection system in a gas turbine engine. The fuel delivery system includes a pump (), a metering device (), and a fuel heating system () for heating the hydrogen fuel to an injection temperature for the fuel injection system.

Fuel delivery system for delivering hydrogen fuel to a fuel injection system in a gas turbine engine

A bladed disc system for a turbine engine having a disk portion and a plurality of blade portions which are associated with a stator section and an intercavity sealing portion, disc portion shaped such that blade portions are able to fit within firtree slot in disc portion, blade portion having aerofoil section and root section, aerofoil section having portion shaped such that they extend proximate to intercavity sealing portion, disc portion extending from portion that connects with drum to outer edge at which blade portions are connected with disc portion having width transition region in which thickness of disc increases from point at which disc connects to drum to outer edge at which it holds blade portions, and wherein width transition region has curved width transition region with radius r, and an overhanging portion which extends into the intercavity spacing between the width transition region and the intercavity sealing portion.

Bladed disc

A gas turbine is provided for an aircraft comprising an engine core and a core flow path, a fan, a front drum cavity arranged radially inward of the core flow path, and a front bearing chamber. The front drum cavity comprises a front drum inlet, for providing air to the front drum cavity from the core air flow, located downstream of a stage of the compressor, and a front drum outlet, for ejecting air from the front drum cavity to the fan air flow, located axially between the fan and the compressor. The front drum inlet is through a seal, and the front drum outlet is through a spaced gap.

Gas turbine engine

A gas turbine engine rotor arrangement comprising at least one blade and a disc is disclosed. The blade extends radially outwards from the disc and is secured thereto by cooperating shank of the blade and recess of the disc. The shank comprises a bottom surface facing a base surface of the recess, the bottom surface having axially extending peripheral edges. The bottom surface is shaped so that when the engine rotor arrangement is in use, liquid in a cavity between the bottom surface and base surface, acted upon by an unbalanced force in the radially outward direction, is guided by the bottom surface to flow between and away from the axial edges.

Gas turbine engine rotor arrangement

A gas turbine engine has a compressor section with rotational compressor components rotatable with respect to static compressor components. A compressed air bleed arrangement is provided to cool one or more other rotational components of the gas turbine engine. The compressed air bleed arrangement takes a flow of compressed air from the compressor section along an off-take passage. The off-take passage opens in the compressor section at an off-take port. The off-take passage is rotatable, in use, with the rotational compressor components. The compressed air bleed arrangement is operable to provide the air in the off-take passage with higher static pressure than the air in the compressor section at the off-take port, by diffusing the air in the off-take passage. The off-take passage further includes off-take vanes, operable to increase the tangential velocity of the air in the off-take passage compared with the air at the off-take port.

Gas turbine engine havinga rotatable off-take passage in a compressor section

A gas turbine engine comprising a turbine section cooling system and a method of cooling a turbine section of a gas turbine engine is provided. The gas turbine engine comprises in flow series a compressor section, a combustor, and a turbine section, the engine further comprising a turbine section cooling system. The turbine section cooling system including a first compressed air bleed arrangement and a second compressed air bleed arrangement. The first compressed air bleed arrangement bleeds a first flow of compressed air from a high pressure stage of the compressor section. The first flow of compressed air bypasses the combustor and arrives at the turbine section to form a sealing and/or cooling flow at a row of stator vanes upstream of an adjacent rotor disc. The second compressed air bleed arrangement bleeds a second flow of compressed air from one or more lower pressure stages of the compressor section. The second flow of compressed air bypasses the combustor and arrives at the turbine section to form a cooling flow. A first portion of the cooling flow is routed to a front face of the rotor disc and a second portion of the cooling flow is routed to a rear face of the rotor disc.

Gas turbine engine with turbine cooling arrangement

A gas turbine engine comprising a rotor and a stator which define first, second and third cavities; the rotor and stator define a seal therebetween and which is located for sealing between the second and third cavities, the rotor comprises an aperture through which a gas flow passes from the first cavity to the second cavity characterized in that the seal comprises a deflector that extends axially over at least a portion of the aperture to deflect at least a part of the gas towards the rotor.

Turbine apparatus

It is known with regard to particularly cavities below high pressure turbine discs that mixing of hot gas leakage flows through an inner seal with cooling flows can diminish the effectiveness of that cooling flow when presented to other parts for cooling. By providing a path within a wall which is particularly shaped in portions it is possible to provide entrainment of a hot leakage gas flow away from entry into the cavity. Thus, the cooling flow retains a higher cooling effect and maintains its swirling nature in comparison with prior arrangements where mixing with the hot leakage flow occurred.

Flow cavity arrangement

A rotor assembly for a gas turbine engine, the rotor assembly comprises at least two rotors defining a cavity therebetween. A first rotor defines a cooling air inlet in its radially inward portion. A second rotor defines a cooling air outlet in its radially outward portion, such that the cooling air passes radially outwardly through the cavity.

Gas turbine engine rotor ventilation arrangement

Assembly of segmented seals can be problematic in areas of limited space for an introductory tapered mandrel and where there are the problems with damage to the seal surfaces. By the present method and arrangement, a seal formed from segmentsis expanded to an expanded state and then retained at that state by the use of a sacrificial materialin the gapsbetween the segmentsopened by expansion. The sacrificial materialis typically a low melting temperature wax and so is removed when the seal arrangement reaches its operating temperature.

Seal arrangement and a method of seal assembly

It is known there is ingestion of hot gas flowfrom an annular flowin gas turbine engines. This flowis prevented from attaching to the disc by coolant flowpassing through gaps in lock platesby creation of a barrier layer. In order to regulate and enhance this effect spacer protrusionsare provided between ends of lock plates. In such circumstances the gapis controlled and therefore the rate of leakage flowregulated for best cooling effect. This cooling effect and cooling flowcan be further enhanced by providing chutesfor presentation of the coolant flow

Turbine engine cooling

Efficient cooling of a stage of gas turbine engine turbine blades () is achieved by first reducing the pressure of the cooling air after it has been bled from the annulus of the compressor () by passing it through a diffuser (), to a pressure magnitude lower than is required at entry to the turbine blades, then re-pressurizing the bled air up to the required entry pressure, by passing it through a radial compressor defined by a cowl () positioned in close spaced, co-rotational relationship with the downstream face of the associated turbine disk ().

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Derby, United Kingdom

Guy David Snowsill