Rolls-royce Corporation

Rolls-royce North American Technologies Inc.

Rolls-royce High Temperature Composites Inc.

Rolls-royce Power Engineering Plc

Keith C Sadler

Michael J Whittle

Michiel Kopmels

Peter Jeffrey Goodman

Ted Joseph Freeman

Matthew D Thomason

Aaron D Sippel

James Kell

Andrew D Norton

Alexandra Baucco

Stephen Harris

Andrew T Holt

Mark Timothy Mitchell

Bogachan Alpan

Michael J Jago

David J Hunt

Keith R F Speed

Andrew Thomas Napper

A method comprises inspecting a ceramic matrix composite component assembled in a gas turbine engine to determine an extent of damage to the ceramic matrix composite component, determining a repair technique to repair the damage to the ceramic matrix composite component based on the extent of damage to the ceramic matrix composite component, and repairing the ceramic matrix composite component using the repair technique.

Methods for health monitoring of ceramic matrix composite components in gas turbine engines

A gas turbine engine includes a shroud ring, a vane ring, and a secondary flow assembly. The shroud ring extends around an associated turbine wheel and includes a carrier segment and a blade track segment comprising ceramic matrix composite materials. The blade track segment and the carrier segment defining a shroud cavity radially therebetween. The vane ring includes a heat shield comprising ceramic matrix composite materials that forms an outer platform, an inner platform, and an airfoil defining a heat shield cavity radially through the heat shield. The secondary flow assembly includes a recirculating flow circuit having a discharge tube and a vane pressurizing tube; and the secondary flow assembly is configured to cool the blade track segment. The flow circuit is further configured to pressurize the heat shield cavity thereby providing a seal against the gases from the primary gas path entering the heat shield cavity.

Turbine engine with reused secondary cooling flow

A turbine assembly for use with a gas turbine engine includes a bladed wheel assembly, a vane assembly, and an inner seal. The bladed wheel assembly is adapted to interact with gases flowing through a gas path of the gas turbine engine. The vane assembly is located upstream of the bladed wheel assembly and adapted to direct the gases at the bladed wheel assembly. The inner seal is configured to block gases from passing around the vane assembly.

Turbine section assembly with ceramic matrix composite components and interstage sealing features

A turbine vane assembly adapted for use in a gas turbine engine includes a plurality of turbine vanes, an outer vane support, and an inner vane support. The plurality of turbine vanes comprise ceramic matrix composite material and are adapted to interact with hot gases flowing through a gas path of the gas turbine engine during use of the turbine vane assembly.

Turbine vane assembly incorporating ceramic matrix composite materials

A turbine vane assembly includes a plurality of turbine vanes, an outer vane support, and an inner vane support. The turbine vanes are adapted to interact with gases flowing through a gas path of the gas turbine engine. The outer vane support is arranged radially outward of the turbine vane and configured to receive aerodynamic loads from the turbine vanes. The inner vane support is arranged radially inward from the outer vane support to locate the turbine vane assembly therebetween.

An assembly adapted for use in a gas turbine engine includes a carrier and a blade track segment. The carrier extends at least partway about an axis. The blade track segment is supported by the carrier radially relative to the axis to define a portion of a gas path of the assembly.

Turbine shroud ring segments with ceramic matrix composite components

A turbine blade having labyrinth of internal channels for circulation of coolant, the novel labyrinth geometry includes; inlet arranged on axially upstream face of terminal portion leading to an upstream duct portion having first section adjacent the inlet and a second section, the second section having reduced cross section compared to first section; a first passage intersecting first section and extending through the blade body towards tip of the blade, a proximal end of leading edge passage arranged to capture incoming air flow; a main blade passage intersecting a downstream duct portion, the downstream duct portion arranged in axial alignment with the upstream duct portion but separate from upstream duct portion; and a restrictor passage intersecting with the mid-blade passage and extending towards a mid-blade duct portion, the mid-blade duct portion in axial alignment with the upstream and downstream duct portions and in fluid communication with the upstream duct portion.

Thermal shielding in a gas turbine

A turbine disc assembly comprises a turbine disc mounted coaxially on a shaft and, in use, arranged in the path of a hot work fluid flow. A cover plate is axially displaced from the turbine disc in a direction upstream with respect to the work fluid flow. The cover plate comprises a radially inner annular disc component and a radially outer annular disc component, the annular disc components are arranged coaxially with each other.

Turbine disc assembly

A guide vane assembly for a turbomachine, comprising a guide vane and a liner: the guide vane comprising an aerofoil portion having a radially outer platform, wherein one of the outer platform and the liner has a hook element, the hook element comprising an opening and a circumferentially extending channel, and the other one of the outer platform and the liner comprises a retaining element having a head portion, wherein in an operational orientation the head portion located within the channel and is too wide to be withdrawn through the opening of the hook element.

Guide vane assembly

A component such as a turbine blade of a gas turbine engine has a cooling arrangement comprising a cascade impingement array in which cooling air flows from a supply chamber through impingement passages in webs to an internal passageway comprising first and second limbs. The cooling air flow through the limbs, provides improved heat transfer compared with continued impingement cooling in the chordwise direction of the blade.

Air-cooled component

The present invention provides an airfoil for a first stage turbine blade having an external surface with first and second sides. The external surface extends spanwise between a hub and a tip and streamwise between a leading edge and a trailing edge of the airfoil. The external surface includes a contour substantially defined by Table 1 as listed in the specification.

First stage turbine airfoil

A cooling arrangement is provided in which an incident coolant flow is presented to a passage such that a proportion of that flow is diverted in order to create a standing relative overpressure in a chamber compared to the actual static pressure of the flow in the passage. In such circumstances through flow transfer apertures in the walls of the chamber forced coolant flows are presented for impingement upon surfaces and therefore greater heat transfer. Generally, coolant flow is presented to both ends of the chamber in order to create the relative standing overpressure in the chamber relative to the current static pressure presented in the passage through the incident flow. In such circumstances through utilizing the dynamic rotation or other motion of a component incorporating the arrangement a higher relative pressure is achieved for greater force coolant flow projection compared to the incident or actual static pressure of the incident coolant flow in the.

Cooling arrangement

A component such as a turbine blade for a gas turbine engine has a cooling arrangement comprising a supply passagewithin the blade from which extend cooling passagesof a first array which open at discharge openings, for example at a trailing edge of the blade. Cooling passagesof a second array extend into the blade from discharge openings, and intersect the passagesThe passagesterminate short of the supply passageAs a result of this arrangement, the limiting flow area for cooling air is defined by the cooling passagesof the first array, and is not affected by the accuracy with which the cooling passagesof the second array intersect the cooling passagesof the first array.

Component having a cooling arrangement

A heat transfer arrangement comprises a target member (or), and an impingement member (). The impingement member () defines a fluid path (or) there through to direct fluid onto the target member. The arrangement includes a fluid directing formation () on the impingement member (). The fluid path (or) extends through the fluid directory formation () such that the fluid path directs fluid to exit there from at an exit angle that is substantially orthogonal to the fluid directing formation.

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Bristol, United Kingdom

Keith C Sadler