Rolls-royce Corporation

Anthony John Rawlinson

Peter Ireland

Ian Tibbott

Lynne H Turner

Matthew Adams

Irene Cresci

Peter Philip Ramwell

Peter Ramwell

Lynne Helen Turner

A stator assembly comprises an annular platform () defining a boundary of an annulus and a plurality of guide vanes () for arranging in a circumferential array on the annular platform (). The platform () includes a circumferential array of slots each slot having, at a first end, walls converging () from an annulus facing side towards an opposite side of the platform and then extending parallel () towards a second end. One or more cooling holes () is arranged in a wall () of the slot. Each vane () includes a root portion () which converges in a first region () distal to the root end and then extends parallel towards the root end in a second region (). The root portion () is configured to engage in a slot of the platform () with the first region () abutting a convergent wall portion () of the slot and the second region () spaced from a parallel wall () of the slot. At least one bore () is provided in a parallel wall () of each slot and at least one bore () is also provided in the second region of the root portion (). The bores () in the two components are positioned to align when a vane () is engaged in a slot. The aligning bores () together are configured to receive a pin () for mechanically securing the vane in the slot. A braze joint () may be added between the root portion () and parallel slot walls

Stator assembly for a gas turbine engine

A gas turbine engine component, especially an aerofoil-sectioned nozzle guide vane (NGV), having at least one internal cooling chamber for passage of cooling air, the chamber including leading edge portion and one inlet portion via which cooling air may enter the chamber from feed source, wherein the component includes a partitioning element, e.g. curved or scoop-shaped partitioning plate or wall, provided in the chamber inlet portion and defining within the inlet portion a sub-chamber adjacent the leading edge portion, and wherein partitioning element is configured so the cooling air velocity in the sub-chamber is less than the cooling air velocity in the remainder of inlet portion. The reduced velocity of the cooling air in the sub-chamber adjacent the leading edge serves to increase pressure therein, thereby maintaining desired backflow pressure margin between the feed pressure of the cooling air delivered to the showerhead holes and the gas-path from the combustor.

Internal cooling of engine components

An aerofoil blade or vane for the turbine of a gas turbine engine is provided. The blade or vane includes an aerofoil portion which, in use, extends radially across a working gas annulus of the engine. A coolant inlet is formed at an end of the aerofoil portion for the entry of a cooling air flow into the portion. A corresponding coolant exhaust is formed at the trailing edge of the aerofoil portion for spent cooling air to flow from the portion. A passage within the aerofoil portion connects the inlet to the exhaust. A fence within the aerofoil portion extends radially and forwardly from a start position at the end of the aerofoil portion adjacent the trailing edge to an end position. The passage forms a loop which extends along one side of the fence, wraps around the end position, and extends along the other side of the fence.

Aerofoil blade or vane

A component of a turbine stage of a gas turbine engine is provided, the component forming an endwall for the working gas annulus of the stage. The component has one or more internal plena behind the endwall which, in use, contain a flow of cooling air. The component further has a plurality of exhaust holes in the endwall. The holes connect the plena to a gas-washed surface of the endwall such that the cooling air effuses through the holes to form a cooling film over the gas-washed surface. Each exhaust hole has a flow cross-sectional area which is greater at an intermediate position between the entrance of the hole from the respective plenum and the exit of the hole to the gas-washed surface than it is at said exit.

Endwall component for a turbine stage of a gas turbine engine

A shroud segment for a turbine stage of a gas turbine engine forms an endwall for the working gas annulus of the stage. The segment also provides a close clearance to the tips of a row of turbine blades which sweep across the segment. In use a leakage flow of the working gas passes through the clearance gap between the blade tips and the segment. The segment has a plurality of cooling holes and respective air feed passages for the cooling holes. The cooling holes are distributed over that part of the gas-washed surface of the segment which is swept by the blade tips. The cooling holes deliver, in use, cooling air which spreads over the gas-washed surface. The feed passages are configured such that the delivered air opposes the leakage flow of the working gas.

Turbine stage shroud segment

An aerofoil () for a gas turbine engine () is hollow to define an interior volume () through which cooling air flows. Passages () interconnect the interior volume () with the exterior of the aerofoil (). Each passage () is provided with an inlet () within the interior volume () which is elongated along an axis () parallel with the direction of cooling air flow through the interior volume (). The arrangement reduces any tendency for the passages () to block though the build up of dirt particles.

Aerofoil having a plurality cooling air flows

An aerofoil for a gas turbine engine comprising a pressure wall and a suction wall and defining leading and trailing edges, the walls define a passage into which is supplied a cooling fluid, an array of cooling holes is provided through at least one of the walls to allow the cooling fluid to flow from an interior surface to an exterior surface. The array of holes comprise two groups, the holes of each group are angled to intersect the holes of the other group and are characterized in that the holes of at least one of the groups comprises two or more holes at different angles to one another to vary the porosity of the wall to account for otherwise varying wall temperatures. This arrangement also allows either less coolant mass flow to maintain a constant metal temperature, or a lower metal temperature for a given coolant mass flow.

Cooling arrangement

An aerofoilincludes cooling channelswith transfer passagesbetween adjacent channels. In normal operation, the constriction in each cooling channelensures direct through coolant airflow in the direction of arrowheads A. However, when a channelis blocked by a blockageairflow in adjacent channelsis forced through the transfer passagesin order to provide airflow in the blocked channeland facilitate cooling. This blocked channel airflow is in the direction of arrowheads B. Generally, the cross-section of the transfer passagesis determined for conformity with the outletcross-section of the channelin order to achieve substantial coolant flow balance across the coolant channelsof the aerofoil

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Derby, United Kingdom

Anthony John Rawlinson