Abstract: There is disclosed a stator vane section for a gas turbine engine. The stator vane comprises an aerofoil, a first platform at a radial end of the aerofoil, a cooling chamber within the aerofoil and a guide duct within the cooling chamber. An impingement chamber is defined between the first platform and a perforated impingement plate provided on the first platform to receive cooling air through the impingement plate. The guide duct is coupled to the first platform so that a collecting chamber extending around the guide duct is defined between the guide duct and the first platform and the collecting chamber is configured to receive cooling air from the impingement chamber via a plurality of collecting channels, and is configured to direct cooling air to the cooling chamber.

Abstract: An end-wall component of the mainstream gas annulus of a gas turbine engine having an annular arrangement of vanes, the component including a cooling arrangement having ballistic cooling holes (33) through which, in use, dilution cooling air is jetted into the mainstream gas upstream of the vanes to reduce the mainstream gas temperature adjacent the end-wall, wherein the cooling holes are arranged in one or more circumferentially extending rows and wherein the axial position of the cooling holes in the or each row varies.

Abstract: Described is a gas turbine engine component (100), comprising a shell having an internal cavity for receiving a multi-part insert; a multi-part insert located within the cavity, wherein the multi-part insert comprises multiple separate parts assembled in an abutting relation with one another within the cavity to provide the multi-part insert; wherein the assembled insert includes at least one retention part, the retention part engaging with a wall of the cavity and at least one other insert part so as to retain the assembled insert within the cavity.

Abstract: 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.

Abstract: An aerofoil component of a gas turbine engine has an aerofoil portion which spans, in use, a working gas annulus of the engine. The aerofoil portion has a pressure side outer wall and a suction side outer wall, each extending from the leading edge to the trailing edge of the aerofoil portion. The aerofoil portion further has one or more main passages which extend in the annulus-spanning direction of the aerofoil portion and which receive, in use, a flow of coolant. The aerofoil portion further has one or more suction wall passages which extend in the annulus-spanning direction of the aerofoil portion and which receive, in use, a flow of coolant, each suction wall passage being bounded on opposing first sides by the suction side outer wall and an inner wall of the aerofoil portion, the inner wall separating the suction wall passages from the main passages.

Abstract: A component for the turbine of a gas turbine engine is provided. The component two facing walls interconnected by one or more generally elongate divider members to partially define side-by-side, generally elongate, cooling fluid passage portions which form a multi-pass cooling passage within the component. The passage portions are connected in series fluid flow relationship by respective bends formed by joined ends of neighbouring of the passage portions. The component further includes one or more core tie linking passages formed in the divider members. One or more differential pressure reducing arrangements are formed in the multi-pass cooling passage adjacent respective of the core tie linking passages.

Abstract: An internally cooled gas turbine engine component has a line of cooling air discharge holes, an internal cooling channel, an internal feed cavity for feeding cooling air from the channel to the discharge holes, and flow disrupting pedestals arranged in rows. A method of configuring the component includes: determining angles ? and ? of the directions of cooling air flow into the first and Nth rows, respectively; defining a change in angle ? of the direction of cooling air flow between rows as ?=(???)/N; and positioning the pedestals such that a line extending forward from the center of each pedestal in the ith row at an angle {?+?(i?1)} intersects the (i?1)th row at a location which is midway between two neighboring pedestals of the (i?1)th row, i being an integer from 2 to N.

Abstract: An arrangement for heating and cooling a turbine casing of a gas turbine engine, the arrangement comprising an inboard duct, adjacent to an inboard surface of the turbine casing, an outboard facing wall of the inboard duct having a plurality of impingement holes opening towards the inboard surface of the casing, through which temperature control fluid can pass from within the inboard duct to impinge upon the inboard surface of the turbine casing.

Abstract: 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.

Abstract: An end-wall component of the mainstream gas annulus of a gas turbine engine having an annular arrangement of vanes, the component including a cooling arrangement having ballistic cooling holes (33) through which, in use, dilution cooling air is jetted into the mainstream gas upstream of the vanes to reduce the mainstream gas temperature adjacent the end-wall, wherein the cooling holes are arranged in one or more circumferentially extending rows and wherein the axial position of the cooling holes in the or each row varies.

Abstract: An aerofoil component of a gas turbine engine is provided. The component has a longitudinally extending aerofoil portion which spans, in use, a working gas annulus of the engine. The aerofoil portion contains an internal chamber for a flow of coolant. The chamber includes a helical passage which spirals in a plurality of turns around an axis that extends in the length direction of the aerofoil portion.

Abstract: A cooled turbine rotor blade for a gas turbine engine is provided. The engine has an annular flow path for conducting working fluid though the engine. The blade has an aerofoil section for extending across the annular flow path. The blade further has a root portion radially inward of the aerofoil section for joining the blade to a rotor disc of the engine. The blade further has a platform between the aerofoil section and the root portion. The platform extends laterally relative to the radial direction of the engine to form an inner boundary of the annular flow path and to provide a rear overhang portion which projects in use towards a corresponding platform of a downstream nozzle guide vane. The platform contains at least one internal elongate plenum chamber for receiving cooling air. The longitudinal axis of the plenum chamber is substantially aligned with the circumferential direction of the engine.

Abstract: 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.

Abstract: A component of a turbine stage of a gas turbine engine is provided. The component forms an endwall for the working gas annulus of the stage. The component has one or more internal passages behind the endwall which, in use, carry a flow of cooling air providing convective cooling for the component at the endwall. Each passage is formed by a plurality of straight passage sections. The passage sections connect end-to-end such that the connections between nearest-neighbour passage sections form angled bends. A first portion of the passage sections lie in a first plane. A second portion of the passage sections lie in a second plane which is spaced from and parallel to the first plane. A third portion of the passage sections extend between the first and the second planes.

Abstract: An aerofoil component of a gas turbine engine has an aerofoil portion which spans, in use, a working gas annulus of the engine. The aerofoil portion has a pressure side outer wall and a suction side outer wall, each extending from the leading edge to the trailing edge of the aerofoil portion. The aerofoil portion further has one or more main passages which extend in the annulus-spanning direction of the aerofoil portion and which receive, in use, a flow of coolant. The aerofoil portion further has one or more suction wall passages which extend in the annulus-spanning direction of the aerofoil portion and which receive, in use, a flow of coolant, each suction wall passage being bounded on opposing first sides by the suction side outer wall and an inner wall of the aerofoil portion, the inner wall separating the suction wall passages from the main passages.

Abstract: A separator device is provided for separating dirt particles from a flow of cooling air fed to airfoils of the turbine section of a gas turbine engine. In use the separator device extends across a conduit which bypasses the combustor of the engine to convey pressurized cooling air carrying dirt particles from the compressor section of the engine to openings which direct the air into the airfoils. The separator device is configured to direct a first portion of the impinging cooling air flow away from the openings and to allow a second portion of the impinging cooling air to continue to the openings. The first portion of cooling air has a higher concentration of the coarsest dirt particles carried by the cooling air than the second portion of cooling air.

Abstract: A rotor blade 40 for a gas turbine engine has an aerofoil portion 42 from a root 48 to a tip 54. In use, combustion gas may leak over the tip 54 from the pressure face 52 to the suction face 50. A gutter 62 extends across the tip 54 to entrain any over tip leakage gap. The floor of the gutter defines an increased depth portion 72 at the exit end of the gutter 62.

Abstract: A turbine blade for a gas turbine engine has an aerofoil portion extending from a root to a tip. The tip carries winglets. A gutter extends across the tip to entrain gas leaking around the tip (over tip leakage). The aerofoil portion has a mean camber line and the gutter has a center line. In the examples described, the conditions that (a) the mean camber line and the centre line coincide at the exit when viewed from the tip towards the root, and (b) the mean camber line and the center line are parallel at the exit when viewed as aforesaid, are not both fulfilled.

Abstract: A cooling arrangement is provided for a turbine disc of a gas turbine engine. The turbine disc has a plurality of circumferentially spaced disc posts forming fixtures therebetween for a row of turbine blades. Each turbine blade has an attachment formation which engages at a respective fixture, a platform radially outwardly of the attachment formation such that the adjacent platforms of the row form an inner endwall for the working gas annulus of the engine, and an aerofoil which extends radially outwardly from the platform. A respective cavity is formed between an exposed radially outer surface of each disc post and the inner endwall. The cooling arrangement has at each disc post, a cooling plate located in the respective cavity and spaced radially outwardly from the exposed outer surface of the disc post to form a cooling channel between the cooling plate and the exposed outer surface.

Abstract: 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.