Patent application number | Description | Published |
20080273988 | Aerofoils - An aerofoil | 11-06-2008 |
20090081024 | Turbine blade - An aerofoil for a gas turbine engine, the aerofoil comprises a leading edge and a trailing edge, pressure and suction surfaces and defines therebetween an internal passage for the flow of cooling fluid therethrough. A particle deflector means is disposed within the passage to deflect particles within a cooling fluid flow away from a region of the aerofoil susceptible to particle build up and subsequent blockage, such as a cooling passage for a shroud of a blade. | 03-26-2009 |
20090214328 | BLADES FOR GAS TURBINE ENGINES - A blade for a gas turbine engine comprises an aerofoil having a root portion, a tip portion located radially outwardly of the root portion, and leading and trailing edges extending between the root portion and the tip portion. A shroud extends transversely from the tip portion of the aerofoil and the aerofoil defines interior cooling passages which extend between the root portion and the tip portion. The aerofoil includes a wall member adjacent the trailing edge and a support structure extending from the wall member to the shroud to support the shroud. The support structure permits a flow of cooling air from a cooling passage to the trailing edge at a region proximate the tip portion of the aerofoil. Optionally, the aerofoil also includes a flow disrupting arrangement. | 08-27-2009 |
20090263235 | Damper | 10-22-2009 |
20090280011 | Blade arrangement - With regard to gas turbine engines it will be appreciated that blades are typically cooled in order to ensure that the materials from which the blades are formed remain within acceptable operational parameters. Coolant is judiciously used in order to maintain engine operational efficiency. Unfortunately with regard to rotor blades horseshoe vortices tend to increase heating towards a pressure side of a blade resulting in localised overheating. Such localised overheating may result in premature failure of the blade component. Traditionally coolant flows have been presented over a forward projection of a blade platform. In such circumstances coolant flow will not be used as efficiently as possible with regard to protecting a pressure side of a platform in a blade assembly and arrangement. By provision of a deflector element on the forward blade platform coolant flow can be proportioned either side of a leading edge of the blade. In such circumstances generally asymmetric coolant flow is provided normally biased towards the pressure side in order to enhance cooling efficiency. A suction side in an adjacent blade assembly is cooled by spent coolant and hot gas flow from the pressure side of a neighbouring blade upstream in the assembly. | 11-12-2009 |
20090317258 | Rotor blade - Cooling within aerofoils ( | 12-24-2009 |
20100040480 | Cooling arrangement - With regard to cooling turbine blades in a gas turbine engine a compromise has to be made between convective cooling within the inner cavity defining a flow path for coolant and the blow rates for developing film cooling on an outer surface of the aerofoil. By providing a chamber between the flow cavity and external apertures reconciliation between the necessary flow rates for convective cooling within the cavity defining the pathway for coolant flow within the aerofoil and the necessary coolant blowing rate for film development can be achieved. | 02-18-2010 |
20100047078 | Blade - Cooling arrangements have been provided for blades and in particular turbine blades utilising gas turbine engines. Generally for internal strength a leading passage has been separate by a solid wall from a feed passage as impingement apertures may diminish structural strength as centres for stress concentration. However, impingement apertures allow impingement jets which have improved cooling efficiency. By providing a leading passage which is divided at least into a lower section and an upper section the lower section can have a wall which is solid for structural integrity whilst an upper section has impingement apertures for greater cooling efficiency. | 02-25-2010 |
20100119377 | Cooling arrangement - Within components such as high pressure turbine blades and aerofoils in a gas turbine engine it is important to provide cooling such that these components remain within acceptable operational parameters. Typically, film cooling as well as convective cooling is utilised. Film cooling requires holes from a feed passage from which the coolant is presented upon an external surface to develop the film. The holes themselves can create cooling through convective cooling effects. In order to maximise the convective cooling effect holes are created which have an indirect path about a direct line between an inlet and an outlet for the hole. By creating an indirect path in the form of a helix or spiral which in turn may have a variable cross sectional area from the inlet to the outlet control of coolant flow can be achieved. The inlet may have a bell mouth shape whilst the hole may have a slot or elliptical cross section to achieve greater diffusion of the coolant flow in order to create an improved exit blow rate for instant film development. | 05-13-2010 |
20100124484 | Aerofoil and method for making an aerofoil - Within aerofoils, and in particular nozzle guide vane aerofoils in gas turbine engines problems can occur with regard to coolant flows from respective inlets at opposite ends of a cavity within the aerofoil. The cavity generally defines a hollow core and unless care is taken coolant flow can pass directly across the internal cavity. Previously baffle plates were inserted within the cavity to prevent such direct jetting across the cavity. Such baffle plates are subject to additional costs as well as potential unreliability problems. Baffles formed integrally with a wall within the aerofoil allow more reliability with regard to positioning as well as consistency of performance. The baffles can be perpendicular, upward or downwardly orientated or have a compound angle. | 05-20-2010 |
20100124485 | Aerofoil cooling arrangement - Within aerofoils ( | 05-20-2010 |
20100254801 | COOLED AEROFOIL FOR A GAS TURBINE ENGINE - A cooled aerofoil for a gas turbine engine has an aerofoil section with pressure and suction surfaces extending between inboard and outboard ends thereof. The aerofoil section includes first and second internal passages for carrying cooling air. The aerofoil section further includes a plurality of holes in the external surface of the aerofoil section which receive cooling air from the internal passages. The external holes are arranged such that cooling air exiting a first portion of the external holes participates in a cooling film extending from the leading edge of the aerofoil section over said pressure surface and cooling air exiting from a second portion of the external holes participates in a cooling film extending from the leading edge over said suction surface. The first portion of external holes receives cooling air from the first internal passage, and the second portion of external holes receives cooling air from the second internal passage. The first and second internal passages are supplied with cooling air from respective and separate passage entrances. Each entrance is located at either the inboard end or the outboard end of the aerofoil section. | 10-07-2010 |
20100284807 | BLADE COOLING - Cooling of turbine blades within a gas turbine engine is important. Coolant flows are taken from the engine to provide cooling effects but diminish the efficiency of the engine. Blades rotate and therefore centrifugal effects stimulate flow and pressure to maintain coolant flow presentation upon the blade. More cooling effectiveness is required towards the root of a blade in comparison with the tip. By providing cavities which incorporate return apertures coolant flow can be recycled. The cavities incorporate return portions on one side of a feed passage and a constriction is provided in passage. Thus, a proportion of coolant within the cavities is returned to the passage with pressure maintained by the rotational and centrifugal effects upon the coolant flow through the feed passage. Coolant flow is presented through outlet apertures as a film upon a surface of a blade. | 11-11-2010 |
20100303635 | COOLING ARRANGEMENTS - Providing cooling within hollow blades such as high pressure turbine blades in a gas turbine engine is important to maintain these components within operational margins for the materials from which they are formed. Traditionally, coolant flows in hollow passages have been used along with impingement apertures towards a leading passage for cooling effectiveness. It is known that opposed undulations or ribs can create rotational vortices within the passage. By shaping shaped portions between the opposed undulations and possibly providing undulations upon these shaped portions themselves it is possible to generate stronger more powerful vortices within the passage. These vortices are coupled with the impingement orifices to create proportionally greater impingement jet flow and pressure and therefore cooling effectiveness within the leading passage. | 12-02-2010 |
20110067378 | SEPARATOR DEVICE - 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 pressurised 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. | 03-24-2011 |
20120027576 | TURBINE STAGE SHROUD SEGMENT - 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. | 02-02-2012 |
20120057961 | TURBINE STAGE SHROUD SEGMENT - 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 mainstream flow of the working gas passes through the passages formed between adjacent turbine blades. 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 has swirl directions which are co-directionally aligned with the swirl directions of the mainstream flow at the segment. | 03-08-2012 |
20120076645 | ENDWALL COMPONENT FOR A TURBINE STAGE OF A GAS TURBINE ENGINE - 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. | 03-29-2012 |
20120082567 | COOLED ROTOR BLADE - 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. The plenum chamber supplies the cooling air to a plurality of exit holes formed in the external surface of the rear overhang portion to cool that portion. | 04-05-2012 |
20120082568 | TURBINE DISC COOLING ARRANGEMENT - 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. | 04-05-2012 |
20120219401 | ENDWALL COMPONENT FOR A TURBINE STAGE OF A GAS TURBINE ENGINE - 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. | 08-30-2012 |
20120251295 | GAS TURBINE ENGINE COMPONENT - A component of a gas turbine engine is provided. The component includes an external wall which, in use, is exposed on one surface thereof to working gas flowing through the engine. The component further includes effusion cooling holes formed in the external wall. In use, cooling air blows through the cooling holes to form a cooling film on the surface of the external wall exposed to the working gas. The component further includes an air inlet arrangement which receives the cooling air for distribution to the cooling holes. The component further includes a plurality of metering feeds and a plurality of supply plena. The metering feeds meter the cooling air from the air inlet arrangement to respective of the supply plena, which in turn supply the metered cooling air to respective portions of the cooling holes. | 10-04-2012 |
20130017060 | TIP CLEARANCE CONTROL FOR TURBINE BLADESAANM BOSWELL; John H.AACI DerbyAACO GBAAGP BOSWELL; John H. Derby GBAANM TIBBOTT; IanAACI LichfieldAACO GBAAGP TIBBOTT; Ian Lichfield GB - 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. | 01-17-2013 |
20130343872 | COOLED COMPONENT FOR THE TURBINE OF A GAS TURBINE ENGINE - 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. | 12-26-2013 |
20140086724 | GAS TURBINE ENGINE COMPONENT - 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: | 03-27-2014 |
20140093379 | GAS TURBINE ENGINE COMPONENT - A gas turbine engine component is described which ( | 04-03-2014 |
20140093392 | GAS TURBINE ENGINE COMPONENT - Described is a gas turbine engine component ( | 04-03-2014 |
20140140860 | AEROFOIL COOLING - 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. | 05-22-2014 |
20150044029 | AEROFOIL - 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. | 02-12-2015 |