| 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 |
