Patent application number | Description | Published |
20090136358 | Blade cooling - In order to couple coolant air flow presented through a coolant gallery | 05-28-2009 |
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 |
20100054955 | Blades - A rotary blade, such as a turbine blade for a gas turbine engine, has an aerofoil portion with a tip partly shrouded by winglets. A gutter extends across the radially outer face of the tip to leave upstands. Cooling air feed galleries are drilled into each upstand, from the trailing edge, toward the upper end of a cooling air feed void, which is spaced from the trailing edge. Cooling passages are drilled from the winglet edges to the gallery. Cooling air supplied through the void passes along the gallery, through the passages and leaves the blade at the cooling holes. This allows cooling to be provided near the trailing edge of the tip without requiring the geometry around the trailing edge to be thickened to accommodate a cooling air void. | 03-04-2010 |
20100061854 | Turbine blade damper arrangement - A turbine blade damper arrangement in which a damper is positioned against the undersides of the platforms of adjacent turbine blades. In operation, the damper is centrifugally urged into engagement with the blade platforms to provide damping of relative movement between the blades. The damper and platform surfaces that it engages are of part-cylindrical configuration in order to minimise gas leakage paths between the damper and blade platforms. | 03-11-2010 |
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 |
20100316486 | COOLED COMPONENT FOR A GAS TURBINE ENGINE - There is disclosed a cooled component for a gas turbine engine, the component preferably taking the form of a shrouded turbine blade, and having a segment region defining a segment of an annulus for the passage of hot gases therethrough. The segment region has a pair of opposed side faces configured to lie substantially adjacent respective corresponding side faces of the segments of similar operationally and circumferentially adjacent components when a series of such components are mounted in an engine such that their respective segments define an annulus. The component of the present invention is characterised by the provision of an elongate cooling slot in at least one of said side faces, said cooling slot being arranged in fluid communication with at least one flow passage within said segment region for the supply of cooling fluid to said slot, the slot being substantially closed at its upstream end and open at its downstream end so as to define an outlet for said cooling fluid at the operationally downstream region of said side face. | 12-16-2010 |
20110255985 | BLADES - A rotor blade | 10-20-2011 |
20110255986 | BLADES - A turbine blade ( | 10-20-2011 |
20110255990 | BLADES - A rotor blade for a gas turbine engine has an aerofoil portion and a tip region. The tip region is at the radially outermost end of the blade. The radially outermost surface carries abrasive material (not shown) to interact with an abradable surface. The tip has a recess in which cooling air outlets are formed. The recess is open in a circumferential direction. This allows cooling air outlets to be formed without interference from the abrasive material, and inhibits any tendency for abrasion debris to collect in the recess and interfere with the flow of cooling air. | 10-20-2011 |