Abstract: A cooled turbine rotor blade for a gas turbine engine includes an airfoil and a profiled root radially inward of the airfoil, a platform between the airfoil and the root, and a cooling air supply channel extending through the platform into an interior of the airfoil and therein up to an outlet opening. An inlet opening of the cooling air supply channel is located at the rear side of the rotor blade, and an inlet part of the cooling air supply channel with the inlet opening is angled into the direction of rotation of the rotor blade and curved into a radially outward direction. Further, a rotor-stator stage for a gas turbine includes a rotor blade as above, and an air cavity radially inwards of sealing features between the rotor stage and a neighboring downstream stator stage to form a source of cooling air for the rotor blade.

Abstract: A turbine arrangement for gas turbine engine, the arrangement comprising a high pressure turbine mounted on a high pressure shaft and a low pressure turbine mounted on a low pressure shaft, wherein the high pressure turbine comprises a single stage turbine rotor configured to rotate in a first direction, and wherein the low pressure turbine comprises three or more rotors (19a, b, c) configured to rotate in a second direction opposite to the first direction. The low pressure turbine is closely coupled to the high pressure turbine with no stator vanes provided between the two.

Abstract: A turbine blade having labyrinth of internal channels for circulation of coolant, the novel labyrinth geometry includes; inlet arranged on axially upstream face of terminal portion leading to an upstream duct portion having first section adjacent the inlet and a second section, the second section having reduced cross section compared to first section; a first passage intersecting first section and extending through the blade body towards tip of the blade, a proximal end of leading edge passage arranged to capture incoming air flow; a main blade passage intersecting a downstream duct portion, the downstream duct portion arranged in axial alignment with the upstream duct portion but separate from upstream duct portion; and a restrictor passage intersecting with the mid-blade passage and extending towards a mid-blade duct portion, the mid-blade duct portion in axial alignment with the upstream and downstream duct portions and in fluid communication with the upstream duct portion.

Abstract: A turbine disc assembly comprises a turbine disc mounted coaxially on a shaft and, in use, arranged in the path of a hot work fluid flow. A cover plate is axially displaced from the turbine disc in a direction upstream with respect to the work fluid flow. The cover plate comprises a radially inner annular disc component and a radially outer annular disc component, the annular disc components are arranged coaxially with each other.

Abstract: There is disclosed a pressure responsive valve 352 for controlling a cooling flow through an orifice 320 in a gas turbine assembly 300. The valve 352 comprises an attachment point 360, a valve element 358, and a compressible valve body 354 defining a chamber 355 for sealing a volume of compressible gas. The valve body 354 is configured to act between the attachment point 360 and the valve element 358 so that in use expansion or contraction of the valve body 354 in response to external pressure causes the valve element 358 to move relative the orifice 320. A corresponding kit, method of installation and gas turbine assembly are disclosed.

Abstract: A turbine disc assembly comprises a turbine disc mounted coaxially on a shaft and, in use, arranged in the path of a hot work fluid flow. A cover plate is axially displaced from the turbine disc in a direction upstream with respect to the work fluid flow. The cover plate comprises a radially inner annular disc component and a radially outer annular disc component, the annular disc components are arranged coaxially with each other.

Abstract: A turbine blade having labyrinth of internal channels for circulation of coolant, the novel labyrinth geometry includes; inlet arranged on axially upstream face of terminal portion leading to an upstream duct portion having first section adjacent the inlet and a second section, the second section having reduced cross section compared to first section; a first passage intersecting first section and extending through the blade body towards tip of the blade, a proximal end of leading edge passage arranged to capture incoming air flow; a main blade passage intersecting a downstream duct portion, the downstream duct portion arranged in axial alignment with the upstream duct portion but separate from upstream duct portion; and a restrictor passage intersecting with the mid-blade passage and extending towards a mid-blade duct portion, the mid-blade duct portion in axial alignment with the upstream and downstream duct portions and in fluid communication with the upstream duct portion.

Abstract: A coolant feed arrangement for delivering coolant to an axially facing surface of a rotor disc configured for carrying a row of turbine blades on its radially outer surface, the coolant delivered through a conduit upstream of the rotor disc, which has an outlet arranged radially inwardly of the blades and directed at the axially facing surface, the arrangement including two or more impediments; at least one impediment including a work extractor device arranged in the conduit to extract work from the coolant on route to a last impediment arranged at or adjacent the outlet, the last impediment including a row of static nozzles configured for accelerating the flow of coolant circumferentially in the direction of rotation of the axially facing surface of the disc whereby to match the speed and direction of rotation of the axially facing surface as the coolant is delivered to the axially facing surface.

Abstract: A guide vane assembly for a turbomachine, comprising a guide vane and a liner: the guide vane comprising an aerofoil portion having a radially outer platform, wherein one of the outer platform and the liner has a hook element, the hook element comprising an opening and a circumferentially extending channel, and the other one of the outer platform and the liner comprises a retaining element having a head portion, wherein in an operational orientation the head portion located within the channel and is too wide to be withdrawn through the opening of the hook element.

Abstract: A rotor stage is provided which comprises a shank passage for separating liquid from secondary air system (SAS) air. The shank passage is formed between a radially outer surface of a disc post, a radially inner surface of one or more blade platforms, and the shanks of two neighbouring blades. In use, SAS air is supplied to the shank passage. Any liquid, such as oil, is separated from the SAS air due to the centripetal acceleration of the rotation of the rotor disc. The separated oil can then flow, or seep, past the blade platform and into the working fluid. The remainder of the SAS air passes through a hole in a sealing plate at the downstream end of the rotor stage, thereby providing SAS air with a lower liquid content for use downstream of the rotor stage.

Abstract: A cooled component for a gas turbine engine, for example a turbine rotor blade, is provided. The component has an internal cooling flow passage. Cooling air is passed through the internal cooling flow passage to remove heat from the component and thereby reduce its temperature. The cooling air is bled from the internal cooling passage after it has passed through a portion of the passage into an internal bleed flow passage. This bled air is then used in a seal. Thus, some of the cooling air that enters the internal cooling flow passage is used both to cool the component and to form a seal, for example with another component.

Abstract: A component such as a turbine blade of a gas turbine engine has a cooling arrangement comprising a cascade impingement array in which cooling air flows from a supply chamber through impingement passages in webs to an internal passageway comprising first and second limbs. The cooling air flow through the limbs, provides improved heat transfer compared with continued impingement cooling in the chordwise direction of the blade.

Abstract: A guide vane assembly for a turbomachine, comprising a guide vane and a liner: the guide vane comprising an aerofoil portion having a radially outer platform, wherein one of the outer platform and the liner has a hook element, the hook element comprising an opening and a circumferentially extending channel, and the other one of the outer platform and the liner comprises a retaining element having a head portion, wherein in an operational orientation the head portion located within the channel and is too wide to be withdrawn through the opening of the hook element.

Abstract: A component such as a turbine blade for a gas turbine engine has a cooling arrangement comprising a supply passage 18 within the blade from which extend cooling passages 28 of a first array which open at discharge openings 29, for example at a trailing edge of the blade. Cooling passages 30 of a second array extend into the blade from discharge openings 31, and intersect the passages 28. The passages 30 terminate short of the supply passage 18. As a result of this arrangement, the limiting flow area for cooling air is defined by the cooling passages 28 of the first array, and is not affected by the accuracy with which the cooling passages 30 of the second array intersect the cooling passages 28 of the first array.

Abstract: A component such as a turbine blade of a gas turbine engine has a cooling arrangement comprising a cascade impingement array in which cooling air flows from a supply chamber through impingement passages in webs to an internal passageway comprising first and second limbs. The cooling air flow through the limbs, provides improved heat transfer compared with continued impingement cooling in the chordwise direction of the blade.

Abstract: A heat transfer arrangement comprises a target member (28 or 30), and an impingement member (32). The impingement member (32) defines a fluid path (38 or 40) there through to direct fluid onto the target member. The arrangement includes a fluid directing formation (42) on the impingement member (32). The fluid path (38 or 40) extends through the fluid directory formation (42) such that the fluid path directs fluid to exit there from at an exit angle that is substantially orthogonal to the fluid directing formation.