A gas turbine engine including an axial high pressure compressor having expansion slots in the outer rim of the rotor section. The expansion slots may be positioned between blades of a rotor segment. The fore end of the slots may have an axial seal which is coupled to the inner surface of the outer rim in the first rotor segment, and may comprise a fin configuration. The axial seal may be integral to the inner surface of the outer rim. The compressor may comprise a plurality of expansion slots and axial seals, including in a plurality of rotor segments.

Aspects of the disclosure are directed to a seal comprising: a shoe, and at least one beam coupled to the shoe, wherein the seal includes a single crystal material with a predetermined crystalline orientation. Aspects of the disclosure are directed to a method for designing a seal, comprising: obtaining a requirement associated with at least one of: a geometrical profile of the seal, a temperature range over which the seal is to operate, a natural frequency associated with the seal, or a range of deflection associated with the seal, selecting a crystalline orientation for a single crystal material of the seal based on the requirement, and fabricating the seal based on the selected crystalline orientation.

Aspects of the disclosure are directed to a seal comprising: a shoe, and at least one beam coupled to the shoe, wherein the seal includes a single crystal material with a predetermined crystalline orientation. Aspects of the disclosure are directed to a method for designing a seal, comprising: obtaining a requirement associated with at least one of: a geometrical profile of the seal, a temperature range over which the seal is to operate, a natural frequency associated with the seal, or a range of deflection associated with the seal, selecting a crystalline orientation for a single crystal material of the seal based on the requirement, and fabricating the seal based on the selected crystalline orientation.

A turbomachine blade may include an airfoil and a shank coupled to the airfoil. The shank may include a cover plate having a first circumferential face and a second, opposing circumferential face. A radial cooling groove is positioned in the first circumferential face and is configured to allow a cooling fluid to pass from a first radial position to a second, different radial position relative to the platform. The radial cooling groove provides cover plate and shank cooling. In addition, the radial cooling groove may deliver fluid for purging gaps between blade platforms and cover plates, which prevents the ingestion of hot gas from the turbine flowpath.

A turbomachine blade may include an airfoil and a shank coupled to the airfoil. The shank may include a cover plate having a first circumferential face and a second, opposing circumferential face. A radial cooling groove is positioned in the first circumferential face and is configured to allow a cooling fluid to pass from a first radial position to a second, different radial position relative to the platform. The radial cooling groove provides cover plate and shank cooling. In addition, the radial cooling groove may deliver fluid for purging gaps between blade platforms and cover plates, which prevents the ingestion of hot gas from the turbine flowpath.

A turbine blade includes an airfoil that extends radially between a root end and a tip end, a platform coupled to the root end, and a shank that extends radially inwardly from the platform. The shank includes a cover plate. The cover plate includes an outer surface, an opposite inner surface, and a contoured face that at least partially defines a damper pin slot. The contoured face extends from the outer surface to a first blend edge. The cover plate also includes a blended surface that extends from the first blend edge to a second blend edge. The second blend edge intersects with the inner surface.

A turbomachinery rotor blade is provided having an aerofoil body, and a hole penetrating the aerofoil body from a suction surface to a pressure surface thereof. The hole is suitable to receive a lacing wire. The blade further has a protrusion from the suction or pressure surface. The protrusion extends in a downstream direction from a downstream side of the hole and/or extends in an upstream direction from an upstream side of the hole, thereby disturbing the suction or pressure surface to locally thicken the aerofoil body adjacent the hole. The maximum radial extent of the protrusion in the radially outward direction of the blade is radially coterminous with the outboard side of the hole, and the maximum radial extent of the protrusion in the radially inward direction of the blade is radially coterminous with the inboard side of the hole.

A turbomachinery rotor blade is provided having an aerofoil body, and a hole penetrating the aerofoil body from a suction surface to a pressure surface thereof. The hole is suitable to receive a lacing wire. The blade further has a protrusion from the suction or pressure surface. The protrusion extends in a downstream direction from a downstream side of the hole and/or extends in an upstream direction from an upstream side of the hole, thereby disturbing the suction or pressure surface to locally thicken the aerofoil body adjacent the hole. The maximum radial extent of the protrusion in the radially outward direction of the blade is radially coterminous with the outboard side of the hole, and the maximum radial extent of the protrusion in the radially inward direction of the blade is radially coterminous with the inboard side of the hole.

A flow-control assembly for guiding a flow of fluid having a variable mass flow rate onto a turbine comprising: a turbine comprising a blade and configured to rotate about an axis of rotation; and a flow-guidance element in fluid communication with the turbine and comprising a flow-guiding vane and configured to guide a flow of fluid at a relative fluid flow angle to rotate the turbine about the axis of rotation; wherein the flow-guidance element is configured to rotate about the same axis of rotation as the turbine so as to alter the variation of the relative fluid flow angle at turbine ingress arising from varying mass flow rate in the flow of fluid.

A method of forming a metal matrix composite (MMC). The method comprises providing a fiber (26) comprising a ceramic material coated with a metal, providing a winding head (12) having a plurality of circumferentially spaced radially extending alternate first and second finger members (18, 20), the finger members each defining a winding surface (22, 24), the winding surface of each first finger member facing a first axial direction, and the winding surface of each second finger member facing a generally opposite axial direction, wherein adjacent winding surfaces (22, 24) of the first and second finger members (18, 20) are spaced in a circumferential direction, and define an axial spacing less than the diameter of the fiber (26), and winding the fiber around the winding head (12) between the winding surfaces (22, 24) of the first and second finger members (18, 20).