A shroud attaching structure of a gas turbine includes: a support provided around an axis of the gas turbine and inside a casing of the gas turbine in a radial direction; and a shroud assembly attached to the support so as to cover an inner peripheral surface of the support, the shroud assembly being formed by laminating a large number of plate-shaped shroud elements containing a ceramic matrix composite. The shroud elements are lined up in a circumferential direction of the support. The adjacent shroud elements are arranged so as to be slidable on each other.

A gas turbine engine having a turbine blade tip clearance control system for increasing the efficiency of the engine by reducing the gap between turbine blade tips and radially outward ring segments is disclosed. The turbine blade tip clearance control system may include one or more clearance control bands positioned radially outward of inner surfaces of ring segments and bearing against at least one outer surface of the ring segments to limit radial movement of the ring segments. During operation, the clearance control band limits radial movement of the ring segments, and the turbine blade tips do not have a pinch point during start-up transient conditions. In addition, the smallest gap during turbine engine operation may be found at steady state operation of the gas turbine engine. Thus, the clearance control system can set the gap between turbine blade tips and ring segments to be zero at steady state operation.

A gas turbine engine having a turbine blade tip clearance control system for increasing the efficiency of the engine by reducing the gap between turbine blade tips and radially outward ring segments is disclosed. The turbine blade tip clearance control system may include one or more clearance control bands positioned radially outward of inner surfaces of ring segments and bearing against at least one outer surface of the ring segments to limit radial movement of the ring segments. During operation, the clearance control band limits radial movement of the ring segments, and the turbine blade tips do not have a pinch point during start-up transient conditions. In addition, the smallest gap during turbine engine operation may be found at steady state operation of the gas turbine engine. Thus, the clearance control system can set the gap between turbine blade tips and ring segments to be zero at steady state operation.

Casings and methods for manufacturing casings are provided. For example, a casing defining radial, axial, and circumferential directions is provided. The casing comprises an annular inner wall and an annular outer wall, each extending along the axial direction, with the outer wall radially spaced apart from the inner wall. The casing also comprises an auxetic structure extending from the inner wall to the outer wall and including a plurality of lattice elements. Each lattice element extends circumferentially and radially from the inner to the outer wall, and the lattice elements are axially spaced apart from one another. The auxetic structure may define at least one aperture for fluid flow from one portion to another of the auxetic structure and/or may be configured to vary the thermal characteristics of the casing along the axial direction. The casing may be integrally formed as a single monolithic component, e.g., by additive manufacturing.

Casings and methods for manufacturing casings are provided. For example, a casing defining radial, axial, and circumferential directions is provided. The casing comprises an annular inner wall and an annular outer wall, each extending along the axial direction, with the outer wall radially spaced apart from the inner wall. The casing also comprises an auxetic structure extending from the inner wall to the outer wall and including a plurality of lattice elements. Each lattice element extends circumferentially and radially from the inner to the outer wall, and the lattice elements are axially spaced apart from one another. The auxetic structure may define at least one aperture for fluid flow from one portion to another of the auxetic structure and/or may be configured to vary the thermal characteristics of the casing along the axial direction. The casing may be integrally formed as a single monolithic component, e.g., by additive manufacturing.

A blade outer air seal (BOAS) for a gas turbine engine includes a seal ring body having a radially inner face and a radially outer face that axially extend between a leading edge portion and a trailing edge portion and a segmented spline that extends from the radially outer face of the seal ring body, the seal secured to the radially inner face of the seal ring body.

A rotor, installable in a casing of a turbine and configured to be rotated by a flow of combustion gas and cooled by a flow of compressed air, has an improved structure to keep a tip clearance constant during operation of a gas turbine. The rotor includes a disk having an outer circumferential surface; a platform installed on the outer circumferential surface of the disk; and a blade airfoil formed on an upper surface of the platform, the blade airfoil having an airfoil end situated opposite to the platform, the airfoil end having an upstream side and a downstream side with respect to a flow direction of the combustion gas, wherein the blade airfoil is formed so that, when the rotor is installed in the casing, the downstream side of the airfoil end is closer to an inner surface of the casing than the upstream side of the airfoil end.

A rotor, installable in a casing of a turbine and configured to be rotated by a flow of combustion gas and cooled by a flow of compressed air, has an improved structure to keep a tip clearance constant during operation of a gas turbine. The rotor includes a disk having an outer circumferential surface; a platform installed on the outer circumferential surface of the disk; and a blade airfoil formed on an upper surface of the platform, the blade airfoil having an airfoil end situated opposite to the platform, the airfoil end having an upstream side and a downstream side with respect to a flow direction of the combustion gas, wherein the blade airfoil is formed so that, when the rotor is installed in the casing, the downstream side of the airfoil end is closer to an inner surface of the casing than the upstream side of the airfoil end.

A control ring for use in a gas turbine engine includes a control ring segment defining a centerline axis. The control ring segment includes an inner diameter surface and an outer diameter surface. A thermally isolating contact is operatively connected to at least one of the inner diameter surface and the outer diameter surface. The thermally isolating contact has lower thermal conductivity than the control ring.

A control ring for use in a gas turbine engine includes a control ring segment defining a centerline axis. The control ring segment includes an inner diameter surface and an outer diameter surface. A thermally isolating contact is operatively connected to at least one of the inner diameter surface and the outer diameter surface. The thermally isolating contact has lower thermal conductivity than the control ring.