A cooling structure for a gas turbine engine comprises a gas turbine engine structure defining a cooling cavity. A cooling component is configured to direct cooling flow in a desired direction into the cooling cavity. A bracket supports the cooling component and has an attachment interface to fix the bracket to the gas turbine engine structure. A first orientation feature associated with the bracket. A second orientation feature is associated with the gas turbine engine structure. The first and second orientation features cooperate with each other to ensure that the cooling component is only installed in one orientation relative to the gas turbine engine structure. A gas turbine engine and a method of installing a cooling structure are also disclosed.

A rotor disk assembly for a gas turbine engine. A multiple of blade attachments each received within a slot in a rotor disk. The rotor blade attachment manufactured of a metal alloy and a rotor blade retained between each two of the multiple of blade attachments, the rotor blade manufactured of non-metal alloy.

An airfoil of either of a turbine blade or a turbine vane includes a cooling passage; at least one disk body disposed on an inner wall of the cooling passage and configured to reduce a flow cross-sectional area of the cooling passage to increase a fluid pressure of cooling fluid flowing through the cooling passage; and at least one through-hole formed in each of the at least one disk body such that the cooling fluid flows through the at least one through-hole and forms a vortex on a downstream side of the at least one through-hole. The cooling passage includes an inlet supplied with the cooling fluid and an end opposite to the inlet, and the at least one disk body is disposed at the inlet of the cooling passage and is configured to increase the fluid pressure of the cooling fluid flowing into the cooling passage.

A turbine blade having a base and an airfoil, the base including a root end. The airfoil including a skin extending from the base and defining a leading edge, a trailing edge, having a tip end opposite from the root end. The turbine blade further including dividers located within the airfoil. The dividers, leading edge, trailing edge, and skin define serpentine channels within the airfoil. A first multi-bend heat exchange path and a second multi-bend heat exchange path extend through the serpentine channels to cool portion of the turbine blade.

Turbine wheels, turbine engines, and methods of forming the turbine wheels are provided herein. In an embodiment, a turbine wheel includes a rotor disk and a plurality of turbine blades. Each turbine blade is operatively connected to the rotor disk through a blade mount, which is bonded to the rotor disk. The blade mount and the rotor disk have a fore surface on a higher pressure side thereof and an aft surface on a lower pressure side thereof. The blade mount includes a blade attachment surface that extends between and connects the fore surface and the aft surface. The turbine blade extends from the blade attachment surface. A gap is defined between adjacent blade mounts. The gap separates the blade mounts and extends into the rotor disk. The gap includes a pocket that has a fore opening in the fore surface. A pocket seal is disposed in the pocket.

An on-board injector that delivers discharge air toward a turbine rotor of a gas turbine engine includes a second wall spaced form a first wall to define an annular inlet about an engine longitudinal axis and a multiple of airfoil shapes between the first wall and the second wall to segregate discharge air from the annular inlet, and a multiple of bypass apertures each along a radial axis transverse to the engine longitudinal axis through each of the multiple of airfoil shapes and the respective first wall, the second wall.

The present invention relates to a structure for cooling a turbomachine's rotor part and a rotor and a turbomachine having the same. The structure for cooling a turbomachine's rotor part includes: a dovetail joint part disposed along an outer circumferential surface of a rotor wheel and having a plurality of mounting grooves in which dovetails of buckets are mounted and cooling slots disposed along the outer circumferential surface of the rotor wheel on the dovetail joint part and having a cooling fluid flowing therethrough.

A shroud segment arranged radially outward of a gas flow path of a gas turbine. The shroud segment includes a body having walls defining an internal pocket for receiving a supply of air. A plurality of pressurization apertures is formed through one of the walls to fluidly connect an internal pocket of the body to an ambient area of the body. A seal slot section is formed in the wall at a position radially inward of the pressurization apertures to receive a seal to connect the shroud segment to an adjacent shroud segment. The pressurization apertures are arranged such that portions of the supply of air are configured to pass through the pressurization apertures and through the seal slot section as leakage into the gas flow path, thereby reducing ingestion of fluid from the gas flow path into the internal pocket of the shroud segment.

A cooling system configured to cool aspects of the turbine engine between a compressor and a turbine assembly is disclosed. In at least one embodiment, the cooling system may include one or more mid-frame cooling channels extending from an inlet through one or more mid-frame torque discs positioned downstream of the compressor and upstream of the turbine assembly. The inlet may be positioned to receive compressor bleed air. The mid-frame cooling channel may be positioned in a radially outer portion of the mid-frame torque disc to provide cooling to outer aspects of the mid-frame torque disc such that conventional, low cost materials may be used to form the mid-frame torque disc rather than high cost materials with capacity to withstand higher temperatures. The cooling fluid routed through the mid-frame cooling channel in the mid-frame torque disc may be exhausted into a cooling system (10) for the downstream turbine assembly.

A turbine blade having a base and an airfoil, the base including cooling air inlets and an internal cooling air passageway, and the airfoil including an internal multi-bend heat exchange path beginning at the base and ending at a cooling air outlet at the trailing edge of the airfoil. The airfoil also includes a skin that encompasses a tip wall, an inner spar, and a tip flag cooling system.