A cooling device for cooling platforms of a guide vane ring of a gas turbine is arranged downstream inside a main flow channel of a combustion chamber. Cooling air passages are arranged in a wall of the platforms or of an intermediate piece that is connected therewith to guide cooling air for film cooling the surfaces of the platforms. At least in certain areas, the wall is configured with at least two layers having—as viewed from the main flow channel—an outer wall and a spaced apart inner wall forming a hollow space, wherein the hollow space can be impinged by cooling air through at least one cooling air blow-in opening inside the outer wall, and at least one cooling air blow-out opening is arranged inside the inner wall extending in the downstream direction to the surfaces of the platforms.

The turbomachine includes a compressor, an inner annular casing, and an outer annular casing. The inner annular casing and the outer annular casing define at least one cavity therebetween. The clearance control system includes a manifold system including at least one conduit disposed within the cavities and configured to channel a flow of cooling fluid between the cavities. The clearance control system also includes an impingement system including a header and at least one plenum configured to channel the flow of cooling fluid to the inner annular casing. The conduits configured to channel the flow of cooling fluid to the impingement system. The clearance control system further includes a channel system including at least one channels configured to channel the flow of cooling fluid to the turbomachine. The channels are configured to control the flow of cooling fluid to the manifold system.

A blade outer air seal support includes, at least one arc body having a first portion and a second portion, a blade outer air seal mounting region defined at least partially between the first portion and the second portion, and an interface feature interfacing the first portion and the second portion. The interface feature is configured such that axially aligned forces are communicated between the first and second portions through the interface feature, bypassing the blade outer air seal mounting region.

A ceramic matrix composite article, method for forming the article, and perforated ply which may be incorporated therein are disclosed. The article includes at least one shell ply forming an exterior wall having first and second portions and defining a plenum. An annular brace formed of at least one structural support ply is disposed within the plenum, including a first integral portion integral with and part of the first portion of the exterior wall, a first curved portion extending from the first integral portion and curving across the article plenum to the second portion of the exterior wall, a second integral portion integral with and part of the second portion of the exterior wall, a second curved portion extending from the second integral portion and curving across the article plenum to the first curved portion, and an overlap in which the first and second curved portions are integral.

An impingement manifold includes a fluid inlet passage and a pressurized chamber. The pressurized chamber includes at least one lobe. The at least one lobe includes a flow improving feature configured to minimize vorticity of a flow field within the pressurized chamber, and at least one flow outlet.

A heat shield panel for a gas turbine engine combustor is disclosed. The heat shield panel includes a hot side defining a first surface having an outer perimeter, a cold side defining a second surface spaced from the first surface and a plurality of holes, each hole including a central axis having vector components defined by a common vector.

A turbine rotor blade includes a mounting portion that partially defines a cooling circuit within the turbine rotor blade and an airfoil portion that extends radially outward from the mounting portion. The airfoil portion further defines the cooling circuit. The turbine rotor blade further includes a platform portion that is disposed radially between the mounting portion and the airfoil. The platform portion includes a bottom wall, a top wall, a forward wall, an aft wall and a pair of opposing side walls. A cooling plenum that at least partially defines the cooling circuit is defined within the platform portion. The cooling plenum is at least partially defined between the forward wall, the aft wall and between the pair of opposing side walls.

A blade outer air seal assembly includes a blade outer air seal that has a plurality of segments that extend circumferentially about an axis and are mounted in a support structure via a carrier. At least one of the segments have a first hook circumferentially spaced from a second hook. A base portion extends from the first hook to the second hook. The carrier has a cavity on a radially inner surface between the carrier and the base portion.

A system may be provided that includes a hot section component of a gas turbine engine. The hot section component includes a dual wall, which includes a first wall and a second wall. The first wall includes multiple impingement cooling holes extending through the first wall. The second wall is positioned adjacent the first wall. The first wall and the second wall together define a cooling passage between the first wall and the second wall. Multiple pockets are in a surface of the second wall. Each of the pockets is positioned opposite a respective one of the impingement cooling holes. Each of the pockets is configured to receive a cooling fluid from the respective one of the impingement cooling holes and direct the cooling fluid into the cooling passage. The cooling passage includes a single cooling passage into which the pockets are configured to direct the cooling fluid.

A turbine rotor blade is additively manufactured and includes an airfoil body including a concave pressure side outer wall and a convex suction side outer wall that connect along leading and trailing edges. A shank is at a radial inner end of the airfoil body, and at least one angel wing extends laterally from at least one side of the shank. A coolant transfer passage is defined through the at least one angel wing. The coolant transfer passage fluidly couples a first wheel space portion defined between the shank and a first adjacent shank of a first adjacent turbine rotor blade and a second wheel space portion defined between the shank and a second adjacent shank of a second adjacent turbine rotor blade. The coolant transfer passage allows coolant to pass between wheel space portions of adjacent turbine rotor blades.