Abstract: A ring-shaped compliant support for a gas turbine engine includes, among other things, an annular case, and an adjustment member that will turn relative to the annular case if exposed to thermal energy.

Abstract: A cooling chamber in a gas turbine engine includes a first side surface, a second side surface, a bottom surface, and a top surface defining a chamber therein. The second side surface is angled at a first angle with respect to the first side surface, the chamber having an inlet end and an exit located downstream of the inlet end, wherein the chamber has a width that narrows from the inlet end toward the exit. An inlet is located in one of the top surface or the bottom surface at the inlet end of the chamber. At least one divider is located within the chamber, the at least one divider configured to separate an airflow flowing from the inlet to the exit into a first airflow and a second airflow. The at least one divider is angled at a second angle with respect to the first side surface.

Abstract: A turbine component for a gas turbine engine includes a multiple of self-opening cooling passages each of which defines a self-opening cooling passage axis that extends through a gas path surface.

Abstract: A gas turbine engine component is provided. The gas turbine engine component has an internal cooling passage having first endwall and a second endwall. A primary inlet aperture is in fluid communication with a coolant supply and the first end of the internal cooling passage. An elongated rib is inside the internal cooling passage and proximate the primary inlet aperture. The elongated rib is configured to direct at least a portion of the coolant from the primary inlet aperture against a proximate one of the first or second endwall and then substantially reverse direction to flow downstream toward a second end of the internal cooling passage. A blade outer air seal (“BOAS”) assembly and a BOAS segment thereof are also provided.

Abstract: Impingement assemblies including an impingement plate having an increased cross-flow structure that forms an increased cross-flow area between the impingement plate and an impingement surface when the impingement plate is installed proximate to the impingement surface. The increased cross-flow structure has a first portion having at least one impingement hole passing through the first portion, the first portion being separated from the impingement surface by an impingement hole height, a second portion separated from the first portion by a separation distance and separated from the impingement surface by a cross-flow structure height, and a third portion extending between the first portion and the second portion. The first portion, the second portion, and the third portion define at least one cross-flow cavity between the impingement plate and the impingement surface, and the cross-flow structure height is greater than the impingement hole height.

Abstract: A probe cooling system for a gas turbine engine includes a probe housing and a bushing disposed between an end of the probe housing and a gas path, the bushing extending from the end of the probe housing to form a cavity. The system also includes a first plenum defined within the probe housing and configured to direct air from an internal cooling air supply towards the cavity. The system also includes a second plenum defined between the bushing and the probe housing and configured to direct air from the internal cooling air supply towards the cavity.

Abstract: A blade outer air seal (BOAS) includes a body defining a core having a forward cavity portion and a plurality of forward hooks extending radially from the body, the hooks extending to a hook height. The BOAS includes a seal surface defined between the forward hooks and which includes a forward side, an aft side, and a seal surface height. The seal surface is aft of the forward cavity portion. The BOAS includes a thermal regulation channel defining an inlet aft of the sealing surface and an outlet defined in fluid communication with the forward cavity portion.

Abstract: According to various embodiments, disclosed is a blade outer air seal assembly for a turbine engine comprising a main body portion that extends generally axially, with respect to a central axis of the turbine engine, from a leading edge portion of the main body to a trailing edge portion of the main body, wherein the leading edge portion includes a leading edge wall having a undercut profile along at least a portion of the leading edge wall, wherein the main body portion comprises cooling passages comprising a leading edge cooling passage adjacent to the leading edge wall, having a leading edge periphery on a side of the leading edge cooling passage adjacent to the leading edge wall, which generally conforms to the undercut profile of the leading edge wall.

Abstract: A cooling chamber in a gas turbine engine includes a first side surface, a second side surface, a bottom surface, and a top surface defining a chamber therein. The second side surface is angled at a first angle with respect to the first side surface, the chamber having an inlet end and an exit located downstream of the inlet end, wherein the chamber has a width that narrows from the inlet end toward the exit. An inlet is located in one of the top surface or the bottom surface at the inlet end of the chamber. At least one divider is located within the chamber, the at least one divider configured to separate an airflow flowing from the inlet to the exit into a first airflow and a second airflow. The at least one divider is angled at a second angle with respect to the first side surface.

Abstract: A gas turbine engine component is provided. The gas turbine engine component has an internal cooling passage having first endwall and a second endwall. A primary inlet aperture is in fluid communication with a coolant supply and the first end of the internal cooling passage. An elongated rib is inside the internal cooling passage and proximate the primary inlet aperture. The elongated rib is configured to direct at least a portion of the coolant from the primary inlet aperture against a proximate one of the first or second endwall and then substantially reverse direction to flow downstream toward a second end of the internal cooling passage. A blade outer air seal (“BOAS”) assembly and a BOAS segment thereof are also provided.

Abstract: A gas turbine engine component includes a wall portion and a leading edge cooling channel that extends through the wall portion. The leading edge cooling channel includes at least one first cooling passage separated from at least one serpentine cooling passage.

Abstract: A blade outer air seal according to an exemplary aspect of the present disclosure includes, among other things, a body to be distributed circumferentially about a blade array. The body has a plurality of grooves, which can, for example, improve the aerodynamic efficiency of a turbine. A fin is between a first groove and a second groove of the plurality of grooves. The fin extends radially from the body and terminates at a radially inner fin face that provides one or more cooling outlets.

Abstract: A blade outer air seal (BOAS) includes a body defining a core having a forward cavity portion and a plurality of forward hooks extending radially from the body, the hooks extending to a hook height. The BOAS includes a seal surface defined between the forward hooks and which includes a forward side, an aft side, and a seal surface height. The seal surface is aft of the forward cavity portion. The BOAS includes a thermal regulation channel defining an inlet aft of the sealing surface and an outlet defined in fluid communication with the forward cavity portion.

Abstract: A gas turbine engine component includes a structure including a first wall and a second wall that provide a cooling passage. The cooling passage extends a length from a first end to a second end. A cluster of impingement inlet holes is provided in the second wall at the first end. An outlet is provided at the second end.

Abstract: One exemplary embodiment of this disclosure relates to a gas turbine engine including a component having a body. The body includes a tortuous cooling passageway, which provides a flow path extending between an inlet in a first surface of the body and an exit in a second surface of the body.

Abstract: A probe cooling system for a gas turbine engine includes a probe housing and a bushing disposed between an end of the probe housing and a gas path, the bushing extending from the end of the probe housing to form a cavity. The system also includes a first plenum defined within the probe housing and configured to direct air from an internal cooling air supply towards the cavity. The system also includes a second plenum defined between the bushing and the probe housing and configured to direct air from the internal cooling air supply towards the cavity.

Abstract: A turbine component for a gas turbine engine includes a multiple of self-opening cooling passages each of which defines a self-opening cooling passage axis that extends through a gas path surface.

Abstract: A turbine component includes a main body with an upstream end and a downstream end. A cooling passage network is interior to the main body and has multiple cooling passages. A shared passage wall in the cooling passage network includes a cross passage opening connecting each passage partially defined by the shared passage wall. The cross passage further including a pressure balancing feature operable to reduce a local pressure differential in a fluid pressure in each adjacent passage at the cross passage.

Abstract: A ring-shaped compliant support for a gas turbine engine includes, among other things, an annular case, and an adjustment member that will turn relative to the annular case if exposed to thermal energy.

Abstract: A multi impingement plate assembly for a Blade Outer Air Seal (BOAS) includes a first impingement plate which defines a multiple of first impingement plate holes and a second impingement plate attached to the first impingement plate. The second impingement plate includes a platform section spaced away from the multiple of first impingement plate holes to define a plate cavity.