Abstract: A component for a gas turbine engine, including: at least one internal channel extending through a first portion of the component, the at least one internal channel having at least one inlet opening and at least one outlet opening each being in fluid communication with the at least one internal channel; a plurality of cooling features extending from a surface of the at least one internal channel; and at least one internal channel extending through a second portion of the component, the at least one internal channel extending through the second portion of the component is in fluid communication with the at least one internal channel extending through the first portion of the component, the second portion being located on top of the first portion, the at least one internal channel extending through the second portion of the component having a plurality of openings extending from the least one internal channel extending through the second portion of the component through the plurality of cooling features to an o

Abstract: A component for a gas turbine engine, including: at least one internal cavity extending through the component, the internal cavity having at least one inlet opening and at least one outlet opening each being in fluid communication with the at least one internal cavity; a plurality of cooling features extending from a surface of the at least one internal cavity, the plurality of cooling features are formed in accordance with at least one of the following groups: i) a plurality of airfoil shaped features that extend upwardly from the surface of the at least one internal cavity and a plurality of wedge shaped features each having a triangular base that has an upstream portion and a downstream portion, the upstream portion extending further from the surface than the downstream portion; ii) a plurality features having a curved or “J” shaped base that extends upwardly from the surface, a plurality features having a double curved or symmetrically “J” shaped base that extends upwardly from the surface, and a pluralit

Abstract: A blade outer air seal includes a blade outer air seal body having a radially inward facing web to be centered on a rotational axis of an associated engine. The blade outer air seal has a leading edge and a trailing edge and there is a leading edge mount hook and a trailing edge mount hook. A cooling air inlet in the leading edge mount hook communicates with a source of cooling air. The web is intermediate the leading edge mount hook and the trailing edge mount hook and the cooling air inlet communicating with a cooling air circuit in the web. Air passes from the leading edge toward the trailing edge and then to an outlet extending radially outwardly and into a central chamber defined between the leading edge mount hook and the trailing edge mount hook. The outlet is upstream of the trailing edge mount hook.

Abstract: Impingement assemblies and components of gas turbine engines are described. The impingement assemblies include an impingement plate having a plurality of impingement holes formed therein and an impingement surface arranged relative to the impingement plate with an impingement cavity defined between the impingement plate and the impingement surface. A raised wall is configured on the impingement surface and extends in a direction from the impingement surface toward the impingement plate and defines a plurality of impingement cells each having a geometric shape on the impingement surface.

Abstract: Impingement assemblies and components of gas turbine engines are described. The impingement assemblies include an impingement plate having a plurality of impingement holes formed therein and an impingement surface arranged relative to the impingement plate with an impingement cavity defined between the impingement plate and the impingement surface. A raised wall is configured on the impingement surface and extends in a direction from the impingement surface toward the impingement plate and defines a plurality of impingement cells each having a geometric shape on the impingement surface.

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. A sacrificial structure for forming internal cooling passages within a blade outer air seal and a method of cooling an interface between a blade outer air seal and a rotating blade array is also disclosed.

Abstract: An impingement plate for providing cooling to a gas turbine engine component includes a first surface with a second surface opposite the first surface. The first surface and the second surface are surrounded by a perimeter defining an edge. A first plurality of holes extend through the impingement plate. A plurality of protrusions extend outward from one of the first surface or the second surface and have a distal surface transverse to one of the first surface or the second surface. At least one second impingement hole extends through the distal surface.

Abstract: A cooling system includes a first surface, a second surface distant from the first surface, and a plenum formed between the first surface and the second surface. The second surface includes a plurality of impingement holes extending through the second surface and configured to provide cooling fluid to the plenum and first surface with the plurality of impingement holes each being angled relative to a line perpendicular to a tangent line corresponding to each of the plurality of impingement holes. The first surface and the second surface can be annular in shape with the first surface radially inward from the second surface or the second surface radially inward from the first surface.

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. A sacrificial structure for forming internal cooling passages within a blade outer air seal and a method of cooling an interface between a blade outer air seal and a rotating blade array is also disclosed.

Abstract: An impingement plate for providing cooling to a gas turbine engine component includes a first surface with a second surface opposite the first surface. The first surface and the second surface are surrounded by a perimeter defining an edge. A first plurality of holes extend through the impingement plate. A plurality of protrusions extend outward from one of the first surface or the second surface and have a distal surface transverse to one of the first surface or the second surface. At least one second impingement hole extends through the distal surface.

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 blade outer air seal for a gas turbine engine includes a seal arc-segment that defines a gaspath side, a non-gaspath side, leading and trailing ends, and first and second circumferential sides. A portion of the gaspath side has a geometrically segmented coating section. The geometrically segmented coating section includes a wall that has an array of cells, and a coating disposed in the array of cells.

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 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: 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: 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: 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 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 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: An blade outer air seal support assembly includes a main support member configured to support a blade outer air seal. The main support member extends generally axially between a leading edge portion and a trailing edge portion. The leading edge portion is configured to be slidably received within a groove established by the blade outer air seal. A support tab extends radially inward from the main support member toward the blade outer air seal. The support tab configured to contact an extension of the blade outer air seal to limit relative axial movement of the blade outer air seal. A gusset spans between the support tab and the main support member.