Abstract: A component for a gas turbine engine includes at least one airfoil that has a radially inner end and a radially outer end. A platform includes a radially outer surface that is attached to the radially inner end of the airfoil. A radially inner side of the platform includes a forward surface and an aft surface. A pocket is recessed into at least one of the forward surface and the aft surface. A cover plate covers the pocket and a seal is attached to the cover plate.

Abstract: An example case for a gas turbine engine includes a bearing compartment, a radially outer ring, and a radially inner ring. The radially outer ring and radially inner ring are connected by a plurality of circumferentially spaced apart struts and define an annular flow path therebetween. The radially inner ring includes a first flange that extends radially outward from the bearing compartment. The first flange includes an undulation that extends axially away from the plurality of struts, and defines a load path between the bearing compartment and the radially outer ring.

Abstract: A component for a gas turbine engine includes a surface adjacent a flow of hot gases. A plurality of cavities is in a portion of the surface. The plurality of cavities have a first group of cavities with a first cross-section and a second group of cavities with a second cross-section different from the first cross-section. The first and second groups of cavities are arranged such that there is no straight line across the portion of the surface that does not intersect one of the plurality of cavities. A thermal barrier coating is over the surface and fills each of the plurality of cavities.

Abstract: A section for a gas turbine engine includes a rotating structure, a stationary structure, and a flow guide assembly arranged generally between the rotating structure and the stationary structure. A flow path is defined between the flow guide assembly and one of the rotating structure and the stationary structure. The flow guide assembly includes a plurality of apertures configured to disrupt acoustic waves of air in the flow path. A seal is configured to establish a sealing relationship between the rotating structure and the stationary structure, and wherein an inlet to the flow path is adjacent the seal. A gas turbine engine and a method of disrupting acoustic waves in a flow path of a gas turbine engine are also disclosed.

Abstract: A component for a gas turbine engine is disclosed. The component includes a platform and a rail having an outer radial surface and an inner radial surface, with the inner radial surface connected to the platform. The rail includes a plurality of apertures spaced circumferentially between the outer radial surface and the inner radial surface, the apertures disposed to form a truss-like structure.

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: Aspects of the disclosure are directed to a floating, non-contact seal comprising: a shoe, and at least three beams, each beam having a first axial end and a second axial end, where the first axial ends are coupled to the shoe and the second axial ends are coupled to a ring structure. Aspects of the disclosure are directed to an engine comprising: a first structure, a second structure configured to rotate relative to the first structure, and a floating, non-contact seal that interfaces the first structure and the second structure, where the seal includes: a shoe, and at least three beams, where each beam has a first axial end and a second axial end, where the first axial ends are coupled to the shoe and the second axial ends are coupled to a ring structure.

Abstract: A multi-ply seal ring includes an inner seal ring and an outer seal ring. The inner seal ring includes a plurality of discrete inner ring segments disposed circumferentially about an axis in an annular array. The inner ring segments are arranged end-to-end within the annular array. The inner seal ring is engaged with and nested within the outer seal ring. In addition or alternatively, the outer seal ring may include a plurality of discrete outer ring segments disposed circumferentially about the axis in an annular array.

Abstract: A gas turbine engine includes a stator vane. A blade outer air seal may be adjacent to the stator vane. A seal assembly may be disposed between the blade outer air seal and the stator vane. The seal assembly may include a first seal disposed between a first airflow path and a second airflow path. A second seal may be disposed between the first airflow path and the second airflow path. The second seal may be disposed in a series arrangement with the first seal.

Abstract: A blade outer air seal (BOAS) for a gas turbine engine according to one disclosed non-limiting embodiment of the present disclosure includes a seal 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 multitude of cooling fins disposed around the radially outer face of the seal body, at least one of the multitude of cooling fins having at least one aperture.

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 brush seal is provided that includes a top plate, a back plate, and a bristle pack. The bristle pack is secured at a joint between the top plate and the back plate. The bristle pack includes a first bristle set extending from the joint, a second bristle set extending from the joint, and a sliding backing plate. The sliding backing plate includes a sliding portion and a support portion. The sliding portion is disposed contiguous with the second bristle set. The support portion has an inner face and a support face. The inner face is in contact with a distal end of the second bristle set. The support face is configured to support at least a portion of the first bristle set. The sliding backing plate is configured to slide relative to the second bristle set.

Abstract: An example case for a gas turbine engine includes a bearing compartment, a radially outer ring, and a radially inner ring. The radially outer ring and radially inner ring are connected by a plurality of circumferentially spaced apart struts and define an annular flow path therebetween. The radially inner ring includes a first flange that extends radially outward from the bearing compartment. The first flange includes an undulation that extends axially away from the plurality of struts, and defines a load path between the bearing compartment and the radially outer ring.

Abstract: Aspects of the disclosure are directed to a brush seal assembly, comprising a compressible core, at least one wire wound around the core, and a backing plate that is substantially ‘c’ shaped or ‘u’ shaped that at least partially encases the at least one wire. Aspects of the disclosure are directed to a method comprising separating compressible cores with a spacer, winding wires around the cores and the spacer, and forming substantially ‘c’ shaped or ‘u’ shaped backings about the wires to obtain a package of brush seal assemblies.

Abstract: Aspects of the disclosure are directed to a system for maintaining a seal about an engine centerline, comprising: a first component, a second component, and a dog-bone seal coupled to the first component and the second component, wherein the first component is configured with at least one spline that is configured to engage a slot in the dog-bone seal to radially center the dog-bone seal relative to the first component and the second component.

Abstract: A brush seal includes a bristle pack disposed between a top plate and a back plate. A sliding backing plate is disposed between the bristle pack and the back plate. The bristle pack includes a first bristle set that extends across a seal gap between a first engine component and a second engine component to prevent leakage across the seal gap and a second bristle set that exerts a force on the sliding backing plate. The sliding backing plate supports at least a portion of the first bristles and extends across the seal cavity to contact the second component. The force exerted on the sliding backing plate drives the sliding backing plate into contact with the second component.

Abstract: A seal assembly includes a blade outer air seal, a downstream vane, and a pressure wall, according to various embodiments. The blade outer air seal may include a radially outer surface and the downstream vane may be coupled to the blade outer air seal via a fluid sealing engagement. The pressure wall may be coupled to the blade outer air seal and may define a metering orifice. In various embodiments, the metering orifice of the pressure wall is configured to meter air flow from a first plenum upstream of the pressure wall to a second plenum downstream of the pressure wall. In various embodiments, at least a preponderance of the radially outer surface of the blade outer air seal at least partially defines the first plenum and the fluid sealing engagement at least partially defines the second plenum.

Abstract: A gas turbine engine article includes a substrate that has at least one step, and the step includes an undercut. A thermally insulating topcoat is disposed on the substrate. The thermally insulating topcoat includes at least one fault that extends from the step.

Abstract: This disclosure relates to a gas turbine engine including a blade outer air seal (BOAS) mounted radially outwardly of a blade. The engine further includes a BOAS support. A radial dimension of the BOAS support is selectively changeable in response to a flow of fluid through the BOAS support to adjust a radial position of the BOAS relative to the blade.

Abstract: A blade tip clearance system includes first and second control rings. The first control ring has a first coefficient of thermal expansion (CTE) and a first thermal response rate. The second control ring is located radially outward of and operatively connected to the first control ring and has a second CTE that is different from the first CTE and a second thermal response rate that is different from the first thermal response rate. Thermal expansion and contraction of the first and the second control rings controls a radial position of the blade tip clearance system relative to a rotating blade component.