Abstract: A seal joint assembly is provided that includes a stationary seal carrier, a stationary seal land and a brush seal. The stationary seal land includes a seal land surface. The brush seal is mounted to the stationary seal carrier. The brush seal includes a first set of bristles and a second set of bristles. The first set of bristles includes a first bristle that contacts the seal land surface. The second set of bristles includes a second bristle that contacts the seal land surface. The second bristle crosses the first bristle.

Abstract: A multi-plane annular brush seal is provided. The brush seal includes a plurality of bristles attached to a front plate and a back plate. The back plate has a BP base end, a first BP inner side surface, a BP transition surface, a second BP inner side surface, and a BP tip end. The first BP inner side surface extends from the BP base end to the BP transition surface, and the second BP inner surface extends from the BP transition surface to the BP tip end. At any given circumferential position of the brush seal, the first BP inner side surface extends substantially parallel to a first plane, and the second BP inner side surface extends substantially parallel to a second plane. The first plane is skewed from the second plane by an obtuse angle.

Abstract: A seal joint assembly is provided that includes a stationary seal carrier, a stationary seal land and a brush seal. The stationary seal land includes a seal land surface. The brush seal is mounted to the stationary seal carrier. The brush seal includes a first set of bristles and a second set of bristles. The first set of bristles includes a first bristle that contacts the seal land surface. The second set of bristles includes a second bristle that contacts the seal land surface. The second bristle crosses the first bristle.

Abstract: A turbulent air reducer assembly for a gas turbine engine includes a windage cover attached to a first standoff and a second standoff, the windage cover tuned to a particular vibration response. A method of managing air within a turbine of a gas turbine engine includes reducing turbulence from a multiple of recesses within a tangential on-board injector with a turbulent air reducer assembly within each of the multiple of recesses, the turbulent air reducer assembly tuned to a particular vibration response from a turbine rotor.

Abstract: A gas turbine engine component has a body extending between two circumferential sides, and between a leading edge and a trailing edge. A refractory metal core within the body forms at least one cooling circuit to utilize fluid to cool the body. When the refractory metal core is removed from the body, the at least one cooling circuit includes an inlet, an outlet, and a passage that varies in cross-sectional area between the inlet and outlet. A method of manufacturing a gas turbine engine, a method of manufacturing a core, and a refractory metal core are also disclosed.

Abstract: A vane cluster for a gas turbine engine includes a first end-of-cluster vane, a second end-of-cluster vane, a neighbor vane adjacent to the second end-of-cluster vane at an interface that includes an angled load interface therebetween; and a multiple of base vanes between the first end-of-cluster vane and the neighbor vane, the angled load interface different than an interface between each of the multiple of base vane.

Abstract: A turbulent air reducer assembly for a gas turbine engine includes a windage cover attached to a first standoff and a second standoff, the windage cover tuned to a particular vibration response. A method of managing air within a turbine of a gas turbine engine includes reducing turbulence from a multiple of recesses within a tangential on-board injector with a turbulent air reducer assembly within each of the multiple of recesses, the turbulent air reducer assembly tuned to a particular vibration response from a turbine rotor.

Abstract: A multi-plane annular brush seal is provided. The brush seal includes a plurality of bristles attached to a front plate and a back plate. The back plate has a BP base end, a first BP inner side surface, a BP transition surface, a second BP inner side surface, and a BP tip end. The first BP inner side surface extends from the BP base end to the BP transition surface, and the second BP inner surface extends from the BP transition surface to the BP tip end. At any given circumferential position of the brush seal, the first BP inner side surface extends substantially parallel to a first plane, and the second BP inner side surface extends substantially parallel to a second plane. The first plane is skewed from the second plane by an obtuse angle.

Abstract: A clearance control ring for a clearance control system of a gas turbine engine includes a contoured radial outer portion that defines a multiple of fins and a multiple of slots. A clearance control system of a gas turbine engine includes a clearance control ring with a radial inner portion from which a contoured radial outer portion extends. The contoured radial outer portion defines a multiple of fins and a multiple of slots. A blade outer air seal assembly with a clearance control ring land which receives the radial inner portion. A method of controlling a radial tip clearance within a gas turbine engine includes tailoring a multiple of fins and a multiple of slots of a clearance control ring for both steady state and transient clearance operations.

Abstract: A casting core for a Blade Outer Air Seal includes a heat exchange cavity core section in communication with a first plenum section and a second plenum section, the first plenum and the second plenum section are of a thickness greater than the heat exchange cavity section.

Abstract: A stator assembly includes a case including an arcuate wall having an aperture with circumferentially spaced first lateral surfaces. A stator vane has an outer platform with a notch. An anti-rotation lug has a base that is received in the notch and a boss extends from the base. The boss is received in the aperture. The boss has second lateral surfaces that engage the first lateral surfaces in an interference fit relationship.

Abstract: A heat shield includes a first end configured to connect to an inner engine case at an inner connecting point, and a second end configured to connect to an outer engine case at an outer connecting point. The heat shield also includes a bulge extending in an aft direction from a radially upstream portion. A compressor section and a gas turbine engine including the heat shield are also disclosed.

Abstract: A gas turbine engine component is provided. The gas turbine engine component comprises a main body having a leading edge and a leading edge wall including an elongated transition portion extending between the leading edge and a proximate flowpath surface of the main body. The elongated transition portion has an axial length that is greater than a radial height. A gas turbine engine is also provided.

Abstract: A clearance control assembly for a gas turbine engine includes a clearance control ring to position a blade outer air seal assembly radially relative to a blade tip. The clearance control ring is compression fit to the blade outer air seal assembly.

Abstract: A gas turbine engine component is provided. The gas turbine engine component comprises a main body and a leading edge cooling passage defined within the main body. The main body has a leading edge and a leading edge wall including an elongated transition portion extending between the leading edge and a proximate flowpath surface of the main body. The leading edge cooling passage comprises an axial flow cooling passage defined within the main body and adjacent to the leading edge wall and has a leading edge periphery that generally conforms to the elongated transition portion of the leading edge wall.

Abstract: A heat shield assembly for a gas turbine engine includes a first heat shield segment defined about an axis and a second heat shield segment defined about the axis. A double circumferential lap joint is defined between the first heat shield segment and the second heat shield segment.

Abstract: A brush seal is provided. The brush seal may comprise an annular backing plate and a first bristle pack coupled to the annular backing plate. The first bristle pack may be oriented in an axial direction. A second bristle pack may be coupled to the first bristle pack and oriented in the axial direction. A bristle of the first bristle pack may have a greater diameter than a bristle of the second bristle pack. A retention structure may be coupled to the second bristle pack.

Abstract: A spring clip for a blade outer air seal (BOAS) includes a clip body defining a first end configured to operatively engage with a BOAS retaining structure on a BOAS and a second end configured to engage with at least one of a mounting hook of a BOAS and a BOAS mount.

Abstract: A gas turbine engine component has a body extending between two circumferential sides, and between a leading edge and a trailing edge. A refractory metal core within the body forms at least one cooling circuit to utilize fluid to cool the body. When the refractory metal core is removed from the body, the at least one cooling circuit includes an inlet, an outlet, and a passage that varies in cross-sectional area between the inlet and outlet. A method of manufacturing a gas turbine engine, a method of manufacturing a core, and a refractory metal core are also disclosed.

Abstract: A gas turbine engine component is provided. The gas turbine engine component comprises a main body having a leading edge and a leading edge wall including an elongated transition portion extending between the leading edge and a proximate flowpath surface of the main body. The elongated transition portion has an axial length that is greater than a radial height. A gas turbine engine is also provided.