Abstract: A particle separator configured to be disposed within a turbine engine diffuser flow path is provided. The particle separator includes a housing that extends between a forward end and an aft end, first and second swirl vane stages, and first and second air flow exits. The first swirl vane stage is configured to cause an air flow passing through the first swirl vane stage to swirl within the housing as the air flow travels axially in a direction from the forward end toward the aft end of the housing. The second swirl vane stage is configured to cause a first portion of the air flow passing through the second swirl vane stage to redirect into a substantially axially directed flow that enters the second air flow exit.

Abstract: An assembly for a gas turbine engine includes at least one rotational assembly, an engine static structure, a compressor, one or more compressed air loads, and a particulate separator assembly. The at least one rotational assembly includes a shaft, a bladed compressor rotor, and a bladed turbine rotor. The engine static structure includes an engine case assembly. The engine case assembly surrounds the at least one rotational assembly. The compressor includes the bladed compressor rotor. The compressor is configured to form a compressed air flow. The one or more compressed air loads are disposed within the engine case assembly. The particulate separator assembly includes a plurality of particulate separators. The plurality of particulate separators are disposed outside of the engine case assembly. The plurality of particulate separators are configured to separate particulate from the compressed air flow and direct the compressed air flow to the one or more compressed air loads.

Abstract: A particle separator configured to be disposed within a turbine engine diffuser flow path is provided. The particle separator includes a housing, a hollow cavity member, and a helical member. The housing has a forward end inlet, an exterior wall, an aft wall, an interior cavity, and an aft exhaust passage extending outwardly from the exterior wall adjacent the aft wall. The hollow cavity member has a first portion, an interior region, and an open aft end. The first portion is disposed within the housing interior cavity and includes perforations. The helical member is disposed within the housing interior cavity. The helical member forms a helical passage between the housing exterior wall and the first portion of the hollow cavity member. The helical passage has a passage inlet disposed adjacent to the housing forward end inlet and a terminal end in communication with the aft exhaust passage.

Abstract: A turbine for a gas turbine engine includes, among other things, a shaft rotatable about a longitudinal axis, a turbine rotor including one or more rows of turbine blades and a disk assembly coupled to the shaft. The disk assembly includes one or more disks each having an attachment region extending radially between an inner boundary and an outer boundary, the outer boundary is established by an outer periphery of the respective disk, the attachment region defines an array of slots distributed about the outer periphery, each of the slots extends radially inwardly from the outer boundary to the inner boundary, and each of the slots is dimensioned to receive a root section of a respective one of the turbine blades to mount the turbine blades to the disk assembly.

Abstract: A fan has a rotating bulge defining a radially outermost point of the rotating cone relative to a rotational axis. The rotating bulge is defined by a radially outwardly extending forward portion leading to the radially outermost point on the bulge and then a radially inwardly extending portion aft of the radially outermost point of the bulge leading into the core engine path within the splitter housing. The core splitter housing has a leading edge, and the bulge radially outermost point being upstream of the leading edge of the splitter housing. The bulge radially outermost point is a first radial distance from a rotational axis, and the leading edge of the splitter housing being at a second radial distance of the rotational axis, and a ratio of the first radial distance to the second radial distance being at least 0.95.

Abstract: An airfoil includes a CMC airfoil wall that defines a platform, a root that extends radially from the platform, and an airfoil section that extends radially from the platform opposite the root. The CMC airfoil wall includes a wishbone-shaped fiber layup structure. The wishbone-shaped fiber layup structure includes first and second arms that merge into a single leg. The first and second arms are comprised of a network of fiber tows, and the single leg includes fiber tows from each of the first and second arms that are interwoven in the single leg. The single leg forms a portion of the root, and the first and second arms extend in opposite directions from the single leg and form a portion of the platform.

Abstract: An assembly is provided for a turbine engine. This assembly includes an engine core extending along an axis. The engine core includes a first compressor section, a second compressor section, a flowpath, a bleed port, an inner passage, an outer passage and a flow diverter. The flowpath extends longitudinally through the first compressor section and the second compressor section. The bleed port fluidly couples the flowpath to the inner passage and the outer passage in parallel. The bleed port is located longitudinally along the flowpath between the first compressor section and the second compressor section. The flow diverter is located at an inlet into the bleed port from the flowpath. The flow diverter is configured to move between a first position and a second position.

Abstract: An airfoil includes a CMC airfoil wall that defines a platform, a root that extends radially from the platform, and an airfoil section that extends radially from the platform opposite the root. The CMC airfoil wall includes a wishbone-shaped fiber layup structure. The wishbone-shaped fiber layup structure includes first and second arms that merge into a single leg. The first and second arms are comprised of a network of fiber tows, and the single leg includes fiber tows from each of the first and second arms that are interwoven in the single leg. The single leg forms a portion of the root, and the first and second arms extend in opposite directions from the single leg and form a portion of the platform.

Abstract: An assembly is provided for an engine. This engine assembly includes a heat exchange apparatus, and the heat exchange apparatus includes a core, a shell and an inner flowpath. The core includes a core sidewall and a plurality of internal passages. The core sidewall extends axially along and circumferentially around an axis. The core sidewall extends radially from a core inner side to a core outer side. The core inner side forms an outer peripheral boundary of the inner flowpath. The internal passages are arranged circumferentially about the axis. Each of the internal passages extends axially in the core sidewall. The shell extends axially along and circumscribes the core. The shell is abutted radially against the core outer side.

Abstract: An assembly for an aircraft engine includes a variable area nozzle. The variable area nozzle includes a nozzle wall, a nozzle sleeve, an actuation system and a flowpath extending axially along an axis through the variable area nozzle and radially between the nozzle wall and the nozzle sleeve. The nozzle sleeve includes a support and a liner. The support includes a first mount and a second mount axially spaced from the first mount. The liner extends axially along and circumferentially about the support. The liner is axially and circumferentially anchored to the support through the first mount. The liner is circumferentially anchored to the support through the second mount. The actuation system is coupled to the support. The actuation system is configured to move the nozzle sleeve axially along the axis relative to the nozzle wall.

Abstract: A blade outer airseal (BOAS) assembly of a gas turbine engine includes a plurality of BOAS segments arrayed circumferentially about the engine central longitudinal axis, and a plurality of BOAS carriers located radially outboard of the plurality of BOAS segments. Each BOAS carrier is supportive of at least one BOAS segment. The BOAS assembly includes plurality of adjustment levers. Each adjustment lever is operably connected to at least two BOAS carriers of the plurality of BOAS carriers. Rotation of each adjustment lever about a respective pivot axis urges movement of the plurality of BOAS segments in a radial direction thereby adjusting a radial gap between the turbine rotor and the plurality of BOAS segments.

Abstract: A blade outer airseal assembly of a gas turbine engine includes a blade outer airseal (BOAS) segment, and a BOAS carrier positioned radially outboard of the BOAS segment relative to an engine central longitudinal axis. The BOAS segment is secured to the BOAS carrier, and an adjustment lever is operably connected to the BOAS carrier. Rotation of the adjustment lever about a pivot axis at a pivot urges movement of the BOAS segment in a radial direction.

Abstract: A variable area nozzle assembly includes a fixed center plug, a fixed outer nozzle, and an inner nozzle. The fixed center plug includes a seal. The inner nozzle is disposed between the fixed center plug and the fixed outer nozzle. The inner nozzle includes an inner nozzle body. The inner nozzle body forms a primary duct between the inner nozzle and the fixed outer nozzle. The inner nozzle body forms a secondary duct and a gap. The gap is a nozzle outlet of the secondary duct. The secondary duct is disposed between the inner nozzle body and the fixed center plug upstream of the seal. The inner nozzle body forms a plurality of apertures. The inner nozzle body is translatable between a first position and a second position. In the first position, the primary duct has a first cross-sectional area and the plurality of apertures are isolated from the gap by the seal.

Abstract: A variable area nozzle assembly includes a fixed center plug, a fixed outer nozzle, and an inner nozzle. The fixed center plug includes a plug body and a protrusion. The fixed outer nozzle extends about the fixed center plug. The inner nozzle extends about the fixed center plug. The inner nozzle is disposed between the fixed center plug and the fixed outer nozzle. The inner nozzle forms a primary duct between the inner nozzle and the fixed outer nozzle. The primary duct includes an exit plane. The inner nozzle body forms a secondary duct and a gap. The gap is a nozzle outlet of the secondary duct. The gap includes a gap inlet and a gap outlet. The gap inlet is disposed upstream of the protrusion. The gap outlet is disposed downstream of the protrusion. The inner nozzle body is translatable between a first position and a second position. In the first position, the primary duct has a first area at the exit plane and the inner nozzle body is disposed at the protrusion.

Abstract: Tangential onboard injector assemblies include a TOBI and a TOBI blocker assembly arranged upstream of an inlet of the TOBI. The TOBI blocker assembly includes a first blocker plate upstream from the inlet and defining a blocker cavity between the first blocker plate and the TOBI. The first blocker plate includes a blocking portion and a filter portion to permit airflow therethrough. A second blocker plate is arranged within the blocker cavity to obstruct a flow of air containing particulate matter and a gap is present between the second blocker plate and the first blocker plate to permit flow of air around the second blocker plate. A third blocker plate is arranged downstream from the second blocker plate within the blocker cavity and includes a feed hole. Air within a region between the second blocker plate and the third blocker plate flows through the feed hole into the TOBI.

Abstract: A method includes providing a ceramic insert on a mandrel, the mandrel and the ceramic insert together define a peripheral working surface, forming a fiber preform by wrapping a fiber layer around the mandrel and the ceramic insert so as to conform to the peripheral working surface, removing the mandrel from the fiber preform to leave a cavity in the fiber preform, the ceramic insert remaining in the fiber preform and bordering the cavity, and densifying the fiber preform with a ceramic matrix to form a gas turbine engine component.

Abstract: A blade outer airseal (BOAS) assembly of a gas turbine engine includes a BOAS segment, and a BOAS carrier located radially outboard of the BOAS segment relative to an engine central longitudinal axis. The BOAS segment secured to the BOAS carrier. The BOAS assembly is selectably radially movable during operation of the gas turbine engine. One or more secondary retention features are configured to limit the radial movement of the BOAS assembly.

Abstract: A vane arc segment includes ceramic matrix composite (CMC) fairing that has an airfoil section, a first single-sided platform at the outer radial end projecting from the suction side wall, and a second single-sided platform at the inner radial end projecting from the pressure side wall. The first single-sided platform is comprised of a fiber layer that extends from the airfoil section and turns into the first single-sided platform, and the second single-sided platform is comprised of the fiber layer that from the airfoil section and turns into the second single-sided platform.

Abstract: Disclosed is a ceramic matrix component having a fibrous core and a ceramic matrix composite shell surrounding at least a portion of the fibrous core. The ceramic matrix composite shell comprises a fibrous preform. The fibrous core has a greater porosity than the fibrous preform. A method of making the ceramic matrix component is also disclosed.

Abstract: A method includes providing a ceramic insert on a mandrel, the mandrel and the ceramic insert together define a peripheral working surface, forming a fiber preform by wrapping a fiber layer around the mandrel and the ceramic insert so as to conform to the peripheral working surface, removing the mandrel from the fiber preform to leave a cavity in the fiber preform, the ceramic insert remaining in the fiber preform and bordering the cavity, and densifying the fiber preform with a ceramic matrix to form a gas turbine engine component.