RTX Corporation Patente

An assembly includes a first powerplant component (142A), a second powerplant component (142B), a first inter-circuit heat exchanger (152), a first fluid circuit (144A) and a second fluid circuit (144B). The first fluid circuit (144A) is configured to service the first powerplant component (142A) using a first working fluid in the first fluid circuit (144A). The first fluid circuit (144A) includes a first circuit reservoir (150A), a first circuit pump (158A) and a first circuit path (148A) extending through the first circuit reservoir (150A), the first circuit pump (158A) and the first inter-circuit heat exchanger (152). The second fluid circuit (144B) is fluidly independent of the first fluid circuit (144A). The second fluid circuit (144B) is configured to service the second powerplant component (142B) using a second working fluid in the second fluid circuit (144B). The second fluid circuit (144B) includes a second circuit reservoir (150B), a second circuit pump (158B) and a second circuit path (148B) extending through the second circuit reservoir (150B), the second circuit pump (158B) and the first inter-circuit heat exchanger (152).

An assembly is provided for an aircraft powerplant (20). This assembly includes an electric machine (106A), an electric machine controller (180A), an electric cable (124), a conduit (142) and an air circuit (168). The electric machine controller (108A) is configured to control operation of the electric machine (106A). The electric cable (124) electrically couples the electric machine controller (108A) to the electric machine (106A). The conduit (142) extends longitudinally from a first end (144) of the conduit (142) to a second end (146) of the conduit (142). The electric cable (124) extends longitudinally through a bore of the conduit (142). The air circuit (168) is configured to cool the electric cable (124) within the conduit (142) using a flow of air. The air circuit (168) includes the bore of the conduit (142).

An assembly for an aircraft powerplant (20) includes a transmission (204), a first powerplant component (142C), a second powerplant component (142D), a first fluid circuit (146A) and a second fluid circuit (146B). The first fluid circuit (146A) is configured to: cool and/or lubricate the first powerplant component (142C); and provide a first working fluid to the transmission (204) to hydraulically actuate the transmission (204). The first fluid circuit (146A) includes a first circuit reservoir (186A), a first circuit pump (194A) and a first circuit path (184A) extending through the first circuit reservoir (186A), the first circuit pump (194A) and the transmission (204). The second fluid circuit (146B) is configured to: cool and/or lubricate the second powerplant component (142D); and provide a second working fluid to the transmission (204) to hydraulically actuate the transmission (204). The second fluid circuit (146B) includes a second circuit reservoir (186B), a second circuit pump (194B) and a second circuit path (184B) extending through the second circuit reservoir (186B), the second circuit pump (194A) and the transmission (204).

An assembly is provided for an aircraft powerplant (20). This assembly includes a first powerplant component (106A; 108A), a second powerplant component (106B; 108B), a third powerplant component (144) and a fluid system (36). The fluid system (36) includes a first fluid circuit (142A), a second fluid circuit (142B), a third fluid circuit (142C) and a heat exchanger (154). The first fluid circuit (142A) extends through the heat exchanger (154) and is configured to service the first powerplant component (106A; 108A). The second fluid circuit (142B) extends through the heat exchanger (154) and is configured to service the second powerplant component (106b; 108B). The third fluid circuit (142C) extends through the heat exchanger (154) and is configured to service the third powerplant component (144). The heat exchanger (154) is configured as or otherwise includes a radiator (162).

A fabric reinforced seal assembly, comprising a substrate; a seal (100) comprising a fabric (108) having a first side (102) facing the substrate, a second side (104) facing away from the substrate and a thickness defined between the first side (102) and the second side (104), and an elastomer (106) infiltrated into the fabric (108) from the second side (102) toward the first side (102), wherein the elastomer (106) extends into the fabric (108) from the second side (104) for a distance of less than or equal to 50% of the thickness of the fabric (108).

An assembly is provided for an aircraft powerplant (20). This assembly includes an engine case (78), an engine core (48), an electric machine (100) and a shield (140). The engine core (48) is housed within the engine case (78). The engine core (48) includes a compressor section (43), a combustor section (44) and a turbine section (45). The electric machine (100) is disposed outboard of and mounted with the engine case (78). The electric machine (100) includes an electric machine rotor (104), an electric machine stator (106) and an electric machine case (108). The electric machine rotor (104) is housed within the electric machine case (108) and is rotatable about an axis (110). The electric machine stator (106) is housed within the electric machine case (108) and is next to the electric machine rotor (104). The shield (140) is disposed between the electric machine case (108) and the engine case (78). The shield (140) extends axially along and partially circumferentially about the electric machine (100).

A method of inspecting a thin walled structure (100) includes the steps of manufacturing, using PBF-L or PBF-EB techniques, a primary structure wall (104) connected to a shielding wall (102) with a plurality of support pins (106) that extend from the shielding wall (102) to the primary structure wall (104), wherein the primary structure wall (104) includes primary structure wall outer and inner surfaces (104a, 104b) and the shielding wall (102) includes shielding wall outer and inner surfaces (102a, 102b); creating, using a vision inspection system, one or more inspection images of the thin walled structure (100), and analyzing the one or more inspection images to determine if any of the shielding wall outer surface (102a) and shielding wall inner surface (102b) and primary structure wall outer surface (104a) and primary structure wall inner surface (104b) exhibit planar surface distortions that exceed tolerances.

A vane arm (22) includes a first end (70) and a second end (72); a shank (74) extending between the first end and the second end and having a first face (76) and a second face (78); a first hole (110) in the shank proximate the first end; a second hole (114) in the shank proximate the second end; a tab (80) projecting distally from the shank proximate the second end and including a bend (84; 86); and a distal portion (90) distally of the bend projecting back toward the first end and less than fully overlapping the second hole, if at all. Another vane arm (22) includes a first end (70) and a second end (72); a shank (74) extending between the first end and the second end and having a first face (76) and a second face (78); a first hole (110) in the shank proximate the first end; a second hole (114) in the shank proximate the second end (72); and an elongate third hole (150) in the shank between the first hole and the second hole and spanning at least 40 percent of a distance between the first hole and the second hole or at least 25 percent of an overall length (LA) of the vane arm. A vane assembly (20), a vane stage, as well as a method for using the vane arm are also provided.

A process for fabricating a melt-infiltrated ceramic matrix composite comprising fabricating a ceramic fiber preform comprising a fiber, a fiber tow, or both; depositing a structural support material coating on the fiber, the fiber tow, or both to form a partially densified ceramic matrix composite comprising at least one pore, at least one porous network or both; depositing a fiber protection material coating on the fiber, the fiber tow, or both; infiltrating a slurry composition comprising a solvent into the pore, the porous network or both; drying the partially densified ceramic matrix composite to remove the solvent; depositing a sacrificial material coating on the fiber protection material coating; and, melt-infiltrating a molten infiltrant into the partially densified ceramic matrix composite to form a melt-infiltrated ceramic matrix composite.

An air-cooled turbine component includes a body (90) extending between and to a first radial surface (112) and a second radial surface (114) relative to a centerline axis. The body (90) includes an outer passage wall (126) and an inner passage wall (128) forming an internal cooling passage (116) between the first radial surface (112) and the second radial surface (114). The internal cooling passage (116) extends lengthwise through the body (90) between and to an upstream end (120) of the internal cooling passage (116) and a downstream end (122) of the internal cooling passage (116) along a lengthwise axis of the internal cooling passage (116). The body (90) forms a flow augmentation feature assembly (134) at the inner passage wall (128) within the internal cooling passage (116). The flow augmentation feature assembly (134) includes a plurality of trip strips (144) arranged in a ladder pattern (146) along the inner passage wall (128). Blade outer air seal (BOAS) segments are provided as the air-cooled turbine component.

A turbine (62) component includes a body including an outer passage wall (126) and an inner passage wall (128) forming an internal cooling passage (116). The internal cooling passage (116) extends lengthwise through the body between and to an upstream end (120) of the internal cooling passage (116) and a downstream end (122) of the internal cooling passage (116) along a lengthwise axis (118). The body forms a flow augmentation feature assembly (134) at the inner passage wall (128) within the internal cooling passage (116). The flow augmentation feature assembly (134) consecutively includes a first flow augmentation feature pattern (138A) extending lengthwise along a first axial portion (144A) of the internal cooling passage (116) at the upstream end (120), a second flow augmentation feature pattern (138B) extending lengthwise along a second axial portion (144B) of the internal cooling passage (116), and a third flow augmentation feature pattern (138C) extending lengthwise along a third axial portion (144A) of the internal cooling passage (116) disposed at the upstream end (120). The second flow augmentation feature pattern (138B) is different than the first flow augmentation feature pattern (138A) and the third flow augmentation feature pattern (138C).Blade outer air seal (BOAS) segments are provided as the air-cooled turbine component.

An air-cooled turbine component (88) includes a body (90) extending between and to a first radial surface (112) and a second radial surface (114) relative to a centerline axis (74). The body includes an outer passage wall (126) and an inner passage wall (128) forming an internal cooling passage (116) between the first radial surface and the second radial surface. The internal cooling passage extends lengthwise through the body between and to an upstream end (120) of the internal cooling passage and a downstream end (122) of the internal cooling passage along a lengthwise axis (118) of the internal cooling passage. The internal cooling passage includes a first passage segment (132) and at least one second passage segment (130). The first passage segment has a first height (H2) between the outer passage wall and the inner passage wall. The at least one second passage segment has a second height (H1) between the outer passage wall and the inner passage wall. The second height is greater than the first height. The at least one second passage segment is disposed on one or both of the upstream end or the downstream end.A blade outer air seal (BOAS) segment (88) is the air-cooled turbine component.A method is provided for forming an outlet aperture (140) of the BOAS segment (88).

An air-cooled turbine component includes a body (90) extending between and to a first radial surface (114) and a second radial surface (112) relative to a centerline axis (74). The body forms an internal cooling passage (116), a supply orifice (130), and a supply plenum (132) interconnected in fluid communication. The internal cooling passage is disposed between the first radial surface and the second radial surface. The internal cooling passage is formed by and radially between an inner passage wall (120) and an outer passage wall (118) of the body. The internal cooling passage includes an enclosed passage segment (122) and a downstream passage segment (126), the enclosed passage segment extending to a terminal end (124) formed by the body. The supply orifice extends between and to an inlet (136) of the supply orifice and an outlet (138) of the supply orifice along an orifice axis (134). The inlet is disposed at the second radial surface. The outlet is disposed at the supply plenum. The orifice axis extends obliquely relative to the inner passage wall and the outer passage wall, the orifice axis extending toward the enclosed passage segment in a direction from the inlet to the outlet. The supply plenum is disposed between and connects the internal cooling passage and the supply orifice together in fluid communication. The supply plenum is disposed between the enclosed passage segment and the downstream passage segment. A blade outer air seal (BOAS) (86) include a plurality of BOAS segments (88) assembled together in circumferential series about a BOAS axis (74), each of the BOAS segments being the air-cooled turbine component. A turbine section (60) of a gas turbine engine (22) includes a bladed turbine rotor (66) configured for rotation about a rotational axis (74), the bladed turbine rotor including a rotor stage (76), the rotor stage including a plurality of rotor blades (80); and the BOAS.

An air-cooled turbine component (80) includes a body (90) extending between and to a first radial surface (112) and a second radial surface (114) relative to a centerline axis (74). The body (90) includes an outer passage wall (126) and an inner passage wall (128) forming an internal cooling passage (116) between the first radial surface (112) and the second radial surface (114). The internal cooling passage (116) extends lengthwise through the body between and to an upstream end (120) of the internal cooling passage (116) and a downstream end (122) of the internal cooling passage (116) along a lengthwise axis (118) of the internal cooling passage (116). The body forms a flow augmentation feature assembly (134) at the inner passage wall (128) within the internal cooling passage (116). The flow augmentation feature assembly (134) includes a plurality of trip strips (144) arranged in a lattice pattern (146) along the inner passage wall (128).Blade outer air seal (BOAS) segments are provided as the air-cooled turbine component.

An air-cooled turbine component includes a body (104) extending between and to a first radial surface and a second radial surface relative to a centerline axis. The body forms an internal cooling passage (130), a supply orifice (136), and a supply plenum (138) interconnected in fluid communication. The internal cooling passage is disposed between the first radial surface and the second radial surface. The internal cooling passage is formed by and radially between an inner passage wall (132) and an outer passage wall (134) of the body. The supply orifice extends between and to an inlet (142) of the supply orifice and an outlet (154) of the supply orifice along an orifice axis (140). The inlet is disposed at the second radial surface. The outlet is disposed at the supply plenum. The orifice axis extends obliquely relative to the inner passage wall and the outer passage wall. The supply plenum is disposed between and connecting the internal cooling passage and the supply orifice together in fluid communication. The supply plenum is formed by a first plenum wall (150) and a second plenum wall (152) of the body. The first plenum wall forms the outlet. The first plenum wall is oriented substantially orthogonal to the orifice axis. The second plenum wall extends between and to the first plenum wall and the outer passage wall. A blade outer air seal (BOAS) (86) includes a plurality of BOAS segments (86) assembled together in circumferential series about a BOAS axis (74), each of the BOAS segments being such a turbine component. A turbine section (60) of a gas turbine engine (22) includes a bladed turbine rotor (66) configured for rotation about the rotational axis (74), the bladed turbine rotor (66) including a rotor stage (76), the rotor stage (76) including a plurality of rotor blades (80); and the BOAS (86).

Apparatus and associated methods relate to geometry of cooling holes (18) in a combustion liner (10) for a gas turbine engine. The combustion liner (10) includes a base-alloy substrate (12) having a bottom surface (22) and a top surface (26), a thermal barrier coat (TBC) metallic layer (14) on the top surface (26) of the base-alloy structure (12), and a TBC ceramic layer (16) on the TBC metallic layer (14). The cooling holes (18) are formed at oblique angles to the top surface (26) of the TBC ceramic layer (16), thereby forming oval-characteristic exit apertures (24) at the top surface (26) of the TBC ceramic layer (16). The oval-characteristic exit apertures (24) define a long axis (L) and a short axis (S). The long axis (L) extends between an upstream side and a downstream side and the short axis (S) extends between opposite lateral sides, thereby defining opposite lateral sidewalls (28), which have stepped sidewall profiles, in the cooling holes (18).

A method for separating contaminants from metal alloy powder, includes subjecting a mixture of the metal alloy powder and the contaminants sequentially to at least two methods selected from the group consisting of mechanical separation; electrostatic separation; washing with solvent, acid or base; subjecting to heat; fluidized bed treatment and aerodynamic separation. A number of specific combinations of steps are disclosed.

A method of assembling a gas turbine engine (20) is provided that includes: mounting a seal (68) to an annular vane support (66), the annular vane support (66) having an annular surface (70) and a plurality of lugs (72) extending out from the annular surface (70), wherein the seal (68) has a first side surface (68B) and a second side surface (68C) opposite the first side surface (68B); and disposing a plurality of retainer clips (74; 674; 774; 874; 974; 1074), each engaged with the first side surface (68B) of the seal (68) and engaged with at least one of the lugs (72), and configured to retain the seal (68) mounted to the annular vane support (66). The retainer clips (74...1074) are configured to remain in the engine (20) after assembling. The retainer clips (74...1074) are configured to disperse when subjected to engine (20) operating temperatures.

Apparatus and associated methods relate to geometry of cooling holes (18) in a combustion liner (10) for a gas turbine engine. The combustion liner (10) includes a base-alloy substrate (12) having a bottom surface (22) and a top surface (26), a thermal barrier coat (TBC) metallic layer (14) on the top surface (26) of the base-alloy structure (12), and a TBC ceramic layer (16) on the TBC metallic layer (14). The cooling holes (18) are formed at oblique angles to the top surface (26) of the TBC ceramic layer (16), thereby forming oval-characteristic exit apertures (24) at the top surface (26) of the TBC ceramic layer (16). The oval-characteristic exit apertures (24) define a long axis (L) and a short axis (S). The long axis (L) extends between an upstream side (30) and a downstream side (32) and the short axis (S) extends between opposite lateral sides, thereby defining an upstream sidewall (32) and a downstream sidewall (34). The upstream sidewalls (32) have stepped sidewall profiles in the cooling holes (18).

Apparatus and associated methods relate to a combustion liner (10) with a plurality of cooling holes (18). The combustion liner (10) includes a base-alloy substrate (12) having a bottom surface (22) and a top surface (26), a thermal barrier coat (TBC) metallic layer (14) on the top surface (26) of the base-alloy structure (12), and a TBC ceramic layer (16) on the TBC metallic layer (14). The TBC ceramic layer (16) has a top surface (26) configured to be exposed to combustion gases during operation within a gas turbine engine. A ratio of a combined hole volume of the plurality of cooling holes (18) to a combustion-liner volume of the region of the combustion liner (10) to which the plurality of cooling holes (18) belong is greater than 20 percent.

A blade outer air seal (BOAS) assembly (56) is provided that includes an annular BOAS support ring (58), a plurality of BOAS segments (60), and a plurality of spring clips (68). The BOAS segments are disposed radially inside of the annular BOAS support ring. The spring clips are engaged with the BOAS segments and the BOAS support ring. Each spring clip includes a forward segment (86), an aft segment (90), and an engagement segment (92). The BOAS assembly is disposable in an engaged configuration wherein the forward segments are engaged with a surface of the BOAS support ring, the aft segments are engaged with an aft external surface (98) of the BOAS support ring, and the engagement segments are engaged with the plurality of BOAS segments. In the engaged configuration, the spring clips are configured to bias the plurality of BOAS segments in an aft axial direction.

A low pressure turbine shaft installation tool (10) including a disk (38) having an inner diameter (40) and an outer diameter (42); the disk has an upper surface (44) with a lower surface (46) opposite the upper surface; a set of disk through holes (47) are arrayed about the disk; and at least one threaded fastener (48) is associated with at least one of the disk through holes and the at least one threaded fastener in operative communication with the upper surface.

A method of controlling access to information includes the steps of storing documents in a vector database in a plurality of distinct indices with each of the distinct indices associated with a distinct access level based upon at least one of a security level and an appropriate geographic location, receiving a query from a user for information, and identifying at least one of a security level and a geographic location of the user, and matching the identified at least one of the security level and the geographic location with at least one of a plurality of aliases, with each of the plurality of aliases providing access to particular ones of the indices in the vector database, accessing the vector database to match the appropriate indices to the identified alias and displaying appropriate information from the appropriate indices. A document management system and access combination is also disclosed.

A system is provided for an aircraft. This aircraft system includes a turbine engine and an air system. The turbine engine includes a flowpath, a compressor section, a combustor section and a turbine section. The flowpath extends through the compressor section, the combustor section and the turbine section from an airflow inlet into the flowpath to a combustion products exhaust from the flowpath. The air system includes an air circuit, a first boost compressor, a second boost compressor and a heat exchanger. A circuit inlet into the air circuit is fluidly coupled to the flowpath upstream of the combustor section. The air circuit projects out from the circuit inlet and extends through the first boost compressor, the second boost compressor and the heat exchanger.

An assembly is provided for an aircraft propulsion system (22). This assembly includes a compressor section (45; 45A; 45B), a combustor section (46), a turbine section (47), a flowpath (52), a tail cone structure (88), an electric machine (108) and an air system. The flowpath (52) includes a flowpath inlet (56) and a flowpath exhaust (58). The flowpath (52) extends through the compressor section (45; 45A; 45B), the combustor section (46) and the turbine section (47) from the flowpath inlet (56) to the flowpath exhaust (58). The tail cone structure (88) forms an inner peripheral boundary of the flowpath (52) at the flowpath exhaust (58). The tail cone structure (88) includes an internal volume (142). The electric machine (108) is disposed within the internal volume (142). The air system is configured to direct air, bled from the flowpath (52) upstream of the compressor section (45; 45A; 45B), into the tail cone structure (88) to cool the electric machine (108).

A seal assembly (100) includes a seal (106) formed of a carbon material, and received within an inner bore in a carrier (108). The seal (106) extends for 360 degrees relative to a central axis of a bore (103) in the seal (106). A groove (116) is formed in an outer periphery (112) of the seal (106) and an inner periphery (110) of the carrier (108). A wire (114) is received within the groove (116) to lock the seal (106) within the carrier (108). A gas turbine engine (20) and a method are also disclosed.

An assembly is provided for an aircraft propulsion system (22). This assembly includes a compressor section, a combustor section, a turbine section, a flowpath (52), a tail cone structure (88), an electric machine (108) and an electric fan (174A; 174B). The flowpath (52) includes a flowpath inlet and a flowpath exhaust (58). The flowpath (52) extends through the compressor section, the combustor section and the turbine section from the flowpath inlet to the flowpath exhaust (58). The tail cone structure (88) forms an inner peripheral boundary of the flowpath (52) at the flowpath exhaust. The tail cone structure (88) includes an internal volume (142) and an outlet (149) fluidly coupled with the internal volume (142). The electric machine (108) is disposed within the internal volume (142). The electric fan (174A; 174B) is configured to flow air within the internal volume (142) such that the air cools the electric machine (108) and is exhausted from the tail cone structure (88) through the outlet (149).

An aircraft propulsion system assembly includes a vane structure (150), a tail cone structure (8), an electric machine (108) and an air system (36). The vane structure (150) includes an inner platform (152), an outer platform (154) and a plurality of vanes (156A-H). Each of the vanes (156A-H) extends radially across the flowpath (52) from the inner platform (152) to the outer platform (154). The vanes (156A-H) include a first vane and a second vane. The tail cone structure (8) forms an inner peripheral boundary of the flowpath (52) downstream of the vane structure (150). The tail cone structure (8) includes an internal volume. The electric machine (108) is disposed within the internal volume. The air system (36) is configured to direct air from an air source radially outboard of the flowpath (52), radially inward across the flowpath (52) through the first vane and the second vane, into the tail cone structure (8) to cool the electric machine (108).

A vane (24) includes an airfoil (26) having a first end (34) and a second end (32); and a first shaft (42) protruding from the first end (34) of the airfoil (26) and having a threaded distal portion (50); and proximally of the threaded distal portion, a first lateral facet (57A) and a second lateral facet (57B) opposite the first lateral facet (57A). An an upper end (61A) of the first lateral facet is axially recessed relative to an upper end (61B) of the second lateral facet. Vane assemblies (20) and a vane stage are also provided.

An aircraft propulsion system (20) includes an open propulsor rotor, a turbine engine (32) and an air system (34). The turbine engine (32) is configured to drive rotation of the open propulsor rotor about an axis (24). The turbine engine (32) includes an engine rotor (82, 82, 84, 85) and a shroud (146). The engine rotor (82, 82, 84, 85) includes a plurality of rotor blades arranged circumferentially around the rotor base in an array. The shroud (146) is next to and circumscribes the array of the rotor blades. The air system (34) includes an electric boost compressor (122), an air circuit (116) and a clearance control device (114). The air circuit (116) extends longitudinally through the electric boost compressor (122) from an air source to the clearance control device (114). The clearance control device (114) is configured to control a clearance between the shroud (146) and the rotor blades using air received from the air source through the air circuit (116).

An attachment for a fan inlet case (12) and bearing housing (14) including the fan inlet case (12) connected to the bearing housing (14); an attachment (18) in operative communication with the bearing housing (14), the attachment (18) being threaded onto threads formed into an outer radial surface of a forward end of the bearing housing (14); and a locking ring (28) in operative communication with the attachment between the attachment (18) and the fan inlet case (12), wherein the attachment (18) secures the locking ring (28) and fan inlet case (12) to the bearing housing (14).

A system for feature recognition of design geometry may include one or more processors coupled to memory, which may be collectively operable to execute a mapping environment. The mapping environment may be operable to access one or more input images of geometric features of at least one component design. The mapping environment may be operable to determine, using a machine learning classifier model, a geometry of the respective geometric features of the at least one component design from the one or more input images. The mapping environment may be operable to assign, using the classifier model, predefined names to the respective geometric features based on the determined geometry. The classifier model may be trained with training data, which may include training images of geometric features and the predefined names assigned to the respective geometric features in the training images. A method for feature recognition of design geometry is also disclosed.

A method is provided for preparing a ceramic matrix composite (CMC) having a matrix with a SiC matrix phase and an oxide matrix phase. In one process embodiment, a CMC preform containing fiber tows is subjected to chemical vapor infiltration of SiC to form a SiC matrix phase. The resultant preform is then subjected to slurry infiltration to introduce ceramic particles, e.g., rare earth oxides [e.g., RE<2>O<3>, REO, REO<2>], rare earth silicates [RE silicates], Ca silicate, and/or silicon carbide (SiC) particles, to fill void spaces in the preform. The CMC preform is then subjected to glass melt infiltration fore forming an oxide phase of the matrix. The melt infiltrated CMC preform is subjected to one or more heat treatments to form a crystalline oxide matrix phase.

An aircraft system includes a turbine engine (24) and an air system (36). The turbine engine (24) includes a compressor section (45), a combustor section (46) and a turbine section (47). A flowpath extends through the compressor section (45), the combustor section (46) and the turbine section (47). A first rotating structure includes a first compressor rotor disposed in the compressor section (45). A second rotating structure includes a second compressor rotor disposed in the compressor section (45) along the flowpath between the first compressor rotor and the combustor section (46). The air system (36) includes an air circuit (136) and an electric boost compressor (138). A circuit inlet (146) into the air circuit (136) is fluidly coupled to the flowpath and disposed along the second compressor rotor. The air circuit (136) projects out from the circuit inlet (146) and extends through the electric boost compressor (138).

A system for feature recognition of design geometry according to an implementation of the present disclosure may include one or more processors coupled to memory. The one or more processors may be collectively operable to execute a mapping environment. The mapping environment may be operable to access at least one component design including a plurality of geometric features. The mapping environment may be operable to determine values for geometric attributes associated with the respective geometric features of the component design. The mapping environment may be operable to assign, using a machine learning classifier model, predefined names to the respective geometric features based on the values of the geometric attributes. The mapping environment may be operable to assign values for one or more parameters to the respective geometric features according to the assigned predefined names. A method for feature recognition of design geometry is also disclosed.

A ceramic matrix composite (CMC) component (400) includes a CMC substrate comprising ceramic fiber tows within a ceramic matrix. The substrate has a bottom surface and an upper surface; and a cooling cavity (450) within an interior region of the substrate, the cooling cavity being defined by a cavity bottom wall, cavity side walls and a cavity top wall (457) provided with curvatures to allow for its expansion. A blade outer air seal (BOAS) assembly includes a plurality of BOAS segments arranged to form an annular shaped structure and each including the CMC component. A method of preparing a CMC includes providing a CMC preform with a preform substrate having ceramic fiber tows, the preform substrate having the bottom surface and the upper surface; forming the cooling cavity within the preform, the cooling cavity being defined by the cavity bottom wall, the cavity side walls, and the cavity top wall; and subjecting the preform to densification to form the ceramic matrix composite. A turbine engine (20) includes the BOAS assembly.

An aircraft powerplant assembly includes a cooling plate (128), a first electric machine controller and a second electric machine controller. The cooling plate (128) includes a fluid cooling circuit (164), a cooling circuit inlet (166) and a cooling circuit outlet (168). The fluid cooling circuit (164) extends within the cooling plate (128) from the cooling circuit inlet (166) to the cooling circuit outlet (168). A first controller housing is removably attached to the cooling plate (128) and overlaps the fluid cooling circuit (164). First controller circuitry is in thermal communication with the cooling plate (128) through a wall of the first controller housing. A second controller housing is removably attached to the cooling plate (128) and overlaps the fluid cooling circuit (164). Second controller circuitry is in thermal communication with the cooling plate (128) through a wall of the second controller housing.

A multi-layer component made with a multi-laser PBF-L AM technique includes a defect region that envelopes overlapping points (18'-1, 18'-2, 18'-3) for sequential layers (16'-1, 16'-2, 16'-3) in different locations in sequential overlap regions(16'c), wherein the defect region does not create a structural debit in the component when the component is operated in its intended service.

An aircraft propulsion system (20) includes a core engine (25) and a heat exchanger assembly (62) that is operable in a heat exchanger configuration (124) where the exhaust gas flow (80) is in thermal communication with a working fluid flow (90) and a bypass configuration (122) where the exhaust gas flow passes through the heat exchanger assembly without thermal communication with the working fluid flow.

A gas turbine engine stator stage has: a case (100A, 100B); a plurality of vane clusters (120); and a plurality of vane singlets (150). The plurality of vane clusters (120) each have: an inner diameter first platform; an outer diameter first shroud mounted to the case (100A, 100B); and a plurality of first airfoils extending between the first platform and first shroud. The plurality of vane singlets (150) each have: an inner diameter second platform; an outer diameter second shroud; and a single second airfoil extending between the second platform and second shroud.