RTX Corporation Patente

A method for detecting and removing at least one atmospheric pollutant from a surface of an aircraft engine component (108) includes detecting the at least one atmospheric pollutant on the surface of the aircraft engine component (108) using a test probe (102) and treating the at least one aircraft engine component (108) to generate a layer of at least one of a soot or other carbonaceous substance on the at least one aircraft engine component (108) to at least partially remove the at least one atmospheric pollutant from the surface of the aircraft engine component (108).

An aerospace piston seal ring including a body shaped as an annular ring, the body defines an axis centered within the body; a sealing surface extending radially relative to the axis between an inner diameter and an outer diameter; and a surface feature formed along the sealing surface, the surface feature configured as an interruption in the sealing surface, wherein the surface feature is configured to contain a lubricant.

An attachment pin (240) includes an elongated body with a longitudinal axis extending from a first end (280) to a second end (282), a cooling passageway (290) extending axially from the first end along the longitudinal axis, and one or more cooling air outlets (295) extend radially from the cooling air passageway to provide for discharge of fluid from the cooling passageway. A ceramic matrix composite (CMC) component includes a base (210) having a radial outer surface (212) and a radial inner surface inward surface (215). The base includes a plurality of ceramic fiber plies and a ceramic matrix, a forward flange structure (220) and an aft flange structure (230) each extending from the radial outer surface of the base, the forward flange structure having a first opening (222) and a second opening (225), and the aft flange structure has a first opening (232) and a second opening (235). The CMC component includes a first (240) and a second (250) of the attachment pin for attaching the CMC component to a support structure, the first attachment pin passing through the first opening of the forward flange structure and the first opening of the aft flange structure, and the second attachment pin passing through the second opening of the forward flange structure and the second opening of the second flange structure. A BOAS assembly comprising a plurality of BOAS segments being the CMC component arranged to form an annular shaped structure. A method of cooling a CMC component includes providing the CMC component and the first and the second attachment pins. A turbine engine includes a fan section (22), a compressor section (24), a combustion chamber (56), a turbine section (28), the turbine section including at least one rotor and one or more turbine blade extending radially outwardly from the at least one rotor, and the blade outer air seal assembly being positioned between the one or more turbine blade(s) and an outer casing to the engine.

An apparatus is provided for an aircraft that includes an open rotor propulsion system (20). The open rotor propulsion system (20) includes an open propulsor rotor (34), an open guide vane structure (36) and a turbine engine (32). The open guide vane structure is axially next to the open propulsor rotor. The turbine engine is configured to drive rotation of the open propulsor rotor about an axis (24). An exterior surface (134) of a component (132) of the open rotor propulsion system is exposed to and borders an environment (22) external to the open rotor propulsion system. The component is configured with an acoustic treatment (116) extending axially and circumferentially along the exterior surface (134).

A cooling flow arrangement (200) for a ceramic matrix composite (CMC) component (210) of a turbine engine is provided, wherein the CMC component includes a hot side (230) configured for exposure to a hot gas path (232) of the turbine engine and an opposing cold side (240). At least one cavity (220) is provided in the CMC component, the at least one cavity having an entrance disposed on the cold side that is configured for receiving a cooling flow. A seal (260) is disposed adjacent the at least one cavity and the cold side of the CMC component and is configured to provide a seal between the CMC component and an upstream turbine component. At least one hole (262) extends through the seal and is positioned to provide the cooling flow to the at least one cavity of the CMC component.A method (300) of providing a cooling flow to the CMC component (210) includes: providing (310) the at least one cavity in the CMC component ; disposing (320) the seal adjacent the at least one cavity and the cold side of the CMC component to provide a seal between the CMC component and an upstream turbine component; and (330) passing the cooling flow from a source (250) through the at least one hole.

A ceramic matrix composite (CMC) component (200) for a turbine engine (20) includes a base (210) having a radial outer surface ( 212) and a radial inner surface inward surface (215). The base includes a plurality of ceramic fiber plies and a ceramic matrix. A cooling cavity (260) extends from the outer radial surface of the base into an interior region of the base. The cooling cavity has a cavity opening (262) at the outer radial surface of the base. The cooling cavity is defined by cavity side walls (264), a cavity bottom wall (266), and an impingement plate (270) or a cover plate covering the cavity opening. At least one spring member (280) includes a contact region (310), a first curved region (320a) and a first end section (330a). The first curved region is connected to both the contact region and the first end section. The first end section is in contact with a restraining member (240), and the contact region is in contact with the impingement plate or cover plate and acts to hold the impingement plate or cover plate in position to cover the cavity opening. The CMC component is a blade outer air seal (BOAS) segment. A BOAS assembly includes a plurality of the BOAS segments arranged to form an annular shaped structure. A method of assembling the CMC component includes: providing the CMC component; providing the impingement plate or the cover plate; retaining the impingement plate or cover plate in a position to cover the cavity opening using the at least one spring member. The turbine engine includes a fan section (22), a compressor section (24), a combustion chamber (56), and a turbine section (28). The turbine section includes at least one rotor and one or more turbine blade(s) extending radially outwardly from the at least one rotor; the blade outer air seal assembly positioned between the one or more turbine blade(s); and an outer casing (36).

A method of repairing a ceramic vane includes providing a ceramic vane (49) that includes a platform (64) and an airfoil section (66) that extends from the platform. The platform includes a mounting surface (78) for supporting the ceramic vane and the airfoil section includes a damaged region (70). An over-wrap fiber ply (72) is wrapped around the airfoil section and covers the damaged region. The over-wrap fiber ply is then densified with a ceramic matrix material to form a ceramic matrix composite over-wrap repair layer (74). A build-up layer (76) is applied to the at least one mounting surface. A desired stagger angle is provided to the ceramic vane by adjusting a thickness and contour of the at least one build-up layer.

A ceramic matrix composite (CMC) component (200) includes a base (210) having a radial outer surface (212) and a radial inner surface inward surface (315). The base includes a plurality of ceramic fiber plies and a ceramic matrix. The radial outer surface has a convex shape, at least one cooling cavity (260a, 260b) within the base that extends from the radial outer surface of the base into an interior region of the base. The at least one cooling cavity has a cavity opening at the radial outer surface of the base. The at least one cooling cavity is defined by cavity side walls (264a, 264b), a cavity bottom wall (266a, 266b), and a cover plate (270a; 270b) positioned within the at least one cooling cavity to cover the cavity opening. The cover plate is angled to vary a depth of the cooling cavity between the cover plate and the cavity bottom wall so that the depth varies from a cavity region of greater depth to a cavity region of lesser depth. A method of controlling cooling air flow within a CMC component includes providing the CMC component. A turbine engine (20) includes a fan section (22), a compressor section (24), a combustion chamber (26), and a turbine section (28). The turbine section includes at least one rotor and one or more turbine blade(s) extending radially outwardly from the at least one rotor, and a blade outer air seal assembly positioned between the one or more turbine blade(s) and an outer casing of the engine. The blade outer air seal assembly is formed of a plurality blade outer air seal segments, wherein each blade outer air seal segment is the CMC component.

A brush seal (200) for use with a ceramic matrix composite (CMC) component (210) of a gas turbine engine may feed a cooling flow to the CMC component, which has a hot side (212) configured for exposure to a hot gas path of the gas turbine engine and an opposing cold side (214) having a cavity (216) for feeding a cooling flow to at least one film cooling hole (218) of the CMC component. The brush seal includes: an outer diameter (OD) backing plate (220); an inner diameter (ID) backing plate (230); a plurality of brush seal bristles (240) sandwiched between the OD backing plate and the ID backing plate; at least one hole (250) forming a passage extending through the OD backing plate, the brush seal bristles, and the ID backing plate; and a cover plate (260) attached to an inner surface of the ID backing plate and configured to extend into the cavity of the CMC component.

A controller (140) includes processing circuitry and memory (144) and is operable to control an additive manufacturing process. The processing circuitry receives additive manufacturing parameters and a part design at an analysis module. The processing circuitry is operable to break the part design into a plurality of elements and assign an initial density to each, and determine a likelihood of a defect in each of the plurality of elements utilizing the received additive manufacturing parameters and assigned density, and evaluate the volume of defects across the plurality of elements. The processing circuitry is operable to update at least one of the part design or the additive manufacturing parameters, and then determine the likelihood of updated defects for each of the plurality of elements. The processing circuitry is operable to reach a solution that is at an acceptable volume of defects, while still satisfying an acceptable part design. A method is also disclosed.

A special tool in intermediate form, comprising: at least one aft core assembly (160) having at least one aft alignment aperture (150) and at least one baffle (120); a forward core assembly having at least one forward alignment aperture, the at least one aft core assembly (160) and the forward core assembly disposed adjacent to and aligned with one another; at least one alignment tool disposed within each at least one aft alignment aperture (150), and aligned with and also disposed within each at least one forward alignment aperture; and a preform material (140) disposed about the adjacent, aligned aft core assembly (160) and forward core assembly.

A brush seal (200) of a turbine engine may provide a cooling flow to a ceramic matrix composite (CMC) component (210) that has a hot side (212) exposed to a hot gas path and an opposing cold side (214). The brush seal includes an inner diameter (ID) backing plate (220) providing a radial face seal across a cavity in the CMC component. The brush seal also includes an outer diameter (OD) backing plate (230), a plurality of brush seal bristles (240) disposed between the ID and OD backing plates, and a hole (250) extending through the brush seal to provide the cooling flow to the cavity of the CMC component. A cross-flow channel (224) is disposed on an inner surface of the ID backing plate such that the cross-flow channel extends from an edge of the ID backing plate to the hole to pass leakage air to mix with the cooling flow and reduce an impingement cooling effect thereof.

A method of operating an aircraft (100) includes determining, for an engine (110) of an aircraft (100), a high delta T5 (MDT5) between a first channel and a second channel over a flight of the aircraft (100). The method also includes updating a multilevel perception neural network (200) with the MDT5, and determining, based on the neural network (200), a fuel coking factor for the engine (110).

A ceramic matrix composite (CMC) component segment (210, 212; 310) forming a portion of a substantially ring-shaped gas turbine engine stage includes a mateface (230, 232) extending axially between an upstream end and a downstream end of the CMC component (210) and configured to form an inter-segment gap (250) with an adjacent CMC component (212), and a plurality of film cooling holes (234) having diffused openings at the mateface (230, 232) to provide film cooling of the mateface (230, 232). A method of providing an inter-segment purge flow to such a CMC component (210, 212) includes disposing a first mateface (230) of the CMC component segment (210) next to a second mateface (232) of an adjacent CMC component segment (212) to form an inter-segment gap (250), and providing first film cooling of the first mateface (230) via the plurality of film cooling holes (234) having diffused openings at the first mateface (230).

A brush seal (200; 201; 202; 330) of a turbine engine (20) may provide a cooling flow (255; 355) to a ceramic matrix composite (CMC) component (210; 310; 320) that has a hot side (212; 312) configured for exposure to a hot gas path of the turbine engine (20) and an opposing cold side (214). The brush seal (200...330) includes an inner diameter (ID) backing plate (220) providing a radial face seal (217) across a cavity (216) in the CMC component (210; 310; 320), with the cavity (216) having an entrance disposed on the cold side (214) to receive the cooling flow (255; 355). The brush seal (200...330) also has an outer diameter (OD) backing plate (230), a plurality of brush seal bristles (240) disposed between the ID backing plate (220) and OD backing plate (230) and attached thereto in a welding zone, and a hole (250; 251; 252; 350) extending through the OD backing plate (230), the brush seal bristles (240), and the ID backing plate (220) in the welding zone to provide the cooling flow (255; 355) to the cavity (216) of the CMC component (210; 310; 320).

A gas turbine engine component (100; 200; 600), comprising a first surface having at least one first aperture (130), and exposed to a first plenum (150; 230; 530, 630) comprising a mixed fluid flow and a first pressure (P₁); at least one second surface having at least one second aperture (140a,b,c), and exposed to a gas path fluid flow and at least one second pressure (P₂); at least one channel (120; 210; 605), at least one internal plenum (125) or both at least one channel (120; 210; 605) and at least one internal plenum (125) disposed between and in fluid communication with each the at least one first aperture (130) and the at least one second aperture (140a,b,c).

Environmental barrier coatings (EBCs) (160) are disclosed having a topcoat layer (130) having a primary hafnon phase (150) and a dispersed hafnia phase (140). The hafnia can be homogeneously dispersed or the layer (130) can exhibit a concentration gradient in which the hafnia concentration increases continuously or stepwise. The hafnon primary phase (150) provides a desirable CTE match with the underlying CMC substrate (110), while the dispersed hafnia (140) provides desirable corrosion resistance.

An apparatus is provided for an aircraft that includes a turbine engine (24). The turbine engine (24) includes a compressor section (42), a combustor section (44), a turbine section (46), a support structure (110), a first rotor system (92) and a second rotor system (92). The support structure (110) extends axially along an axis (34). Each rotor system (92) may include an electric machine (96) and a bladed rotor (94). The electric machine (96) includes an electric machine stator (108) and an electric machine rotor (106). The electric machine stator (108) is mounted to the support structure (110). The electric machine rotor (106) circumscribes the electric machine stator (108) and is operatively coupled to the bladed rotor (94). The electric machine (96) is configured to generate an electromagnetic field with the electric machine stator (108) and the electric machine rotor (106). The bladed rotor (94) is configured to rotate about the axis (34).

A gas turbine engine (20) disposed within a nacelle (42) is provided that includes an open rotor propulsion system (28), a compressor section (30), an airflow inlet (24), and an inlet panel structure actuation system (58). The airflow inlet (24) is in fluid communication with the compressor section (30). The airflow inlet (24) is defined by a plurality of inlet panel structures (56). Each inlet panel structure (56) includes an inlet side segment (60) and an outer panel segment (62). The inlet side segments (60) collectively define an outer radial surface (52) of the airflow inlet (24). The inlet side segment (60) and the outer panel segment (62) of each inlet panel structure (56) intersect at an inlet panel leading edge. The inlet panel structure actuation system (58) is configured to selectively actuate the inlet panel leading edge of each inlet panel structure (56) in a radial inward direction or in a radial outward direction.

An aircraft propulsion system (22) includes a gas turbine engine (24) and a power transfer assembly (26; 126). The gas turbine engine (24) includes a high-pressure (HP) spool (40) and a low-pressure (LP) spool (42). The power transfer assembly (26; 126) includes a first motor-generator (80; 128), a second motor-generator (82; 130), a motor control assembly (76; 132), and a common assembly housing (108). The first motor-generator (80; 128) is operably coupled with the HP spool (40). The second motor-generator (82; 130) is operably coupled with the LP spool (42). The first motor-generator (80; 128) is electrically connected to the second motor-generator (82; 130). The motor control assembly (76; 132) is electrically connected to the first motor-generator (80; 128) and the second motor-generator (82; 130). The motor control assembly (76; 132) is configured to operate the first motor-generator (80; 128) to apply a first rotational force to the HP spool (40) or to operate the second motor-generator (82; 130) to apply a second rotational force to the LP spool (42). The common assembly housing (108) houses the first motor-generator (80; 128) and the second motor-generator (82; 130).

A gas turbine engine (20) is provided that includes a fan section (28), a fan bypass air path (38), a compressor section (30), a combustion section (32), a turbine section (34), an annular compressor inlet (42), and an inlet panel structure actuation system (54). The annular compressor inlet (42) is defined by an inner radial surface (46) and a plurality of inlet panel structures (52). Each inlet panel structure (52) includes an inlet side segment (56) and a bypass side segment (58). The inlet side segments (56) collectively define an outer radial surface (48) of the compressor inlet (42). The inlet side segment (56) and the bypass side segment (58) of each inlet panel structure (52) intersect at an inlet panel leading edge. The inlet panel structures (52) are disposable in a radial inner most configuration and in a radial outer most configuration. The inlet panel structure actuation system (54) is configured to selectively actuate the inlet panel structures (52) between the radial inner most configuration and the radial outer most configuration.

A gas turbine engine (20) is provided that includes compressor, combustor, and turbine sections, an engine shaft (30; 32), an accessory gearbox (62), and first and second starters (74; 74A, 74B). The engine shaft (30; 32) is in communication with the compressor section and the turbine section. The accessory gearbox (62) has first and second accessory mounting pads (72; 72A, 72B), and is in communication with the engine shaft (30; 32). The first starter (74; 74A) is mounted on the first accessory mounting pad (72; 72A), and the second starter (74; 74A) is mounted on the second accessory mounting pad (72; 72B). The first starter (74; 74A) is configured to selectively provide rotational drive to the accessory gearbox (62) and the accessory gearbox (62) in turn provides rotational drive to the engine shaft (30; 32). The second starter (74; 74B) is configured to selectively provide rotational drive to the accessory gearbox (62) and the accessory gearbox (62) in turn provides rotational drive to the engine shaft (30; 32).

An aircraft propulsion system (22) includes a gas turbine engine (24), a power transfer assembly (26; 126), and a controller (78). The gas turbine engine (24) includes a first spool (40) and a second spool (42). The power transfer assembly (26; 126) includes a first motor-generator (80; 128), a second motor-generator (82; 130), and a motor control assembly (76; 132). The first motor-generator (80; 128) is operably coupled with the first spool (40). The second motor-generator (82; 130) is operably coupled with the second spool (42). The motor control assembly (76; 132) is electrically connected to the first motor-generator (80; 128) and the second motor-generator (82; 130). The motor control assembly (76; 132) is configured to control an operating voltage of the power transfer assembly (26; 126). The controller (78) is connected in signal communication with the motor control assembly (76; 132). The controller (78) is configured to identify a target operating voltage of the power transfer assembly (26; 126) selected in response to an aircraft altitude input to the controller (78) and control the motor control assembly (76; 132) to control the operating voltage at the target operating voltage.

An aircraft (20) assembly includes first and second propulsion systems (24A, 24B). Each propulsion system includes an open propulsor rotor (60) and a turbine engine (56) configured to drive rotation of the open propulsor rotor (60). The turbine engine (56) includes a compressor section (91), a combustor section (92), a turbine section (93), a first rotating structure, a second rotating structure and a flowpath (96). The flowpath (96) extends through the compressor section (91), the combustor section (92) and the turbine section (93) with the first bladed rotor disposed between the second bladed rotor and the combustor section (92) along the flowpath (96). The open propulsor rotor (60) of the first propulsion system (24A) is configured to rotate a first rotational direction. The open propulsor rotor (60) of the second propulsion system (24B) is configured to rotate a second rotational direction that is opposite the first rotational direction.

An assembly (20) is provided for an aircraft. This assembly (20) includes an engine core (50), an engine case (68), a nacelle wall (166) and an air circuit. The engine core (50) includes a compressor section (45), a combustor section (46) and a turbine section (47). The engine case (68) houses the engine core (50). The nacelle wall (166) is radially outboard of and covers the engine case (68). A compartment (140) is formed by and extends radially between the engine case (68) and the nacelle wall (166). The air circuit includes a manifold disposed within the compartment (140). The air circuit is configured to direct air from an air source into the manifold. The manifold includes a plurality of outlets. The manifold is configured to direct the air into the compartment (140) through the outlets.

An aircraft assembly includes first and second propulsion systems (24A, 24B). Each of the propulsion systems includes an open propulsor rotor (60) and a turbine engine (56) configured to drive rotation of the open propulsor rotor (60). The turbine engine (56) includes a compressor section (91), a combustor section (92), a turbine section (93), a first rotating structure (112), a second rotating structure (116) and a flowpath (96). The first rotating structure (112) includes a first bladed rotor (105, 106). The second rotating structure (116) includes a second bladed rotor (104, 107) and is operable to rotate independent of the first rotating structure (112). The open propulsor rotor (60) of the first propulsion system (24A) is configured to rotate a first rotational direction. The open propulsor rotor (60) of the second propulsion system (24B) is configured to rotate a second rotational direction that is opposite the first rotational direction.

A method for oxide formation on the surface of a ring includes: mounting a ring in a fixture (312); driving rotation of the ring about a first axis (314); contacting an outer diameter surface of the ring with a roller (340); and laser heating the outer diameter surface of the ring to form an oxide layer. The rotation causes compaction of the oxide by the roller (340).

An apparatus is provided for an aircraft that includes an open rotor propulsion system (20). The open rotor propulsion system (20) includes a turbine engine (32), a first electric machine (100A), a second electric machine (100B), a first controller (102A), a second controller (102B) and a first fluid circuit (134). The first controller (102A) is configured to control operation of the first electric machine (100A). The second controller (102B) is configured to control operation of the second electric machine (100B). The first fluid circuit (134) is configured to circulate a first liquid to cool and/or lubricate the first electric machine (100A), the first controller (102A), the second electric machine (100B) and the second controller (102B).

An aircraft apparatus includes a turbine engine (24), a first electric machine (100A), a first controller (102A), a first fluid circuit (134A), a second electric machine (100B), a second controller (102B) and a second fluid circuit (134B). The turbine engine includes a first rotating structure (60A) and a second rotating structure (60B). The first electric machine is operatively coupled to the first rotating structure. The first controller is configured to control operation of the first electric machine. The first fluid circuit is configured to circulate a first liquid and service the first electric machine and the first controller. The second electric machine is operatively coupled to the second rotating structure. The second controller is configured to control operation of the second electric machine. The second fluid circuit is configured to circulate a second liquid and service the second electric machine and the second controller. The second fluid circuit is fluidly discrete from the first fluid circuit.

A system for traceability of manufacturing data 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 a production set of manufacturing instructions associated with a component design. The mapping environment may be operable to access real manufacturing data associated with execution of the production set of manufacturing instructions. The mapping environment may be operable to generate an evaluation set of manufacturing instructions associated with respective unique identifiers. The unique identifiers may be assigned to respective geometric features of the component design. The mapping environment may be operable to generate a mapped set of manufacturing instructions including the unique identifiers assigned to respective portions of the production set of manufacturing instructions based on the evaluation set of manufacturing instructions.

An aircraft assembly (20) is provided that includes a propulsor rotor, a turbine engine core (44), an inner case, a first gearbox (108A), a second gearbox (108B), a first electric machine (96A) and a second electric machine (96B). The turbine engine core (44) includes a first rotating structure (54A) and a second rotating structure (54B). The first gearbox (108A) is mounted with the inner case. The second gearbox (108B) is located remote from the inner case. The first electric machine (96A) is operatively coupled to the first rotating structure (54A) through the first gearbox (108A). The second electric machine (96B) is operatively coupled to the second rotating structure (54B) through the second gearbox (108B). The first electric machine (96A) may be configurable as a first electric motor and/or a first electric generator. In addition or alternatively, the second electric machine (96B) may be configurable as a second electric motor and/or a second electric generator.

An aircraft assembly (20) includes a propulsor rotor, a turbine engine core (44), an inner case, an outer case, a first gearbox (108A), a second gearbox (108B), a first electric machine (96A) and a second electric machine (96B). The turbine engine core (44) is configured to drive rotation of the propulsor rotor about an axis. The turbine engine core (44) includes a first rotating structure (54A) and a second rotating structure (54B). The first rotating structure (54A) includes a first bladed rotor. The second rotating structure (54B) includes a second bladed rotor. The inner case houses the turbine engine core (44). The outer case houses the propulsor rotor. The first gearbox (108A) is mounted with the outer case. The second gearbox (108B) is mounted with the outer case. The first electric machine (96A) is operatively coupled to the first rotating structure (54A) through the first gearbox (108A). The second electric machine (96B) is operatively coupled to the second rotating structure (54B) through the second gearbox (108B).

An apparatus is provided for an aircraft. This apparatus includes a turbine engine (24), a first electric machine (100A), a first controller (102A), a second electric machine (102A), a second controller (102B) and a first fluid circuit (134). The first electric machine is operatively coupled to the turbine engine. The first controller is configured to control operation of the first electric machine. The second electric machine is operatively coupled to the turbine engine. The second controller is configured to control operation of the second electric machine. The first fluid circuit is configured to circulate a first liquid to cool and/or lubricate the first electric machine, the first controller, the second electric machine and the second controller.

An aircraft assembly includes a turbine engine (24), a first electric machine (100A), a first controller (102A), an electric machine fluid circuit (134A) and a controller fluid circuit (134B). The turbine engine includes a compressor section (45), a combustor section (46), a turbine section (47), a flowpath (84) and a first rotating structure (60A). The first rotating structure includes a first bladed rotor disposed in one of the compressor section or the turbine section. The first electric machine includes a first electric machine rotor (104A). The first electric machine rotor is operatively coupled to the first rotating structure. The first controller is configured to electrically couple the first electric machine to an electrical system (98). The electric machine fluid circuit is configured to circulate a first liquid and service the first electric machine. The controller fluid circuit is configured to circulate a second liquid and service the first controller. The controller fluid circuit is fluidly discrete from the electric machine fluid circuit.

An aircraft assembly includes a first electric cable (124A) and an electric device (122). The electric device (122) includes a device housing (126) and a first electric terminal (156A) disposed within an internal volume (130) of the device housing (126). The device housing (126) includes a first metal sidewall (140A), a first wall cable port (150) and an open first wall aperture (152A-C) next to the first wall cable port (150). The first metal sidewall (140A) is between and borders the internal volume (130) and an external environment (148). The first wall cable port (150) projects through the first metal sidewall (140A) from the external environment (148) to the internal volume (130). The open first wall aperture (152A-C) projects through the first metal sidewall (140A) from the external environment (148) to the internal volume (130). The first electric cable (124A) projects longitudinally through the first wall cable port (150) into the internal volume (130). The first electric cable (124A) is received by and is electrically coupled to the first electric terminal (156A) within the internal volume (130).

An apparatus is provided for an aircraft. This apparatus includes an electric machine controller (102A; 102B) and an air cooling circuit (134). The electric machine controller (102A; 102B) includes controller circuitry (114A; 114B) and a controller housing (112A; 112B). The controller circuitry (114A; 114B) is disposed within an interior of the controller housing (112A; 112B). The air cooling circuit (134) includes a cooling boot (142A; 142B). The air cooling circuit (134) is configured to direct air from an air source into the cooling boot (142A; 142B). The cooling boot (142A; 142B) forms an air plenum (168) with the electric machine controller (102A; 102B). The cooling boot (142A; 142B) includes a plurality of air outlets (172). The cooling boot (142A; 142B) is configured to direct the air through the air outlets (172) and into the air plenum (168) to air cool the electric machine controller (102A; 102B).

An apparatus is provided for an aircraft. This apparatus includes an electric machine (100A; 100B) and an air cooling circuit (134). The electric machine (100A; 100B) includes an electric machine rotor and an electric machine stator. The electric machine (100A; 100B) is configured to generate an electromagnetic field with the electric machine rotor and the electric machine stator. The electric machine rotor is configured to rotate about an axis (109A; 109B). The air cooling circuit (134) includes a cooling boot (142A; 142B). The air cooling circuit (134) is configured to direct air from an air source into the cooling boot (142A; 142B). The cooling boot (142A; 142B) forms an air plenum (166) with an exterior surface (174) of the electric machine (100A; 100B). The cooling boot (142A: 142B) includes a plurality of air outlets (178). The cooling boot (142A; 142B) is configured to direct the air through the air outlets (178) and into the air plenum (166) to air cool the electric machine (10A; 100B).

A test rig with a multi-specimen fixture system including an actuator including a central shaft; a moving component in operative communication with the central shaft; a stationary component located proximate the moving component arrayed around the central shaft; a first connector coupled to the stationary component, the first connector configured to secure a test specimen; a second connector coupled to the moving component, the second connector configured to secure the test specimen; a load cell in operative communication with at least one of the first connector or the second connector; and at least one compliant element in operative communication with the moving component, wherein the at least one compliant element is configured to produce an independent load to the test specimen.

A mid-turbine frame (64) of a gas turbine engine (24) includes a fairing (76) and a damping system (102). The fairing (76) includes an outer fairing wall (82), an inner fairing wall (84), and a plurality of struts (86). Each of the outer fairing wall (82) and the inner fairing wall (84) extends between and to an inner surface and an outer surface. The inner surface forms a core flow path (70) through the fairing (76). One of the outer fairing wall (82) or the inner fairing wall (84) forms a damping wall of the fairing (76). The damping system (102) includes at least one damping ring (106) and a plurality of clamp assemblies (104). The at least one damping ring (106) is disposed at the outer surface of the damping wall. The plurality of clamp assemblies (104) is circumferentially arrayed on the damping wall about the axis (44). Each of the plurality of clamp assemblies (104) clamps the at least one damping ring (106) on the damping wall.

An axial compressor rotor assembly (52) includes a plurality of rotor disks (76) extending axially along and circumferentially about a central axis (53) between an upstream end (54) and a downstream end (56). The rotor disks (76) extend radially between an inner radial hub (78) and an outer radial rim (84), and axially between a forward rim end (92A) and a rearward rim end (92B). An airfoil (80) extends spanwise from each of the rotor disks (76). A first rotor disk (76) is disposed forward of a second rotor disk (76). The second rotor disk (76) includes an interstage shaft (106). The interstage shaft (106) includes an outer radial end (108), an inner radial end (110), and an arm (130). The arm (130) extending axially from an outer surface (124) of the interstage shaft (106) to a distal end (132). The rearward rim end (92B) of the first rotor disk (76) is fixedly joined to the distal end (132) of the arm (130), for example, by inertia welding.