RTX Corporation Patente

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.

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.

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).

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 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 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 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.

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) 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 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.

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 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 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 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.

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.

Methods and devices are provided for performing ultrasonic shot peening (USP) on a component of an engine. A USP head unit (300) is configured for performing USP. The USP head unit (300) includes a treatment enclosure (302) containing peening media. USP is performed on the component using the USP head unit (300). The USP head unit (300) also includes a peening block (1202) coupled to the treatment enclosure (302), and a guide window (902) extending from within the treatment enclosure (302), through the peening block (1202) and to the component. Performing USP includes vibrating the peening media within the treatment enclosure (302) and the guide window (902) at an ultrasonic frequency such that the peening media impacts against the component from within the guide window (902).

An assembly (30) is provided for an aircraft propulsion system (20). This assembly (30) includes an open propulsor rotor (34) and an open guide vane structure (36). The open propulsor rotor (34) is configured to rotate about an axis (24). The open propulsor rotor (34) includes a plurality of open propulsor blades (42) arranged circumferentially about the axis (24) in a blade array. Each of the open propulsor blades (42) projects spanwise into an environment (22) external to the aircraft propulsion system (20) to a respective blade tip (46). The open propulsor blades (42) include a first span propulsor blade (42A) and a second span propulsor blade (42B). The first span propulsor blade (42A) has a first span length (48A). The second span propulsor blade (42B) has a second span length (48B) that is less than the first span length (48A). The open guide vane structure (36) is axially next to the open propulsor rotor (34).

A tail cone ventilation system (10) including a tail cone case (16) defining a tail cone interior (22) and a tail cone exterior (24), the tail cone having a forward portion (44) and an aft portion (30) separated axially along a tail cone axis; an air inlet (38) fluidly coupled with the tail cone interior through ducting, the air inlet located externally from the tail cone interior; a conduit (53) attached to the tail cone case at the tail cone exterior and to a pylon (54) at an opposite end, the conduit comprising a cable chase (56) having cable (52), the cable being in operative communication with at least one electronic component (12) within the tail cone interior; a cooling supply fluidly coupled between the air inlet and a pylon exhaust (68), the cooling supply configured to cool the cable within the cable chase; and a fan (66) fluidly coupled with the cooling supply.

A component (10) includes a wall (12) and a passage (14). The wall (12) bounds an interior region (20) and extends along an axis (A) from a first end (10A) to a second end (10B) of the component (10). The passage (14) extends within the wall (12) between the first end (10A) and the second end (10B). The passage (140) includes an undulating profile comprising a plurality of undulations and extends along an extension axis. The undulating profile includes a first period measured along the extension axis between a first pair of adjacent undulations that is different from a second period measured along the extension axis between a second pair of adjacent undulations.

TURBINE AIRFOIL COATINGA method for coating a metallic substrate includes applying an MCrAlY coating. Machining removes the MCrAlY coating from one or more regions of the substrate. A simultaneous aluminizing and chromizing: aluminizes an interior surface region of the substrate lacking the MCrAlY and at least a portion of a region where the MCrAlY remains; and chromizes an exterior surface region of the substrate lacking the MCrAlY and at least a different portion of the region where the first MCrAlY remains.

An aircraft assembly includes an airframe, a propulsion module (56) and a turbine engine (58). The airframe includes a body, a first wing and a second wing. The propulsion module (56) is mounted to the body. The propulsion module (56) includes a first open propulsor rotor (60). The turbine engine is configured to drive rotation of the first open propulsor rotor (60) about a propulsion module axis (64). The turbine engine (58) includes a flowpath (92), a compressor section (88), a combustor section (89), a turbine section (90) and a first rotating structure (104A,104B). The flowpath (92) extends through the compressor section (88), the combustor section (89) and the turbine section (90) from an inlet (124) into the flowpath to an exhaust from the flowpath (126). The first rotating structure (104A,104B) includes a bladed engine rotor disposed in the compressor section (88) or the turbine section (90). The first rotating structure (104A,104B) is rotatable about a rotating structure axis (106A,106B) that is offset from the propulsion module axis (64).

An aircraft assembly includes an engine core (48), a first electrical power system (34A) and a second electrical power system (34B). The first electrical power system (34A) includes a first electric machine (96A) and a first electric machine controller (98A). The first electric machine (96A) is operatively coupled to the engine core (48). The first electric machine controller (98A) is electrically coupled to the first electric machine (96A). The first electric machine controller (98A) is configured to control operation of the first electric machine (96A). The second electrical power system (34B) is electrically discrete from the first electrical power system (34A). The second electrical power system (34B) includes a second electric machine (96B) and a second electric machine controller (98B). The second electric machine (96B) is operatively coupled to the engine core (48). The second electric machine controller (98B) is electrically coupled to the second electric machine (96B). The second electric machine controller (98B) is configured to control operation of the second electric machine (96B).

An aspect includes a hybrid gas turbine engine system (201; 301) of a hybrid electric aircraft (200; 300). The hybrid gas turbine engine system includes a gas turbine engine (20), an electric motor (312) operable to perform an electric taxiing of the hybrid electric aircraft, and a controller (216). The controller is operable to prevent fuel flow to the gas turbine engine (20) during at least a portion of the electric taxiing and monitor for a powered warm-up request during the electric taxiing. A powered warm-up state (405) of the gas turbine engine (20) is initiated based on detecting the powered warm-up request. The powered warm-up state (405) adds heat to one or more components of the gas turbine engine prior to transitioning to a takeoff power state. The gas turbine engine transitions from the powered warm-up state to the takeoff power state after reaching a target temperature (408) of the one or more components in the powered warm-up state.

An airframe (22) includes a body (26), a first wing (28A), a second wing (28B) and a vane (54). The body extends longitudinally along a longitudinal centerline (34) from a forward end (36) of the body to an aft end (38) of the body. The first wing and the second wing are disposed to opposing lateral sides of the body. The vane projects vertically out from the body to a tip of the vane. The vane extends longitudinally along the body from a leading edge (66) of the vane to a trailing edge (68) of the vane. A first open rotor propulsion system (24A) is vertically above and mounted to the body. A second open rotor propulsion system (24B) is vertically above and mounted to the body. The vane is configured as a structural barrier laterally between the first open rotor propulsion system and the second open rotor propulsion system.

An apparatus for performing formal verification of a system design comprises a first large language model (LLM) system configured to receive a natural language specification of the system design and generate a first description of the system design in a temporal logic language and a second LLM system configured to receive the natural language specification of the system design and generate a second description of the system design in the temporal logic language. A first finite state machine (FSM) is configured to generate verification properties for the system design responsive to the first description of the system design in the temporal logic language. A second FSM configured to generate verification drivers for the system design. A formal verification (FV) tool configured to determine whether a model of the system design passes or fails. The FV tool is further configured to generate a property witness template and an optimal witness template.

A case mounted bumper pad including a case contact region supporting a runner pad support; the case contact region is in operative communication with the case, the runner pad support is in operative communication with a runner pad, the runner pad is in operative communication with a sync ring of a variable vane; and a gap formed in the bumper pad proximate the case.

A cover plate (140) for enclosing a cooling cavity (130) of a ceramic matrix composite (CMC) component (100) of a turbine engine may include a base plate (141) dimensioned to cover an open cooling cavity (130) of the CMC component (100). The base plate (141) has a first side facing the cooling cavity (130) and at least one feature plate (145) is attached to the first side of the base plate (141). At least one heat transfer augmentation feature (145) is formed on the at least one feature plate (142) and extends into the cooling cavity (130), wherein the at least one heat transfer augmentation feature (145) is configured to cause turbulence in a cooling air flow within the cooling cavity (130). Thus, rather than forming complex heat transfer augmentation features (145) on the cooling cavity (130) of the CMC component (100) itself, such features can be easily provided using metal plate or sheet material or separately-formed CMC elements on the cover plate (141).

A method of manufacture is provided during which a component (20') with an external artifact (54) and an internal artifact (42) is scanned using a computed tomography machine (58) to provide scan data. The internal artifact (42) is disposed within an interior of the component (20'). External artifact data and internal artifact data are determined based on the scan data. The external artifact data is indicative of a physical characteristic of the external artifact (54). The internal artifact data is indicative of a physical characteristic of the internal artifact (42). A first external coordinate system is registered for the component (20') based on the external artifact data. An internal reference feature is registered for the component (20') based on the internal artifact data. Manufacturing data is provided for the component (20'). The internal reference feature is located within the first external coordinate system during the providing of the manufacturing data. A feature is formed into the component (20') using the manufacturing data.

A method (200) includes determining a current burn flow value of a turbine engine (205), measuring a fuel pressure of a metered fuel flow of liquid fuel to a combustor of the turbine engine at a measurement point in a fuel system upstream of the combustor (210) and determining a current burner pressure as a function of the measured fuel pressure and the current burn flow value (215).

An aircraft powerplant (20) assembly includes a gearbox (88) having a first power transfer apparatus (120) and a second power transfer apparatus (122) independent of the first power transfer apparatus (120). A power coupler (90) includes a clutch (164). The first power transfer apparatus (120) is configured to operatively couple and transfer mechanical power through the gearbox (88) between a first engine apparatus (134) and an engine accessory (92). The power coupler (90) is configured to operatively couple a power coupled accessory (96) to the second power transfer apparatus (122) when the clutch (164) is engaged. The second power transfer apparatus (122) and the power coupler (90) are collectively configured to operatively couple and transfer mechanical power through the gearbox (88) between a second engine apparatus (122) and the power coupled accessory (96) when the clutch (164) is engaged. The power coupler (90) is configured to operatively decouple the second power transfer apparatus (122) from the power coupled accessory (96) when the clutch (164) is disengaged.

A vane structure (140) includes a structure passage (142) that extends internally from a passage inlet (160) to a passage outlet (162). The passage inlet (160) fluidly couples the structure passage (142) to an external environment outside of the vane structure (140). The first heat exchanger (134A) is disposed within the vane structure (140) along the structure passage (142) between the passage inlet (160) and the passage outlet (162). The second heat exchanger (134B) is disposed within the vane structure (140) along the structure passage (142) between the passage inlet (160) and the passage outlet (162). The electric machine system (32) includes a plurality of electric components including a first electric machine (96A) and a first electric machine controller (108A) for the first electric machine (96A). The first electric machine (96A) is configurable as an electric motor and/or an electric generator. The first fluid circuit (128A) is configured to cool and/or lubricate a first of the electric components. The first fluid circuit (128A) includes the first heat exchanger (134A) and/or the second heat exchanger (134B).

A method for inspecting a component (66) for an aircraft propulsion system (22, 68) with an impedance-measuring inspection system (64) includes, for each of a plurality of different positions on the component (66), positioning a probe assembly (70) of the impedance-measuring inspection system (64) at one of the plurality of different positions with a first electrode (80) and a second electrode (82) of the probe assembly (70) contacting the component (66), measuring an impedance of the component (66) between the first electrode (80) and the second electrode (82) over a frequency range extending between a first alternating current (AC) frequency and a second AC frequency, and determining an impedance differential (112) of the impedance over the frequency range. The method further includes identifying a presence or an absence of a defect condition of the component (66) using an impedance differential profile (118, 120) for the component (66) including the impedance differential (112) at each of the plurality of different positions on the component (66).

An apparatus comprising a first large language model (LLM) system configured to receive a natural language specification of a system design and generate a first description of the system design in a temporal logic language responsive to the natural language specification. A second LLM system is configured to receive the natural language specification and generate a second description of the system design in the temporal logic language. A first finite state machine (FSM) is configured to generate verification properties for the system design responsive to the first description of the system design in the temporal logic language. A second FSM is configured to generate verification drivers for the system design responsive to the first description of the system design. A formal verification (FV) tool is configured to indicate whether the system design passes or fails responsive to the first description of the system design in a temporal logic language.

A fan blade fairing assembly (68) for a gas turbine engine (24) includes a fairing body (74), a replaceable component (132B), and a bonding assembly (128). The fairing body (74) includes an outer wall (82), a first sidewall (84), and a second sidewall (86). The outer wall (82) extends between and connects the first sidewall (84) and the second sidewall (86). The replaceable component (132B) is bonded onto the fairing body (74). The bonding assembly (128) includes an electrically debonding adhesive (130), a first conductive member (134), and a second conductive member (136). The electrically debonding adhesive (130) bonds the replaceable component (132B) onto the fairing body (74). The electrically debonding adhesive (130) is disposed between the first conductive member (134) and the second conductive member (136).

A method for forming an access passage (168) for a gas turbine engine (24) of an aircraft propulsion system (22) includes installing a tooling jig (144) on an engine case (70) of the gas turbine engine (24) at a port (88) at a turbine section (34) radially outward of a turbine vane stage (72), aligning the tooling jig (144) on the engine case (70) and fixedly mounting the tooling jig (144) on the engine case (70), and drilling a first access aperture (172) of the access passage (168) through the inner ring (92) along the tooling axis (152) with a drill bit (180) inserted into the first turbine vane through a tooling aperture (150) of the tooling jig (144). The access passage (168) extends through the turbine vane stage (72) to a rotor cavity (86) disposed radially inward of the turbine vane stage (72) and between an upstream rotor stage (66) and a downstream rotor stage (68) of a bladed turbine rotor.

A system for an aircraft includes an open propulsor rotor (46), a turbine engine (24) and an attachment (98). The turbine engine (24) is configured to drive rotation of the open propulsor rotor (46) about an axis (32). The turbine engine (24) includes an engine case (64). The attachment (98) is configured to mount the turbine engine (24) to a pylon structure (88) and transfer one or more loads between the turbine engine (24) and the pylon structure (88). The attachment (98) is configured to mount to the pylon structure (88) through a pylon pin connection (160) such that the attachment (98) is operable to pivot about the pylon pin connection (160) and move relative to the pylon structure (88). The attachment (98) includes a first link (102A) and a second link (102B). The first link (102A) is mounted to the engine case (64) through a first link-case pin connection (146A). The second link (102B) is mounted to the engine case (64) through a second link-case pin connection (146B).

An aircraft system includes a turbine engine case and an attachment that is configured to mount the turbine engine case (64) to a pylon structure (88) and transfer one or more loads between the turbine engine case (64) and the pylon structure (88). The attachment includes a bracket (100), a first link (102A) and a second link (102B). The bracket is configured to mount to the pylon structure (88) through a bracket-pylon pin connection (160) such that the attachment (98) is operable to pivot about the bracket-pylon pin connection (160) and move relative to the pylon structure (88). The bracket is mounted to the first link (102A) through a bracket-first link pin connection (144A). The bracket is mounted to the second link (102B) through a bracket-second link pin connection (144B). The first link (102A) is mounted to the turbine engine case (64) through a first link-case pin connection (146A). The second link (102B) is mounted to the turbine engine case (64) through a second link-case pin connection (146B).

An assembly is provided for an aircraft propulsion system (20). This assembly includes an open propulsor rotor (34) and a nose cone (38). The open propulsor rotor (34) is configured to rotate about an axis (24). The nose cone (38) is upstream of and is configured to rotate with the open propulsor rotor (34). The nose cone (38) extends axially along the axis (24) from a forward tip end (108) of the nose cone (38) to an aft base end (110) of the nose cone (38) that is adjacent the open propulsor rotor (34). The nose cone (38) includes an air scoop (140) arranged at an intermediate location (130) between the forward tip end (108) of the nose cone (38) and the aft base end (110) of the nose cone (38). The air scoop (140) is configured to direct air from an environment (22) external to the aircraft propulsion system (20) into an interior of the nose cone (38).