RTX Corporation Patente

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).

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 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 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 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).

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.

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.

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.

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.

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 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 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 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 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 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 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).

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).

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.

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 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.

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.

An aircraft assembly includes a fan case (70), a core case (62), an accessory gearbox (104), one or more electric machines (94), a first electric device (120) and an electrical system (92). The core case is radially inboard of the fan case. The accessory gearbox is mounted with the fan case or the core case. The one or more electric machines are mounted to the accessory gearbox. The accessory gearbox is dedicated to the one or more electric machines. The first electric device is mounted with the fan case or the core case. The electrical system is electrically coupled to the one or more electric machines and the first electric device. The electrical system is configured to receive a first current of electricity from at least one of the one or more electric machines. The electrical system is configured to provide a second current of electricity to the first electric device.

Gas turbine engine (20) including a fan (28), compressor, combustion (32), and turbine sections, a fan bypass air path (38), a lubrication system (70), and a turbine active clearance control system (54). The turbine section has a rotor stage and a turbine case clearance control segment (46A). The turbine ACC system includes first (60) and second (64) valves and a heat exchanger (58). The heat exchanger may receive a first fan bypass air flow and a lubricant flow and permits heat transfer therebetween to produce a flow of conditioned air. The first valve receives conditioned air from the heat exchanger. The second valve may receive conditioned air and a second fan bypass air flow. The turbine ACC system may operate in a heating mode in which conditioned air flow, or a combination of conditioned air flow and the second fan bypass air flow is passed to the clearance control segment.

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. A second LLM system is 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 generator and compiler is configured to generate finite state machines. At least one finite state machine (FSM) generated by the generator and compiler is configured to analyze operation of the system design using each of the at least one FSM including optimization properties. A formal verification (FV) tool is configured to determine whether the system design passes or fails responsive to the first description of the system design in the temporal logic language.

A gas turbine engine (20) disposed within a nacelle (40) is provided that includes an open rotor propulsion system (28), compressor (30), combustion (32), and turbine (34) sections, a lubrication system (70), and a turbine active clearance control system (54). The turbine active clearance control system (54) includes first and second valves (60, 64), a heat exchanger (58), and a nacelle air inlet device (56). The heat exchanger (58) receives air from the air inlet device (56), accepts a lubricant flow therethrough, and permits heat transfer between the air flow and the lubricant to produce a conditioned air flow. The first valve (60) receives conditioned air flow from the heat exchanger (58) and may pass the conditioned air flow to the second valve (64). The second valve (64) receives a second air flow from the air inlet device (56).

An engine system (200) includes a turbine engine (202) including a compressor section (206), a combustor section (208) having a burner, a turbine section (210), and a nozzle (212) in an open-loop configuration. The engine system (200) also includes a bottom-cycle apparatus (204) and an exhaust heat exchanger (228) downstream of the turbine section (210) of the turbine engine (202) configured to reject heat from the turbine engine (202) to the bottoming-cycle apparatus (204) and create a cooled turbine exhaust (226) in the turbine engine (202). The engine system (200) further includes an exhaust compressor (214) arranged downstream of the exhaust heat exchanger (228) and upstream of the nozzle (212) of the turbine engine (202) configured to compress the cooled turbine exhaust stream (226) and increase a pressure of the cooled turbine exhaust stream (226) prior to exiting the nozzle (212) of the turbine engine (202).

A gas turbine engine (20) includes compressor (30) and turbine (34) sections, and a turbine cooling air (TCA) system (54). The turbine section (34) has upstream (44A) and downstream (44B) rotor stages, and turbine sub-sections (47A,47B). The TCA system (54) includes a flow mixing device (70) and a compressor bleed air flow valve (72), and is configured to receive first and second compressor bleed air flows. The compressor bleed air flow valve (72) may be in an open or a closed configuration. The TCA system (54) operates in a first or a second mode. In the first mode, the compressor bleed air flow valve (72) is closed and a conditioned air flow is from the first compressor bleed air flow. In the second mode, the compressor bleed air flow valve (72) is open and the conditioned air flow is a mixture of the first and second compressor bleed air flows. The conditioned air flow is directed to the downstream turbine subsection (47B).

An apparatus for performing formal verification of a system design, comprising 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 finite state machine (FSM) is configured to verify the system design responsive to the second description of the system design in the temporal logic. A formal verification (FV) tool is configured to indicate whether the system design passes or fails, wherein the FV tool is further configured to provide a verification coverage estimation of a model implemented by the FV tool using a finite elements method (FEM) triangulation mesh.

A gas turbine engine (20) disposed within a nacelle (40) includes an open rotor propulsion system (28), compressor (30), combustion (32), and turbine (34) sections, and a turbine active clearance control system (54). The turbine section has a turbine case clearance control segment (46A) disposed radially outside of a rotor stage (44). The turbine active clearance control system includes a bleed air mixing system (60), a nacelle air inlet device (56), and a compressor air bleed structure (58). The turbine active clearance control system is configurable in a turbine case heating mode in which the bleed air mixing system is used to mix an exterior air flow and a compressor bleed air flow to produce a conditioned air flow that is operable to heat the turbine case clearance control segment, and the turbine active clearance control system is configured to direct the conditioned air flow to the turbine case.

A composite vane with a locking feature including a leading edge and a trailing edge opposite chordwise from the leading edge; a radially inner attachment region opposite spanwise from a radially outer attachment region; a span dimension extending between the radially inner attachment region and the radially outer attachment region; a chord dimension extending between the leading edge and the trailing edge; a pressure side opposite a suction side of the composite vane; a dovetail section formed within the composite vane proximate the radially outer attachment region; a shoe member in operative communication with the dovetail section; a boot in operative communication with an inner portion opposite spanwise from the dovetail portion; and a failsafe fastener in operative communication with the boot and the inner portion of the composite vane.

A gas turbine engine (20) is provided that includes fan (28), compressor (30), combustion (32), and turbine (34) sections, a fan bypass air path (38), and a turbine active clearance control system (54). The turbine section has a rotor stage (44) and a turbine case clearance control segment (46A). The turbine active clearance control system includes a bleed air mixing system (60), a fan bypass air inlet device (56) configured to receive a fan bypass air flow, and a compressor air bleed structure (58) configured to receive a compressor bleed air flow. The turbine active clearance control system is configurable in a turbine case heating mode in which the bleed air mixing system mixes the fan bypass air flow and the compressor bleed air flow to produce a conditioned air flow that is operable to heat the turbine case clearance control segment.

A propulsion system (20) includes an open propulsor rotor (34), an open guide vane structure (36), a turbine engine (32) and a blade actuation system (48). The open propulsor rotor (34) includes a plurality of open propulsor blades (42). The turbine engine (32) includes a compressor section (65) and a rotating structure (86). The rotating structure (86) is operatively coupled to the open propulsor rotor (34). The rotating structure (86) includes a compressor rotor in the compressor section (65). The blade actuation system (48) is configured to adjust pitch of at least one of the open propulsor blades (42). A controller (110) is configured to control operation of the blade actuation system (48) during aircraft flight to tailor the pitch. The turbine engine (32) is configured to drive rotation of the open propulsor rotor (34) about the axis during a first mode of operation. The open propulsor rotor (34) is configured to windmill about the axis to drive rotation of the rotating structure (86) during a second mode of operation.

A propulsion system assembly includes an open propulsor rotor (34), an open guide vane structure (36) and a nose cone (38). The open propulsor rotor (34) is configured to rotate about an axis (24). The open guide vane structure (36) is aft of the open propulsor rotor (34) along the axis (24). The nose cone (38) is forward of the open propulsor rotor (34) along the axis (24). The nose cone (38) includes a forward tip end (108), an aft base end (110), a stationary section (116) and a rotating section (118). The aft base end (110) is adjacent to the open propulsor rotor (34). The stationary section (116) projects axially along the axis (24) from a junction (124) between the stationary section (116) and the rotating section (118) to the forward tip end (108). The rotating section (118) projects axially along the axis (24) from the junction (124) between the stationary section (116) and the rotating section (118) to the aft base end (110). The rotating section (118) is configured to rotate with the open propulsor rotor (34) about the axis (24).

An assembly is provided for an aircraft (20). This assembly includes a propulsion system (24) and a control system (138). The propulsion system (24) includes an open propulsor rotor (60) and a turbine engine (56) configured to drive rotation of the open propulsor rotor (60) about an axis. The control (138) system includes a first sensor (140), a second sensor (142) and a controller (144) in signal communication with the first sensor (140) and the second sensor (142). The first sensor (140) is configured to provide first sensor data indicative of a first environmental parameter. The second sensor (142) is mounted to the propulsion system (24) and is configured to provide second sensor data indicative of a second environmental parameter. The controller (144) is configured to monitor and/or control operation of the propulsion system (24) using the first sensor data. The controller (144) is also configured to monitor the first sensor data using the second sensor data.

A first core turbine rotor of a first propulsion system (24A) is disposed in a first core turbine section (89) and is configured to rotate in a first rotational direction. A first power turbine rotor (106) is disposed in a first power turbine section (90) and is configured to rotate in the first rotational direction. The first power turbine rotor is operatively coupled to a first open propulsor rotor (58). A second core turbine rotor of a second propulsion system (24B) is disposed in a second core turbine section and is configured to rotate in the first rotational direction. A second power turbine rotor is disposed in a second power turbine section and is configured to rotate in a second rotational direction that is opposite the first rotational direction. The second power turbine rotor is operatively coupled to a second open propulsor rotor (58).

A shaft assembly (95) includes a shaft (100) that will be rotating in use. The shaft (100) receives a seal runner (102), and a locking nut (101). The seal runner (102) has a portion secured within an inner peripheral bore (106) of the shaft (100), and the locking nut (101) is mounted on an outer peripheral surface (98) of the shaft (100). The seal runner (102) further has a flange (108) axially beyond the shaft (100). A key washer (113) has non-rotation structure (114) for preventing rotation of the key washer (113) relative to the locking nut (101). A retaining ring (118) is mounted in a channel (116) in the key washer (113), and extends into a groove (120) in the seal runner (102). A gas turbine engine (20) and a method are also disclosed.

An assembly is provided for an aircraft (20). This assembly includes an open propulsor rotor (60), a turbine engine (58), an accessory gearbox (152), a first electric machine (142A), a second electric machine (142B), an aircraft system (24) and an electrical system (140). The turbine engine is configured to drive rotation of the open propulsor rotor about an axis (50). The accessory gearbox is mounted with the turbine engine. The first electric machine is operatively coupled to the turbine engine through the accessory gearbox. The second electric machine is operatively coupled to the turbine engine through the accessory gearbox. The aircraft system is configured as an actuation system, a fuel system, a lubrication system and/or a cooling system. The aircraft system includes a plurality of electric devices. The electrical system is electrically coupled to the first electric machine, the second electric machine and each of the electric devices.

An integrated mount structure (42) including a foundation section (46) located adjacent to a shoulder section (40) of an aft case section (30) of an engine (10); a column section (50) adjacent the foundation section, the column section is configured to contact a vehicle (44) and couple the engine with the vehicle; and an internal hollow (58) formed in both the column section and the foundation section.