RTX Corporation Patente

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 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 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.

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.

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) 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).

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.

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 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 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.

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.

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.

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 turbine engine assembly (86) includes a vane structure (88) and an inner structure (182). The vane structure (88) includes a plurality of structure segments (106) arranged circumferentially about an axis (32). Each of the structure segments (106) includes an inner platform segment (114), an outer platform segment (162) and an airfoil (177). The inner platform segment (114) includes a base (128) and a tab (130). The tab (130) is connected to and projects laterally out from the base (128). The tab (130) is radially next to and extends laterally along the base (128) of the inner platform segment (114) of a respective circumferentially neighboring one of the structure segments (106). The airfoil (177) extends radially across a flowpath (78) and is connected to the inner platform segment (114) and the outer platform segment (162). The inner structure (182) extends circumferentially around the axis (32). The inner structure (182) is configured with a plurality of pins (198). Each of the pins (198) projecting axially into an aperture (146) in the base (128) of a respective one of the structure segments (106).

A turbine engine (24) assembly includes an airfoil (102), a first baffle (104) and a second baffle (106). The first baffle (104) is disposed in a cavity (128) in the airfoil (102). The first baffle (104) extends spanwise from a first baffle base end (146) to a first baffle tip end (148) disposed at an airfoil tip end (110). The second baffle (106) is disposed in the cavity (128) longitudinally between and next to the first baffle (104) and an airfoil interior wall (121). The second baffle (106) extends spanwise from a second baffle base end (166) to a second baffle tip end (168) disposed at the airfoil tip end (110).

An assembly is provided for a turbine engine (20). This assembly includes a carrier (80), a control ring (82) and a bladed outer air seal (78) mounted to the carrier (80). The carrier (80) extends circumferentially around an axis (22). The carrier (80) includes an annular cavity (114), a first section (106) and a second section (108). The annular cavity (114) is formed within the carrier (80) by the first section (106) and the second section (108). The first section (106) is welded and/or brazed to the second section (108) at a sealed full-hoop bond joint (148; 150). The first section (106) is further connected to the second section (108) at a second sealed full-hoop joint (148; 150). The control ring (82) extends circumferentially around the axis (22) within the annular cavity (114). A method of manufacture is provided for an assembly for a turbine engine (20).

A seal assembly (60A, 60B; 160B) includes a carbon seal (64A, 64B; 164B) that has a sealing surface (66B, 66B; 166B). A seal seat (62A, 62B; 162B) has a sealing surface (66B, 68B; 168B) and is positioned for rotation relative to the carbon seal (64A, 64B; 164B). A diamond-like carbon coating (80) at least partially forms the sealing surface (68B, 68B; 168B) on the seal seat (62A, 62B; 162B).

A process for melt-infiltrating a partially densified ceramic matrix composite, comprising the steps of: disposing in a chamber of a thermal processing apparatus a partially densified ceramic matrix composite comprising a ceramic matrix including at least silicon carbide and at least one oxide phase precursor; introducing at least one gas into the thermal processing apparatus; elevating a temperature of the chamber to cause a reaction between the silicon carbide and the least one gas; achieving a chemical equilibrium of the gases within the chamber; identifying a gas regime for minimizing the liberation of SiO and SiO<2>; maintaining a temperature and CO<2>:CO ratio of the identified gas regime within the chamber; and, melt-infiltrating the partially densified ceramic matrix composite with the at least one oxide phase precursor to form a molten oxide ceramic matrix composite comprising a matrix including a silicon carbide phase and an oxide phase.

An aircraft propulsion system (20) assembly includes an open propulsor rotor (34) and a housing structure (54). The housing structure (54) includes an engine case (106), a nacelle structure (108) and a receptacle (126). The nacelle structure (108) is disposed radially outboard of and partially covers an engine case (106). The nacelle structure (108) forms a first axial section of an exterior flow boundary (56) for the aircraft propulsion system (20) that borders an environment (22) external to the aircraft propulsion system (20). The receptacle (126) is disposed radially outboard of the engine case (106) and axially between the open propulsor rotor (34) and the nacelle structure (108). The receptacle (126) is configured to receive a structure base (50) of an open guide vane structure (36).

An aircraft propulsion system includes aircraft systems having at least one hydrogen tank and an aircraft-systems heat exchanger and engine systems having at least a main engine core, a high pressure pump, a hydrogen-air heat exchanger, and a turbo expander (800). The main engine core includes a compressor section, a combustor section having a burner, and a turbine section. Hydrogen is supplied from the at least one hydrogen tank through a hydrogen flow path, passing through the aircraft-systems heat exchanger, the high pressure pump, the hydrogen-air heat exchanger, and the turbo expander (800), prior to being injected into the burner for combustion. The turbo expander (800) includes a rotor (804) separated into a first expander portion (812) and a second expander portion (814) arranged about an output shaft (806) and the output shaft (806) is operably connected to a generator configured to generate electrical power.

A high cycle fatigue test rig including a base; a first preload element attached to the base; a second preload element attached to the base; a test specimen attached to the base; the test specimen having a first end and a second end opposite the first end; the test specimen located in between the first preload element and the second preload element; the first preload element and the second preload element attached to the test specimen proximate the second end; a first clamp in operative communication with the first preload element; a second clamp in operative communication with the second preload element; an end cap in operative communication with the second end of the test specimen, the first preload element and the second preload element; and a dynamic force generator in operative communication with the end cap.

A system (60) for establishing components of a gas turbine engine (20) may include a constraints engine (62). The constraints engine (62) may be operable to determine a set of constraints (81) based on one or more input parameters (80) and selected values (83) for a group of interrelated design parameters (70) that may be associated with respective elements of an array. The design parameters (70) may correspond to respective components of a gas turbine engine (20). The constraints engine (62) may be operable to communicate the constraints (81) to a design tool (64). The design tool (64) may be operable to select values for the elements of the array from respective ranges of selectable values (82) within a design space (68) of the respective components based on a design model (66) associated with the design parameters (70). A method for establishing components of a system (60) is also disclosed.

A combustor wall (76) includes a panel (88), a boss (86), a wall aperture (90) and multiple cooling apertures (118A, 118B). The panel (88) extends axially along and circumferentially about an axial centerline (22). The panel (88) extends radially between a first panel surface (128) and a second panel surface (130). The first panel surface (128) forms a peripheral boundary of a combustion chamber (58). The boss (86) projects out from the second panel surface (130). The boss (86) extends circumferentially around and forms an outer peripheral boundary of the wall aperture (90). The wall aperture (90) extends along a wall aperture centerline (148) through the combustor wall (76) to the first panel surface (128). The cooling apertures (118A, 118B) are arranged circumferentially about the wall aperture (90). Each of the cooling apertures (118A, 118B) extends along a cooling aperture centerline (180, 192) through the panel (88) and/or the boss (86) to the wall aperture (90). The cooling aperture centerline (180) of a first of the cooling apertures (118A) is angularly offset from the wall aperture centerline (148) by a first acute angle (190).

A system for providing powered device operation on a rotating mechanism comprises static hardware (104) configured to remain in a fixed position and rotating hardware (102) configured to rotate about a central axis. Power generation circuitry associated with the static hardware (104) and the rotating hardware (102) is configured to generate a DC power signal on the rotating hardware (102) and output the DC power signal. A powered device (202) located on the rotating hardware (102) is configured to receive the DC power signal and perform a powered operation on the rotating hardware (102).

A combustor liner (100; 100-1, 100-2; 200-1; 300-1) for a combustor (56) is disposed about an axis (A) in a gas turbine engine (20) having an engine axis (A). The combustor liner (100...300-1) includes a first wall (106; 206; 306) comprising a first opening (110; 210; 310) therethrough, the first opening (110; 210; 310) configured to receive a cooling fluid, a second wall (108; 208; 308) spaced from the first wall (106; 206; 306), the second wall (108; 208; 308) configured to define a portion of a combustion chamber (104) and comprising a second opening (112; 212; 312), the second opening (112; 212; 312) configured to discharge the cooling fluid, a lattice structure (134; 234; 334) disposed between the first wall (106; 206; 306) and the second wall (108; 208; 308), a flow guide (116; 216; 316) defining a portion of the first opening (110; 210; 310) and connecting the first wall (106; 206; 306) to the second wall (108; 208; 308).

An aircraft powerplant (20) assembly includes a gearbox (88) and a power coupler (90). The gearbox (88) includes a first engine power transfer apparatus (120), a second engine power transfer apparatus (122) and an accessory power transfer apparatus (124). The power coupler (90) is mounted to the gearbox (88). During a first operating mode, the power coupler (90) is configured to operatively couple the first engine power transfer apparatus (120) to the accessory power transfer apparatus (124) and operatively decouple the second engine power transfer apparatus (122) from the accessory power transfer apparatus (124). During a second operating mode, the power coupler (90) is configured to operatively couple the second engine power transfer apparatus (122) to the accessory power transfer apparatus (124) and operatively decouple the first engine power transfer apparatus (120) from the accessory power transfer apparatus (122).

A system for indicating rotor to stator clash including a rotor including a forward section and an aft section opposite the forward section; a stator located near the rotor; and a clash indication feature attached to the rotor, the clash indication feature configured to contact the stator responsive to a predetermined deflection dimension of the stator.

A cryogenic fueled aircraft propulsion system includes a core engine (22) where fuel is mixed with compressed air and ignited to generate a high energy exhaust gas flow, a fuel system (24) that is configured to supply the fuel to the core engine (22), a cryogenic fuel tank (26) for storing fuel cryogenically in a liquid phase, a gas collection system (28) where boil off gas (36) from the fuel system (24) is collected, a fuel cell system (34) where a flow of collected boil off gas (36) is utilized to generate electric power (52), and an electrical coupling (48) where electric power (52) generated by the fuel cell is communicated to location outside of the aircraft (20).

A preform (200) for a ceramic matrix composite (CMC) component having a flow path surface, such as a blade outer air seal (BOAS), a blade, a vane, or a combustor panel, includes a base preform (200) includes a polyvinyl butyral (PVB) tackifier for dimensional stability and has a corresponding flow path surface. One or more machined channels (210) are formed on the corresponding flow path surface of the base preform (200), and may include additional features such as ribs, trip strips, or pin fins. One or more overwrap plies are then disposed on the base preform (200) and cover the one or more machined channels (210) to form a cooling feature. The overwrapped base preform may then be densified to provide a CMC component with a cooling feature on the flow path surface.

A rotor blade (58) for a gas turbine engine (20) is provided that includes an airfoil (68), an attachment section (64), and a neck section (66). The airfoil (68) includes a plurality of internal cooling air passages (90). The attachment section (64) has a base surface (70). A forward center cooling air passage (190AF) is open at the base surface (70) and extends through the attachment and neck sections (64). An aft center cooling air passage (190AA) is open at the base surface (70) and extends through the attachment and neck sections (64). A wall to wall, W2W, cooling air passage (190B) is open at the base surface (70) and extends through the attachment and neck sections (64). The W2W cooling air passage (190B) is disposed between the forward and aft center cooling air passages (190AF). In a transition in a direction from a cross-sectional plane disposed in the attachment section (64) to the airfoil root end (84), a side segment (194, 196) progressively extends outwardly from the W2W cooling air passage.

An aircraft powerplant (20) assembly includes a gearbox (102), a first electric machine (90A), a first electric cable (112A), a second electric machine (90B) and a second electric cable (112B). The first electric machine (90A) is mounted to the gearbox (102) at a first side (135A) of the gearbox (102). A first machine terminal (132A) is disposed at a first side (130A) of the first electric machine (90A). The first electric cable (112A) is electrically coupled to the first machine stator (96A) through the first machine terminal (132A). The second electric machine (90B) is mounted to the gearbox (102) at the first side (135) of the gearbox (102). A second machine terminal (132B) is disposed at a second side (130B) of the second electric machine (90B) that faces the first side (130A) of the first electric machine (90A). The second electric cable (112B) is electrically coupled to the second machine stator (96B) through the second machine terminal (132B).

An attachment assembly (100) for a ceramic matrix composite (CMC) vane (140) includes an inner diameter ring (110) having a strut recess (118) and a first flange (115) on a downstream side configured for engaging an inner platform (142) of the CMC vane (140), an outer diameter ring (120) having a support structure (125) disposed to form a gap (150) relative to a portion (148) of an outer platform (146) of the CMC vane (140) during a non-operating engine condition, and a strut (130) cantilevered from an inner side of the outer diameter ring (120) and disposed to pass through a hollow portion (148) within an airfoil (145) of the CMC vane (140) to engage the strut recess (118) in the inner diameter ring (110) to support the inner diameter ring (110) and the CMC vane (140). The inner and outer diameter rings and the strut (130) may be metallic and carry the loads rather than the CMC carrying the loads.

A method is described for introducing dormant cooling holes (120) into CMC components that have a base (200) made of a first group of plies (110) and a second group of plies (210). Before or after densification of a CMC preform, dormant cooling holes (120) are provided that pass through the first group of plies (110) but are covered by the second group of plies (210). The dormant cooling holes (120) become open upon erosion of the second plies (210). The opening of the dormant cooling holes (120) allows air to flow through the cooling holes (120) to the area of erosion to slow the erosion process.

A method for fabricating an interface coating on a substrate, comprising (i) providing a substrate; (ii) depositing on the substrate at least one layer of boron nitride; and (iii) depositing on the boron nitride layer a Si<3>N<4>-BN multilayer coating to form a coated substrate, the Si<3>N<4>-BN multilayer coating comprises a pattern of alternating layers having at least one boron nitride layer and at least one silicon nitride layer according to the following steps: (iiia) depositing at a first deposition temperature a layer of silicon nitride at a first thickness; (iiib) depositing at a second deposition temperature a layer of boron nitride at a second thickness; and repeating steps (iiia) and (iiib) until forming the Si<3>N<4>-BN multilayer coating. The coated substrate is a ceramic fibrous preform.

A gas turbine engine (22) for an aircraft propulsion system includes a turbine section (32), an exhaust section (34), an engine case (60), and a thermophotovoltaic (TPV) assembly (70). The turbine section (32) and the exhaust section (34) form a core flow path (66) along an axis (44) of the gas turbine engine (22). The engine case (60) is disposed within the exhaust section (34). The engine case (60) includes an inner liner (72) and an outer liner (74). Each of the inner liner (72) and the outer liner (74) extend circumferentially about the axis (44). The inner liner (72) forms the core flow path (66) through the exhaust section (34). The outer liner (74) is disposed radially outward of the inner liner (72) and forms an annular cavity (84) between the inner liner (72) and the outer liner (74). The TPV assembly (70) includes a plurality of TPV cells (86). The plurality of TPV cells (86) are disposed within the annular cavity (84).

An assembly for an aircraft powerplant (20) includes an engine case (72), a gearbox (102), a conduit (122), a first electric machine (92A), a first controller (94A) and a first electric cable (114A). The gearbox (102) is offset from the engine case (72). The conduit (122) extends longitudinally from a first end (124) of the conduit (122) to a second end (126) of the conduit (122). The conduit (122) is anchored to the engine case (72) at the first end (124) of the conduit (122) by a first conduit mount (136). The conduit (122) is anchored to the gearbox (102) at the second end (126) of the conduit (122) by a second conduit mount (138). The conduit (122) is self-supported longitudinally between the first conduit mount (136) and the second conduit mount (138). The first electric machine (92A) is mounted to the gearbox (102). The first controller (94A) is mounted to the engine case (72). The first electric cable (114A) electrically couples the first electric machine (92A) to the first controller (94A). The first electric cable (114A) extends longitudinally through a bore (134) of the conduit (122).

A gas turbine engine case structure (150) includes the unitary combination of: a case wall having an inner surface and an outer surface; at least one fuel inlet (171); circumferentially-distributed fuel injectors protruding inward from the case wall and having an outlet; a fuel plenum fluidically between the at least one fuel inlet (171) and the fuel injectors and configured so that each inlet (171) of the at least one inlet (171) is coupled to feed multiple of the fuel injectors; and a fuel inlet conduit (260) fluidically between the fuel inlet (171) and the fuel plenum. The inlet conduit (260) has a forward inlet and is held spaced from the case wall at multiple longitudinally-spaced locations.

A fuel delivery system for a combustor (300) of a gas turbine engine (20) has a primary fuel tank (518) configured to store liquid hydrogen (510) and a secondary fuel tank (558) configured to store fuel (550), such as Jet A or SAF. The fuel (510,550) being provided and mixed in the jet engine to alleviate the production of contrails. In addition, the primary fuel tank (518) may be a cryogenic tank (518) to enable the storage of liquid hydrogen, or liquid natural gas. As a result, the primary fuel (510) may be utilized to cool the secondary fuel (550) and/or the gas turbine engine (20).