RTX Corporation Patente

An aircraft propulsion system (20) assembly includes a fan rotor (44), an inner flowpath (40), an outer flowpath (42) and an inter-flowpath structure (122). The fan rotor (44) includes a plurality of fan blades (48) arranged circumferentially about an axis (26). The outer flowpath (42) includes an outer flowpath inlet (94) downstream of the fan rotor (44). The outer flowpath inlet (94) is radially outboard of an inner flowpath inlet (90). The inter-flowpath structure (122) is disposed radially between and partially forms the inner flowpath inlet (90) and the outer flowpath inlet (94). The inter-flowpath structure (122) includes a splitter (128) and an actuator (142). The splitter (128) forms a leading edge (134) of the inter-flowpath structure (122). The actuator (142) is configured to translate the splitter (128) axially along the axis (26) between a first position and a second position.

An assembly includes a compressor section (33), a core flowpath (40), a bypass flowpath (42), a flowpath wall and an air circuit (124). The flowpath wall forms an inner peripheral boundary of the bypass flowpath (42). The flowpath wall includes a wall base (92) and a wall protrusion (94). The wall base (92) extends axially along and circumferentially about the axis. The wall protrusion (94) projects radially out from the wall base (92). The wall protrusion (94) extends axially along the axis from a protrusion first end to a protrusion second end. The protrusion first end is upstream of the protrusion second end along the bypass flowpath (42). The air circuit (124) extends from a circuit inlet (126) to a circuit outlet (128). The circuit inlet (126) is fluidly coupled to the core flowpath (40). The circuit outlet (128) is fluidly coupled to the bypass flowpath (42). The circuit outlet (128) is formed in the wall protrusion (94) axially between the protrusion first end and the protrusion second end.

An assembly (100) for a gas turbine engine (20), the assembly (100) comprising: a CMC vane (105) having a CMC vane inner diameter platform (145), wherein the CMC vane inner diameter platform (145) includes an upper sealing interface surface (144a) comprising an upper surface (140a) of a groove (140); one of a tangential onboard injector (110) or an inner air seal outer diameter, wherein said one of the tangential onboard injector (110) or the inner air seal outer diameter includes a lower sealing interface surface (144b) comprising a lower surface (140b) of the groove (140), wherein the CMC vane inner diameter platform (145) and said one of the tangential onboard injector (110) or the inner air seal outer diameter are disposed in contact to form an interface (144) comprising the upper sealing interface surface (144a) and the lower sealing interface surface (144b); and a seal (135; 200; 300; 400) disposed within the groove (140) formed by the upper surface (140a) and the lower surface (140b).

A system for sensing clearance with a turbine within a gas turbine engine (20) comprising a controller (506) configured to transmit an electromagnetic wave signal and receive a reflected electromagnetic wave signal. A waveguide (502) configured to propagate the electromagnetic wave signal from the controller (506) toward the turbine and propagate the reflected electromagnetic wave signal to the controller (506). A blade outer air seal (BOAS) (66) defining a window (608) therein through which the electromagnetic wave signal passes and through which the reflected electromagnetic wave signal passes.

A hybrid-electric aircraft propulsion system (20) includes an engine (22), a first electric machine assembly (80), and a second electric machine assembly (82). The engine (22) includes a first rotational assembly and a second rotational assembly. The first electric machine assembly (80) includes a first electric machine (86), a first control unit (88), first electrical cables (94), and a first cable conduit (96). The first electric machine (86) is coupled with the first rotational assembly. The first electrical cables (94) extend between and electrically connect the first electric machine (86) and the first control unit (88) through the first cable conduit (96). The second electric machine assembly (82) includes a second electric machine (90), a second control unit (92), second electrical cables (98), and a second cable conduit (100). The second electric machine (90) is coupled with the second rotational assembly. The second electrical cables (98) extend between and electrically connect the second electric machine (90) and the second control unit (92) through the second cable conduit (100).

A replaceable tool tip system (200) for an ultrasonic impact grinder includes a tool (230) having a proximal end adjacent a horn (232) of the ultrasonic impact grinder and a distal end. At least one threaded blind hole (386) is disposed in the distal end of the tool. A replaceable tool tip (235) includes at least one threaded through hole (384) in registration with each threaded blind hole, wherein a first threading of each threaded blind hole is different from a second threading of each corresponding threaded through hole. At least one headless differential screw fastener (280) has first threads (381) for engaging a respective threaded blind hole on a first end and different second threads (383) for engaging a respective through hole on a second end, and is used to attach and preload the replaceable tool tip.

A method of building a component using additive manufacturing techniques includes depositing and sintering one or more layers of build powder using a PBF-L process to form an intermediate consolidated build powder structure having a sintered surface. The PBF-L process is paused and a laser peening process is performed in situ on the sintered surface to induce a desired residual stress profile in selected layers of the intermediate consolidated build powder structure. The PBF-L process is resumed to deposit and sinter one or more layers of additional build powder onto the intermediate consolidated build powder structure to form an additional intermediate consolidated build powder structure having an additional sintered surface. The steps are repeated a predetermined number of times to form a finished component. The complete component includes one or more zones of residual stress imparted by the laser peening process in one or more interior layers.

A propulsion system (20) includes an electrical assembly (26). The electrical assembly (26) includes an electric machine (86), a control unit (88), and a thermal management assembly (84). The control unit (88) is electrically connected to the electric machine (86). The thermal management assembly (84) includes an electric machine cooling circuit, a control unit (CU) cooling circuit, and a pump assembly. The electric machine cooling circuit is connected in fluid communication with the electric machine (86). The CU cooling circuit is connected in fluid communication with the control unit (88). The pump assembly includes a pump shaft, a first pump rotor, and a second pump rotor. The first pump rotor and the second pump rotor are mounted on the pump shaft. The first pump rotor forms a coolant pump of the electric machine cooling circuit. The second pump rotor forms a coolant pump of the CU cooling circuit.

A hybrid-electric aircraft propulsion system includes a gas turbine engine, an electrical assembly, and a gear train (60). The gas turbine engine includes a first rotational assembly (40) rotatable about a rotational axis of the gas turbine engine. The electrical assembly includes a plurality of electric machines (70). The plurality of electric machines (70) includes a first electric machine (76) and a second electric machine (78). The gear train (60) includes a first shaft (84), a second shaft (92), a third shaft (96), a first gear (88), a second gear (98), and a transmission assembly (94). The first shaft (84) couples the first electric machine (76) with the first rotational assembly (40). The second shaft (92) is coupled with the second electric machine (78). The second shaft (92) is selectively coupled with the third shaft (96) by the transmission assembly (94). The first gear (88) is mounted on the first shaft (84). The second gear (98) is mounted on the third shaft (96) and engaged with the first gear (88).

A method for controlling clearance (68) between a blade outer air seal (BOAS) (66) and a rotating portion (64) of a gas turbine engine (20) comprising detecting an operating mode of the gas turbine engine (20), monitoring a current clearance between the BOAS (66) and the rotating portion (64) of the gas turbine engine (20) and adjusting a position of the BOAS (66) to alter the clearance from the current clearance to a baseline clearance associated with the detected operating mode.

A method for developing a mistuning model to facilitate dispositions of potential repair blend profiles on damaged blades of a bladed rotor can include collecting a vibration data, the vibration data comprising blade-alone frequency for each of a plurality of blades in the bladed rotor on a gas-turbine engine; compiling the vibration data with vibration data sets across a fleet of gas-turbine engines to generate a vibration fleet data; calculating a system vibration amplitude as a function of frequency shift for each of a plurality of system modes based on the vibration fleet data to generate a frequency shift impact on the system vibration amplitude; and developing, based on the frequency shift impact, the mistuning model, the mistuning model configured to receive a blade-alone frequency parameter as an input perturbation. The developed mistuning model can be used for potential repair blend dispositions.

An ion gas injection and reclamation system (10) including a heat exchanger (12) fluidly coupled with at least one flow passage; a re-valence/recirculator (36) fluidly coupled with the at least one flow passage (18, 24); an ion fluid reclaimer fluidly coupled with the re-valence/recirculator; and an injection gas comprising at least one ion fluid (18) constituent fluidly coupled with at least one of the re-valence/recirculator, the ion gas reclaimer and the at least one flow passage.

An aircraft system is provided that includes an open propulsor rotor, a turbine engine and a power turbine electric machine. The turbine engine includes a core compressor section, a core combustor section, a core turbine section, a power turbine section, a flowpath, a first core rotating structure and a power turbine rotating structure. The first core rotating structure includes a first core compressor rotor and a first core turbine rotor. The first core compressor rotor is disposed in the core compressor section. The first core turbine rotor is disposed in the core turbine section. The power turbine rotating structure is configured to drive rotation of the open propulsor rotor. The power turbine rotating structure includes a power turbine rotor disposed in the power turbine section. The power turbine electric machine is operatively coupled to the power turbine rotating structure.

An air-cooled turbine component includes a body (90) extending between and to a first radial surface (112) and a second radial surface (114) relative to a centerline axis (74). The body (90) includes an outer passage wall (126) and an inner passage wall (128) forming an internal cooling passage (116) between the first radial surface (112) and the second radial surface (114). The internal cooling passage (116) extends lengthwise through the body (90) between and to an upstream end (120) of the internal cooling passage and a downstream end (122) of the internal cooling passage along a lengthwise axis (118) of the internal cooling passage. The body (90) forms a flow augmentation feature assembly (134) at the inner passage wall (128) within the internal cooling passage (116). The flow augmentation feature assembly (134) includes a plurality of trip strips (136) arranged in a tread pattern along the inner passage wall (128).

An assembly is provided for an aircraft. This assembly includes an aircraft propulsion system (20) and a propulsion system pylon (108). The aircraft propulsion system (20) includes a turbine engine (32) and an open propulsor rotor (34). The turbine engine (32) is configured to drive rotation of the open propulsor rotor (34). The propulsion system pylon (108) is configured to mount the aircraft propulsion system (20) to an airframe (106) of the aircraft. The propulsion system pylon (108) includes an engine pylon structure (114) and an airframe pylon structure (112). The engine pylon structure (114) is configured to connect the aircraft propulsion system (20) to the airframe pylon structure (112). The engine pylon structure (114) is attached to the turbine engine (32) by a plurality of engine-pylon mounts (140, 152). The airframe pylon structure (112) is configured to connect the engine pylon structure (114) to the airframe (106) of the aircraft. The airframe pylon structure (112) is removably attached to the engine pylon structure (114).

A combustor system for an aerial vehicle can include an outer case (212; 612; 712) with a first and second end (214, 216; 614, 616), an inner combustor case (218; 618; 714) with a lumen (222; 622), a central combustor chamber (220; 620), an air flow path (224; 624; 716) between the outer case (212; 612; 712) and inner combustor case (218; 618; 714), an integrated air scoop (236; 632; 720) attached to the inner combustor case (218; 618; 714) at the lumen (222; 622) that directs air from the air flow path (224; 624; 716) into the central combustor chamber (220; 620), a wedge (724) within the integrated air scoop (236; 632; 720), a second integrated air scoop (634; 722) attached to the outer case (212; 612; 712), a bifurcated orifice (726) integrated into the second integrated air scoop (634; 722), a fuel manifold (226) and fuel orifice (228; 718) within the first integrated air scoop (236; 632; 720), and a splash plate (230; 630) with a first and second face (234), the splash plate (230; 630) extending through the lumen (222; 622) into the central combustor chamber (220; 620). The first and second integrated scoops (236; 632, 634; 720, 722) can interlock and seal by pushing the wedge (724) into the bifurcated orifice (726).

A combustor system for an aerial vehicle includes a diffuser case (212) with a first and second end (214) and a large and small radius section, an inner diffuser wall contained within the diffuser case (212), the inner diffuser wall (218) with a lumen (222), a central combustor chamber (220), a first air flow path (224) extending between the diffuser case (212) and inner diffuser wall (18), an integrated air scoop (236) attached to the inner diffuser wall (218) at the first lumen (222) that directs air from the first air flow path (224) into the central combustor chamber (220), a fuel manifold (226) and fuel orifice (228) within the integrated air scoop (236), and a splash plate (230) comprising with a first and second face (232, 234). The splash plate (230) extends through the lumen (222) into the central combustor chamber (220). The large radius section is aligned with the inner diffuser wall (218) away from the splash plate (230) and the small radius section is aligned with the splash plate (230).

A blade for a gas turbine engine (22) includes an airfoil body (68). The airfoil body (68) includes a pressure side wall (76), a suction side wall (78), and a tip end wall (130). The tip end wall (130) forms a blade tip (70). The airfoil body (68) forms a plurality of suction side wall passages (91) and a main body cavity (88). The suction side wall (78) includes an exterior wall segment (92) and an interior wall segment (94). The suction side wall (78) forms the plurality of suction side wall passages (91) between the exterior wall segment (92) and the interior wall segment (94). The interior wall segment (94) and the exterior wall segment (92) extend to the tip end wall (130). The plurality of suction side wall passages (91) extend through the tip end wall (130) to the blade tip (70). The interior wall segment (94) and the pressure side wall (76) form the main body cavity (88).

A pylon for an aircraft projects spanwise along a span line out from a base end towards a first mounting station and a second mounting station. The first mounting station is disposed a first spanwise distance along the span line from the base end. The second mounting station is disposed a second spanwise distance along the span line from the base end. The second spanwise distance is greater than the first spanwise distance. The pylon is configurable to mount to: a portion of a first propulsion system at the first mounting station, wherein the first propulsion system comprises a first propulsor rotor with a first diameter; or a portion of a second propulsion system at the second mounting station, wherein the second propulsion system comprises a second propulsor rotor with a second diameter that is greater than the first diameter.

A process for fabricating a melt-infiltrated ceramic matrix composite, comprising the steps of fabricating a ceramic fiber preform comprising fibers, fiber tows, or both fibers and fiber tows (20, 30); depositing a structural support material on the fibers, the fiber tows, or the fibers and fiber tows to form a partially densified ceramic matrix composite comprising pores, a porous network or both pores and a porous network; preparing a slurry composition comprising sacrificial particles (10); infiltrating the slurry composition into the partially densified ceramic matrix composite and depositing the sacrificial particles into the pores, the porous network or both the pores and porous network; freeze-drying a slurry infiltrated partially densified ceramic matrix composite; sublimating a frozen slurry infiltrated partially densified ceramic matrix composite; and, melt-infiltrating a sublimated partially densified ceramic matrix composite to form a melt-infiltrated ceramic matrix composite.

An assembly includes an engine core (30), a first propulsor rotor (22), a first power coupler (90) and a drive unit (32). The engine core (30) includes a compressor section (66), a combustor section (67), a turbine section (68), a flowpath (128) extending therethrough and a first rotating structure (88). The first rotating structure (88) includes a first turbine rotor (80) disposed in the turbine section (68). The first power coupler (90) is configured to operatively couple the first rotating structure (88) to the first propulsor rotor (22) during a first mode such that the first rotating structure (88) is configured to drive rotation of the first propulsor rotor (22) through the first power coupler (90). The first power coupler (90) is also configured to operatively decouple the first rotating structure (88) from the first propulsor rotor (22) during a second mode. The drive unit (32) is configured to drive the rotation of the first propulsor rotor (22) during the second mode.

An aircraft system (20) includes an open propulsor rotor (38), an engine core (94), an engine flowpath (36) and a flow regulator (146). The open propulsor rotor (38) includes a plurality of open propulsor blades (46) arranged circumferentially about an axis (24). Each of the open propulsor blades (46) is disposed in an external environment (22). The engine core (94) includes a compressor section (89), a combustor section (90) and a turbine section (91). The engine flowpath (36) includes an end flowpath inlet (134), a side flowpath inlet (136) and a flowpath exhaust (138). The engine flowpath (36) extends longitudinally through the compressor section (89), the combustor section (90) and the turbine section (91) from the end flowpath inlet (134) to the flowpath exhaust (138). The side flowpath inlet (136) is disposed to a radial outer side of the engine flowpath (36) longitudinally between the end flowpath inlet (134) and the compressor section (89). The flow regulator (146) is configured to regulate a flow of air from the external environment (22) into the engine flowpath (36) through the side flowpath inlet (136).

An assembly is provided for an aircraft. This assembly includes an aircraft propulsion system (20), a propulsion system pylon (108) and a mount (184). The propulsion system pylon (108) is configured to mount the aircraft propulsion system (20) to an airframe (106) of the aircraft. The mount (184) attaches the propulsion system pylon (108) to the aircraft propulsion system (20). The mount (184) includes a bracket (186), a stud (188) and a bushing (190). The bracket (186) includes a receptacle (194) with a frustoconical receptacle surface (198). The stud (188) includes a shoulder (212) and projects axially along an axis (202) into the receptacle (194). The bushing (190) is disposed axially between and is engaged with the frustoconical receptacle surface (198) and the shoulder (212).

An apparatus (100; 200; 300; 404) includes a waveguide (104; 204; 304), and at least one transceiver (102; 204; 302; 406) configured to transmit and receive signals via the waveguide (104; 204; 304). The apparatus (100...404) also includes a plurality of sensor interrogation points (106; 206; 306). Each sensor interrogation point (106; 206; 306) of the plurality of sensor interrogation points (106; 206; 306) is configured to receive, via the waveguide (104; 204; 304), an interrogation signal (210; 201; 310) from the at least one transceiver (102...406), and radiate the interrogation signal (210; 201; 310) toward a target to perform at least one measurement. Each sensor interrogation point (106; 206; 306) of the plurality of sensor interrogation points (106; 206; 306) is also configured to receive a return signal (212, 214, 216; 312, 314, 316) reflected from the target corresponding with the at least one measurement, and redirect the return signal (212...316) into the waveguide (104; 204; 304) for reception by the at least one transceiver (102...406) via the waveguide (104; 204; 304).

A fan blade (20) for a gas turbine engine includes a main body (28) that extends between a leading edge (21) and a trailing edge (22). The main body (28) includes channels (30) formed into the main body (28) and a plurality of ribs (26) extending between the channels (30) and a ledge. The ledge is formed as an integral part of the main body (28) and is configured to define an outer perimeter the channels (30). A cover skin (32) is formed on and metallurgically bonded to the main body (28), including the ledge and the plurality of ribs (26) of the main body (28) using a hybrid additive manufacturing/subtractive manufacturing process. The cover skin (32) is configured to cover (32) the channels (30) and form an aerodynamic surface. The main body (28) and cover skin (32) form an airfoil (18) that extends radially outwardly from a dovetail, which is configured to permit the fan blade (20) to be attached to a fan rotor (16) of the gas turbine engine (210).

A pin-guided spring (240) is used for mounting a ceramic matrix composite (CMC) component (210) of a turbine engine, wherein the CMC component having a hot side configured for exposure to a hot gaspath (205) of the turbine engine and an opposing non-gaspath side having a first flange (212) and a second flange (213) extending radially outward therefrom, the first and second flanges each including axial holes (214, 215) disposed therethrough. The pin-guided spring has a first end (352) with first axial holes (460) in registration with the axial holes in the first flange and dimensioned to engage a plurality of pins (220), a second end (354) with second axial holes in registration with the axial holes in the first flange and dimensioned to engage the plurality of pins, and a compressible central section (356) dimensioned for an interference fit between a support and the second flange. The interference fit on the pin-guided spring can protect the CMC component from damage.

A gas turbine engine (20) includes a propulsor (42), a compressor section (24) and a turbine section (28). The propulsor (42), the compressor section (24) and the turbine section (28) each include rotors for rotation with at least one shaft (104) about an axis (A). At least one bearing compartment (112, 114) includes at least one bearing (108) for supporting the at least one shaft (104). A lubricant supply system (111) supplies lubricant to the at least one bearing (108), and into the at least one bearing compartment (112, 114). The lubricant supply system (111) includes a pump (124) for delivering lubricant into an oil supply tube (110) when the engine (20) is operating for communicating the lubricant into the at least one bearing compartment (112, 114). The lubricant pump (124) further delivers lubricant into a refill reservoir (133) during operation of the engine (20). The refill reservoir (133) communicates with the oil supply tube (110) through a metering plug (131) such that at shutdown of the gas turbine engine (20), lubricant may flow from the refill reservoir (133) into the oil supply tube (110) through the metering plug (131). There is a second of the at least one bearing compartment (112, 114) including at least a second bearing (108), and a second oil supply tube (129) for delivering lubricant to the second bearing compartment (114) when the oil pump (124) is operating. The second bearing compartment (114) further receives lubricant from the refill reservoir (133) through the at least one metering plug (131). There is a single refill reservoir (133) and a single metering plug (131) for providing lubricant to the first and second bearing compartment (112, 114) at engine shutdown. A method is also disclosed.

A rotor stage for a gas turbine engine is provided that includes a disk, first and second rotor blades (152, 252), a platform seal (72), and a damper (70). The first and second rotor blades (152, 252) are attached to the disk and adjacent. The first rotor blade (152) has a first neck portion and a first platform (158), and the second rotor blade (252) has a second neck portion and a second platform (258). The first rotor blade (152) has a first lateral rail (166) that extends laterally outward from a first neck portion surface, and has a first damper post (168) that extends laterally outward from the first neck portion surface. The second rotor blade (252) has a second lateral rail (266) that extends laterally outward from a second neck portion surface, and has a second damper post (268) that extends laterally outward from the second neck portion surface. The damper (70) is circumferentially disposed between the first and second lateral rails (166, 266).

A heat exchanger including at least one heat exchanger wall comprising a first side integral with a second side, the first side comprising a first material with a first coefficient of thermal expansion and the second side comprising a second material with a second coefficient of thermal expansion, wherein the first side is configured to grow equally to the second side responsive to a temperature gradient between the first side and the second side; and a thermally constrained feature formed adjacent the at least one heat exchanger wall.

An aircraft propulsion system (20) includes a core engine (25) that is configured to drive a propulsive fan (22), at least one electric machine (64, 66, 68) that generates thermal energy during operation, a thermal management system (62) that is configured to generate a lubricant flow (80) that accepts the thermal energy generated from the at least one electric machine (64, 66, 68), an oil tank (76, 78) that is configured to supply and store lubricant from the thermal management system (62), the oil tank (76, 78) includes a heat exchange surface (104) in thermal communication with a cooling flow (84) for cooling lubricant stored within the oil tank (76, 78), and a passage (82) that is configured to direct the cooling flow (84) to the heat exchange surface (104) of the oil tank (76, 78).

An assembly is provided for an aircraft. This aircraft assembly includes a turbine engine (32) and a mounting structure (36). The turbine engine (32) extends axially along an axis (24). The mounting structure (36) is configured to mount the turbine engine (32) to an airframe (106) of the aircraft. The mounting structure (36) includes a frame (124) and a linkage (168). The frame (124) is spaced radially outboard from the turbine engine (32). The frame (124) extends circumferentially about the turbine engine (32). The linkage (168) includes a linkage first end (170) and a linkage second end (172). The linkage (168) extends longitudinally between the linkage first end (170) and the linkage second end (172). The linkage (168) is attached to the frame (124) at the linkage first end (170) and the linkage second end (172). The linkage (168) is arranged tangent to and attached to the turbine engine (32) at an intermediate section (174) of the linkage (168) longitudinally between the linkage first end (170) and the linkage second end (172).

An aircraft propulsion system includes an open propulsor rotor, an open guide vane structure, a turbine engine and a heat exchanger. The open guide vane structure includes a plurality of open guide vanes. The turbine engine is configured to drive rotation of the open propulsor rotor. The turbine engine includes a flowpath, a compressor section, a combustor section and a turbine section. The flowpath extends through the compressor section, the combustor section and the turbine section from a flowpath inlet to a flowpath exhaust. The flowpath inlet is disposed axially downstream of the open propulsor rotor and is fluidly coupled to an external environment. The heat exchanger is disposed radially outboard of the flowpath. At least a portion of the heat exchanger is disposed axially between the flowpath inlet and the open guide vanes. The heat exchanger is configured to transfer heat energy between air and a liquid working fluid.

An aircraft has a gas turbine engine. A first fuel flowpath (410) passes from a fuel source (202) to a combustor (70). An air-air heat exchanger has a heat donor leg along a bleed flowpath and a heat recipient leg along a bypass flowpath. An air-fuel heat exchanger (86) has a heat donor leg (120) along the bleed flowpath downstream of the air-air heat exchanger. A second fuel flowpath (420) passes from the fuel source (202) through the air-fuel heat exchanger (86) and includes a heat recipient leg (122) of the air-fuel heat exchanger (86) and at least partially non-overlapping the first fuel flowpath (410). One or more third fuel flowpaths (430) pass from the fuel source (202) through fuel-driven actuators (230). The second fuel flow path (420) and the one or more third fuel flowpaths (430) return to the fuel source (202) or the first fuel flowpath (410).

Environmental barrier coatings (EBCs) comprising a bond coat (520) and a top coat (530) are disclosed. An EBC bond coat (520), an EBC top coat (530), or both coats may comprise (i) one or more transition metal silicate having an initial particle size that is coarse ("coarse TM silicate") (580) and (ii) one or more transition metal silicate having an initial particle size that is fine ("fine TM silicate") (590). Fine TM silicate (590) may partially or completely plate one or more coarse TM silicate particles (580). Methods of producing such EBCs are disclosed.

A gas turbine engine (20) includes a propulsor (22), a compressor section (24), and a turbine section (28). The propulsor (22), the compressor section (24) and the turbine section (28) include rotors for rotation with at least one shaft (104). At least one bearing compartment (112, 114) includes at least one bearing (108) for supporting the at least one shaft (104). A lubricant supply system (111) supplies lubricant to the at least one bearing (108), and into the at least one bearing compartment (112, 114). The lubricant supply system (111) includes a pump (124) for delivering lubricant into an oil supply tube (110, 129) when the engine (20) is operating for communicating the lubricant into the at least one bearing compartment (112, 114). A check valve (130) is on a fluid line (97) intermediate the pump (124) and the oil supply tube (110, 129). The check valve (130) allows flow of lubricant from the pump (124) into the at least one bearing compartment (112, 114) when the engine (20) is operating, and resists migration of lubricant from the oil supply tube (110, 129) when the engine (20) is shut down. A method is also disclosed.

A method of repairing a part (100') with a directed energy deposition (DED) repair operation includes identifying a repair zone (108') on the part (100'). The repair zone (108') represents a region that includes at least one zone of operational wear or damage requiring the DED repair operation. A layer of repair material is applied to the repair zone (108') and the repair zone (108') is preheated using a first energy source (110b'). The layer of repair material in the repair zone (108') is melted using a second energy source (110') and is permitted to cool and solidify. The applying, preheating, melting, and cooling steps are repeated a desired number of times to accomplish the desired repair. Use of the first energy source (110b') and the second energy source (110') results in a temperature-affected zone (116') that does not extend beyond the repair zone (108') such that heat-sensitive portions of the part (100') do not experience distortion that exceeds allowable tolerances.

An assembly for an aircraft powerplant includes an electric device (108A, 108B) and a closed loop fluid circuit (142A, 142B). The electric device includes a device housing (120A, 120B) and electrical circuitry (122A, 122B) housed within an interior of the device housing. The closed loop fluid circuit is configured to cool the electrical circuitry and includes a circuit path (146A, 146B) and a plurality of circuit components (148, 150, 152 and 154). The circuit path extends through the circuit components. Each of the circuit components are at least partially housed within the interior of the device housing. The circuit components may include a heat exchanger (150A, 150B).

An assembly for an aircraft powerplant (20) includes a first powerplant component, a differential geartrain (156), a first clutch (164), a second clutch (166) and a first turbine engine (26). The first powerplant component includes a first component rotor. The differential geartrain (156) includes a first geartrain component and a second geartrain component. The first clutch (164) is configured to operatively couple the first geartrain component to the first component rotor during a first mode. The first clutch (164) is configured to operatively decouple the first geartrain component from the first component rotor during a second mode. The second clutch (166) is configured to operatively decouple the second geartrain component from the first component rotor during the first mode. The second clutch (166) is configured to operatively couple the second geartrain component to the first component rotor during the second mode. The first turbine engine (26) includes a first engine rotating structure (116) operatively coupled to the second geartrain component.

An assembly for an aircraft powerplant (20) includes a vane structure (148), a first electric device (108A, 108B) and a cooling system (36). The first electric device is disposed within the vane structure (148) and includes a first device housing (12A, 120B) and first electrical circuitry (122). The first electrical circuitry is housed within an interior of the first device housing. The cooling system is configured to cool the first electrical circuitry and includes a first heat exchanger (146) and an air circuit (155). The first heat exchanger is disposed within the vane structure and thermally coupled to the first electric device. The air circuit extends within the vane structure from an airflow inlet (158) into the air circuit, through the first heat exchanger, to an airflow outlet (160) m the air circuit. The airflow inlet fluidly couples the air circuit to a volume (162) external to the vane structure.

A turbine engine assembly (24) includes an airfoil (96) and a baffle (98). The baffle is disposed in an airfoil cavity (119,120) with an outer passage (148,149,150) formed between the baffle and a wall (110) of the airfoil. The baffle extends longitudinally from a baffle end (140) to a baffle trailing edge (142) and laterally between a baffle first side (144) and a baffle second side (146) with the baffle second side meeting the baffle first side at the baffle trailing edge. The baffle includes an endwall (158), a first sidewall (160), a second sidewall (162) and a baffle cavity (164). The endwall forms the baffle end laterally along the baffle cavity. The first sidewall and the second side wall form respectively the baffle first side and second side longitudinally along the baffle cavity and are bonded at the baffle trailing edge.