Pratt & Whitney Canada Patente

A combustor (26) for a gas turbine engine is provided that includes an annular combustor shell (34), a plurality of fuel nozzles (36), a rotary fuel slinger (38), and a plurality of first igniters (54). The annular combustor shell (34) includes front and rear annular liners (34A, 34B) that collectively form the combustor shell (34) and define a combustion chamber (40) that includes inner and outer radial combustion zones (40A, 40B). The fuel nozzles (36) are engaged with the front annular liner (34A) and the outer radial combustion zone (40B), and the fuel nozzles (36) are configured to inject a first fuel. The rotary fuel slinger (38) is engaged with the combustor shell (34) and configured to deliver a second fuel into the inner radial combustion zone (40A). The first fuel is a different type of fuel than the second fuel. The first igniters (54) are engaged with the outer radial combustion zone (40B) and are configured to ignite the first fuel.

An impending bypass indicator (IBI) (22; 22'; 22") includes a housing (30) having a fluid inlet (24) for admitting a fluid (11) into an interior (28) of the housing (30), and a piston (34) displaceable in the interior (28) of the housing (30) in response to a pressure variation of the fluid (11) in the interior (28), the piston (34) displaceable between a first position and a second position spaced apart from the first position. A piston magnet (50) is coupled to the piston (34). An indicator (22; 60) indicates an impending bypass of a component (20) in response to the piston (34) being in the second position. A housing magnet (54) engages the piston magnet (50) in the second position and generates a magnetic force acting on the piston magnet (50). A biasing element (56) biases the piston (34) towards the first position with a biasing force inferior to the magnetic force. An electromagnetic coil (58) adjacent the housing magnet (54) is selectively electrifiable to generate a second magnetic force opposite to the magnetic force.

An apparatus may include a first terminal, a second terminal, a direct current (DC) motor (202a, 302a, 402a, 502a, 602a), a brake mechanism (204a, 304a, 404a, 504a, 604a), and a voltage blocking mechanism. The DC motor (202a, 302a, 402a, 502a, 602a) may be configured to operate when a voltage is applied across the first terminal and the second terminal. The brake mechanism (204a, 304a, 404a, 504a, 604a) may be configured to apply braking to the DC motor (202a, 302a, 402a, 502a, 602a) when no voltage is present across the first terminal and the second terminal, and remove the braking from the DC motor (202a, 302a, 402a, 502a, 602a) when the voltage is applied across the first terminal and the second terminal. The voltage blocking mechanism may be configured to prevent the DC motor (202a, 302a, 402a, 502a, 602a) from operating when the voltage is applied across the first terminal and the second terminal until the braking from the brake mechanism is removed from the DC motor (202a, 302a, 402a, 502a, 602a).

A propulsion system (20) for an aircraft (1000) includes a battery (64), an electrical distribution system (66), a first control channel (110, 114), and a second control channel (112, 116). The battery (64) includes a plurality of battery strings (72). The electrical distribution system (66) includes at least one string contactor (132, 142) for each of the plurality of battery strings (72). The at least one string contactor (132, 142) includes a first gate (138, 148) connected to the first control channel (110, 114) and a second gate (140, 150) connected to the second control channel (112, 116). The at least one string contactor (132, 142) is operable to a closed position with a first control signal present at the first gate (138, 148) and a second control signal present at the second gate (140, 150). The at least one string contactor (132, 142) is operable to an open position with one or both of the first control signal absent at the first gate (138, 148) or the second control signal absent at the second gate (140, 150).

A method of monitoring a battery system (42) utilized within an aircraft propulsion system is provided that includes: providing a parallel electrical circuit having a first and second battery units (38A, 38B) and first and second electrical contactors (60A, 60B) disposed in a parallel electrical circuit, positive and negative bus connectors (60A, 60B) electrically connected to first and second legs of the circuit, and an electrical device (34) electrically connected to the positive and negative bus connectors (56, 58). The method further includes controlling the first and second electrical contactors (60A, 60B) to be in a closed configuration; using a first voltage sensor (62A) to determine a first voltage across the first battery unit (38A); using a second voltage sensor (62B) to determine a second voltage across the second battery unit (38B); using a bus voltage sensor (66) to determine a bus voltage between the positive and negative bus connectors (56, 58); and monitoring the battery system (42) using the first, second, and bus voltages.

Systems (38; 138) and methods (1000; 2000; 3000) for inspecting components (12) with thermography are provided. The systems (38; 138) and methods (1000; 2000; 3000) facilitate the selection of thermography parameters (45) for a particular inspection situation. A method (1000; 2000; 3000) includes executing a machine learning algorithm (70A, 70B) to determine one or more thermography parameters (45) based on one or more characteristics (37) of the component (12), and performing a thermographic inspection according to the one or more thermography parameters (45). The thermographic inspection includes exciting the component (12) to induce a thermal response in a region of the component (12) containing a defect (22), and acquiring a thermographic image (46) of the region of the component containing the defect (22).

Systems (38; 138) and methods for non-destructive evaluation of components (12) using computed tomography and infrared thermography are provided. A method includes acquiring, using X-ray computed tomography, a digital three-dimensional representation of the component (12), and identifying a region of the component (12) containing the defect (22) from the digital three-dimensional representation of the component (12). The component (12) is mechanically excited to induce a thermal response in the region of the component (12) containing the defect (22). A thermographic image (46) of the region of the component (12) containing the defect (22) is acquired with an infrared sensor (44) while the thermal response is exhibited. The component (12) is evaluated using the thermographic image (46).

An anti-windmilling control method for a gas turbine engine (20...600) including an engine shaft (404;610) with a rotating component and an accessory gearbox (234,236) each operably coupled to the engine shaft (404;610), an accessory locking brake (312;314;604) operably connected to the accessory gearbox and configured to selectively control operation of at least one gear of the accessory gearbox, and a controller (212;602) arranged in communication with the accessory locking brake, the method including: determining, with the controller, that the gas turbine engine is at altitude; determining, with the controller, that the gas turbine engine is in a state of windmilling; and in response to determining that the gas turbine engine is in the state of windmilling, activating the accessory locking brake to engage with the at least one gear to prevent rotation of the at least one gear, thereby halting rotation of the engine shaft and the rotating component.

An aircraft engine includes an air inlet duct and at least one strut having a leading edge and a trailing edge (31B). The at least one strut (30) extends across at least part of the air inlet duct and has a strut passage. The trailing edge (31B) has one or more edge contours (39), each defining a contour edge wall (39W) recessed from a baseline surface (31BS) of the trailing edge (31B). The one or more edge contours (39) have a recessed tap (38) in fluid communication with the strut passage.

A hybrid-electric propulsion system (26) for an aircraft (20) is provided that includes a propulsion unit (32), a propulsion unit gear box (30), a gas turbine engine (28), and an electric motor drive system (34). The gas turbine engine (28) has compressor, combustion, and turbine sections, and is configured to provide motive force to the propulsion unit gear box (30). The electric motor drive system (34) is disposed to provide motive force to the propulsion gear box (30) or to the gas turbine engine (28). The electric motor drive system (EMDS) (34) is adapted to interchangeably use a plurality of EMDS packages, each EMDS package including an electric motor (36) and a battery (38). The electric motor drive system (34) has a set of performance parameters that vary depending upon the particular EMDS package that is used with the electric motor drive system (34). The electric motor drive system (34) includes interfaces that allow each EMDS package to be interchangeably used.

A propulsion system for an aircraft (1000) includes a battery (64), a battery management system (68), and an engine controller (28). The battery management system (68) includes a plurality of first battery temperature sensors (110), at least one second battery temperature sensor (110), and a single-channel battery management system (BMS) controller (92). The single-channel BMS controller (92) includes a BMS control channel connected in signal communication with the plurality of first temperature sensors (110). The engine controller (28) includes a first engine control channel (98) and a second engine control channel (100). The first engine control channel (98) is connected in signal communication with the BMS control channel to form a first control lane (142A). The second engine control channel (100) is connected in signal communication with the at least one second battery temperature sensor (110) to form a second control lane (142B).

A propulsion system for an aircraft (20) is provided that includes a thermal engine (30), an electrical power motive system (28), a propulsion unit (40), and a thermal management system (38). The electrical power motive system (28) has an electric motor (46). The thermal management system (38) includes a first heat exchanger (56), and is configured to cycle a flow of lubricant (64) into and out of the thermal engine (30) and into and out of the first heat exchanger (56) in a first loop (60), and is configured to cycle a flow of coolant (66) into and out of thermal engagement with the electrical power motive system (28) and into and out of the first heat exchanger (56) in a second loop (62). The first heat exchanger (56) is configured to permit heat transfer between the flow of lubricant (64) and the flow of coolant (66) to occur within the first heat exchanger (56).

An aircraft engine (10) having a centerline (11) has: an oil system (28; 100) including: a component (101) located above the centerline (11) and having an inlet (101A) and an outlet (101B); an oil circuit (110) interconnecting the components (101) and including an input line (111) connected to the inlet (101A) and an output line (112) connected to the outlet (101B), the input line (111) defining an input U-shaped section (111A), the output line (112) defining an output U-shaped section (112A); a valve (113, 114) connected on the input line (111) or the output line (112) and located between the component (101) and a respective one of the input U-shaped section (111A) and the output U-shaped section (112A); and an anti-siphon line (120) connecting the input line (111) to the output line (112) and defining a flow path of a lesser flow circulating area of the input line (111) and the output line (112), the input U-shaped section (111A), the output U-shaped section (112A), and the component (101) located between connection points (121, 122) defined between the anti-siphon line (120) and the input and output lines (111, 112).

A method of operating a dual fuel gas turbine engine (102) includes operating the engine (102) with a first fuel according to a first control loop (202). The method also includes while the engine (102) is operating with the first fuel, bumplessly transitioning the engine (102) to operate with a second fuel, and operating the engine (102) with the second fuel according to a second control loop (204).

A rotating structure (70) for an aircraft powerplant (20) includes a shaft (74), a retainer (78), a balancing device (76) and a bladed rotor. The shaft (74) includes a bore (82) and a shoulder (102) axially along the bore (82). The retainer (78) is disposed in the bore (82) and mounted to the shaft (74). The balancing device (76) is configured to rotationally balance the rotating structure (70) about the axis (72). The balancing device (76) is disposed in the bore (82) and axially retained between the shoulder (102) and the retainer (78). The balancing device (76) includes a balancing mass (114) with a plurality of mass teeth (120) arranged circumferentially about the axis (72). During a first condition, the balancing mass (114) is rotationally repositionable about the axis (72) relative to the shaft (74) while the balancing device (76) remains in the bore (82) and axially retained between the shoulder (102) and the retainer (78). During a second condition, the mass teeth (120) are meshed with a plurality of shaft teeth (96) to rotationally fixed the balancing mass (114) to the shaft (74).

A rotating structure (70) for an aircraft powerplant (20) includes a balancing device (76) and a bladed rotor. The balancing device (76) is configured to rotationally balance the rotating structure (70) about a rotational axis (72). The balancing device (76) includes a housing (82) and a plurality of balancing masses (84). The housing (82) includes a first sidewall (94), a second sidewall (96), an outer endwall (98), an annular channel (102) and an annular slot (114). The outer endwall (98) circumscribes the annular channel (102). The annular channel (102) extends axially within the housing (82) between the first sidewall (94) and the second sidewall (96). The annular slot (114) extends radially through the outer endwall (98) to the annular channel (102). Each of the balancing masses (84) includes a mass body (130) and an engagement feature (134). The mass body (130) is captured within the annular channel (102). The engagement feature (134) is configured to be engaged by a tool (162) from outside of the housing (82) to adjust a circumferential position of the mass body (130) about the rotational axis (72).

An aircraft system is provided that includes a first propulsor rotor (42A), a first transmission (106A), a second propulsor rotor (42B), a second transmission (106B) and an intermittent combustion engine (68). The first propulsor rotor (42A) is rotatable about a first propulsor axis (48A). The first transmission (106A) is coupled to the first propulsor rotor (42A). The second propulsor rotor (44) is rotatable about a second propulsor axis (48B). The second transmission (106B) is coupled to the second propulsor rotor (44). The intermittent combustion engine (68) is configured to drive rotation of the first propulsor rotor (42) through the first transmission (106A). The intermittent combustion engine (68) is configured to drive rotation of the second propulsor rotor (44) through the second transmission (106B).

An oil supply system (130) for an aircraft engine (10) includes: an oil tank (32); a closed-loop oil circuit (52) including a component (22, 36), a pump (48) having a pump inlet (48A) and outlet (48B), and a de-aerator (54); a make-up flow conduit (60, 60', 60") connecting the oil tank (32) to the closed-loop oil circuit (52); a make-up pump (62, 62', 62") connected to the closed-loop oil circuit (52); a recirculation conduit (63) in parallel to the make-up pump (62, 62', 62") and bypassing the make-up pump (62, 62', 62"); a fixed orifice (65) fluidly connected to the make-up flow conduit (60, 60', 60") upstream of the closed-loop oil circuit (52) and downstream of the recirculation conduit (63), the fixed orifice (65) having an area selected to increase a pressure at an outlet (62B) of the make-up pump (62, 62', 62") above a pressure at the de-aerator inlet (54A); and a metering orifice (64) connected on the recirculation conduit (63) and having a flow circulating area selected to reduce a flow rate inputted into the closed-loop oil circuit (52) below a flow rate generated by the make-up pump (62, 62', 62").

An aircraft powerplant (20) assembly includes a compressor section (36), a combustor section (37), a turbine section (38), a flowpath and a rotating structure (70). The flowpath extends through the compressor section (36), the combustor section (37) and the turbine section (38). The rotating structure (70) includes a balancing device (76) and a bladed rotor disposed in the turbine section (38) or the compressor section (36). The balancing device (76) is configured to rotationally balance the rotating structure (70) about a rotational axis (72) of the rotating structure (70). The balancing device (76) includes a mounting platform (108) and a plurality of balancing masses (88). The mounting platform (108) includes a plurality of mounting apertures (120). The balancing masses (88) are arranged circumferentially about the axis (72). Each of the balancing masses (88) includes a shank (130) and a head (128). The shank (130) of each of the balancing masses (88) is threaded into a respective one of the mounting apertures (120). The head (128) of each of the balancing masses (88) is disposed radially inboard of the mounting platform (108).

A coating system (16) for an aircraft component (18) is described herein. The component (18) includes a substrate (28) having a surface (30) for the application of the coating system (16). The coating system (16) includes a diffusion barrier coating layer (20) disposed on the substrate surface (30). A bond coating layer (22) may be disposed on the diffusion barrier coating layer (20). The diffusion barrier coating layer (20) may be formed from a first MCrAlX alloy (24), where M includes, for example, cobalt and X includes, for example, yttrium. The bond coating layer (22) may be formed from a second MCrAlX alloy (26), where M includes, for example, nickel and X includes, for example, one or more of yttrium, silicon, or hafnium. The diffusion barrier coating layer (20) and/or the bond coating layer (22) may have a microstructure that includes gamma and beta phases.

An aircraft system (20) includes a compressor section, a combustor section, a turbine section, an exhaust duct (92) and a flowpath (46). The exhaust duct (92) includes an upstream duct section (96), a plurality of intermediate duct sections (97A,97B) and a downstream duct section (98). A first (97A) of the intermediate duct sections is fluidly discrete from a second (97B) of the intermediate duct sections. The intermediate duct sections (97A,97B) are fluidly coupled in parallel between the upstream duct section (96) and the downstream duct section (98). The flowpath (46) extends sequentially longitudinally through the compressor section, the combustor section, the turbine section, the upstream duct section (96), the intermediate duct sections (97A,97B) and the downstream duct section (98) from an inlet into the flowpath to an outlet (64) from the flowpath.

A method is provided which includes: processing first fuel data to determine a first trend of a powerplant performance characteristic for a first fuel, wherein the first fuel data is indicative of a plurality of first values of a tracked operational parameter for an aircraft powerplant model over a period of time, and the first values were recorded when combusting the first fuel; processing second fuel data to determine a second trend of the powerplant performance characteristic for a second fuel, wherein the second fuel data is indicative of a plurality of second values of the tracked operational parameter for the aircraft powerplant model over the period of time, and the second values were recorded when combusting the second fuel; and scheduling use of the first fuel and the second fuel based on the first trend of the powerplant performance characteristic and the second trend of the powerplant performance characteristic.

A pusher aircraft power plant (10) for propelling an aircraft is configured to improve serviceability of swirl recovery vanes (22). The pusher aircraft power plant (10) includes a prime mover (12) operable to generate motive power, an enclosure (14) housing at least part of the prime mover (12), a bladed rotor (16) drivingly coupled to the prime mover (12) and disposed aft of the prime mover (12), and a plurality of vanes (22) angularly distributed about a rotation axis (RA) of the bladed rotor (16) and being non-rotatable about the rotation axis (RA) of the bladed rotor (16). The vanes (22) are disposed aft of the bladed rotor (16) to interact with a flow of air propelled by the bladed rotor (16) during rotation of the bladed rotor (16). The vanes (22) are disposed aft of the enclosure (14) permitting the vanes (22) to be uninstalled from the pusher aircraft power plant (10) without accessing the enclosure (14).

A system (10) and method (100) for manufacturing a component (14) with optimum geometry includes determining a target mass flow rate (102) through an aperture (48) of the component (14) and determining geometric variances (104) associated with the aperture (48). Optimum geometry is selected based on the geometric variances and the target mass flow rate. Input variables for a manufacturing tool are determined that correspond to the optimum geometry and the component (14) is manufactured based on the selected input variables.

A system for feeding pre-cooled air to an aircraft environmental control system (ECS) (28) comprises an aircraft engine (16) including a compressor section (20) having a bleed port, and an engine oil circuit including an oil cooler (25) for cooling engine oil. The system further includes an air pre-cooler (34) including an air passage (38) in heat exchange relationship with a coolant passage (36). The air passage (38) has an air inlet (38a) fluidly connected to the bleed port of the compressor section (20) and an air outlet (38b) fluidly connected to an inlet (32) of the ECS (28). The coolant passage (36) has a coolant inlet (36a) fluidly connected to an oil outlet of the oil cooler (25) and a coolant outlet (36b) fluidly connected to the engine oil circuit at a location upstream of the oil cooler (25).

A tractor aircraft power plant (10) for propelling an aircraft is configured to improve serviceability of swirl-recovery vanes (22). The tractor aircraft power plant (10) includes an intermittent combustion engine (12) such as a Wankel engine operable to generate motive power. A bladed rotor (16) is drivingly coupled to the Wankel engine (12) and is disposed forward of the Wankel engine (12) relative to a propulsion direction. A gearbox (38) drivingly couples the bladed rotor (16) to the Wankel engine (12). A plurality of vanes (22) are angularly distributed about a rotation axis (PA) of the bladed rotor (16) and are non-rotatable about the rotation axis (PA) of the bladed rotor (16). The vanes (22) are disposed aft of the bladed rotor (16) to interact with a flow of air propelled by the bladed rotor (16) during rotation of the bladed rotor (16). The vanes (22) are mounted to the gearbox (38) to facilitate serviceability of the vanes (22).

A method (200; 300) includes sending a first control signal (205; 305), wherein the first control signal causes a turbine engine of a plurality of turbine engines of an aircraft to initiate an entry into an asymmetric operating regime (210). The method further includes receiving a control input for an exit from the asymmetric operating regime (215; 310), determining a progression state of the turbine engine into the asymmetric operating regime (220a, 225a, 225b; 315), and sending a second control signal for accelerating the turbine engine to a high-power operating regime (227, 229; 320), wherein the second control signal specifies an acceleration rate based on the progression state of the turbine engine into the asymmetric operating regime.

A lubrication system (78) includes a first rotating shaft (52) and a second rotating shaft (50). The first rotating shaft (52) and the second rotating shaft (50) are separated by a flowpath (58). The flowpath (58) includes a fluid (72). The fluid (72) is directed towards an outer surface (80) of the first rotating shaft (52). The fluid (72) thereafter moves towards the outer surface (84) of the second rotating shaft (50). The second structure includes a recess (86) configured to collect the fluid (72). The outer surface (80) may include a protrusion (90). The protrusion (90) is configured to direct fluid (72) from the outer surface (80) to the recess (86).

An assembly for a powerplant includes a housing (34), a primary fuel injector (38) and an ignition system (40). The housing (34) forms a combustion volume (66) within the housing (34). The primary fuel injector (38) is configured to inject primary fuel into the combustion volume (66). The ignition system (40) is configured to ignite the primary fuel within the combustion volume (66). The ignition system (40) includes a pilot fuel injector (70), a pilot ignitor (72), a pilot chamber (74), a first component (78) and a second component (80). The pilot fuel injector (70) is configured to inject pilot fuel into the pilot chamber (74). The pilot ignitor (72) is configured to ignite the pilot fuel within the pilot chamber (74). The pilot chamber (74) is fluidly coupled with the combustion volume (66) through an aperture in the first component (78). The pilot chamber (74) is formed by and disposed between the first component (78) and the second component (80). The first component (78) is configured from or otherwise include a ceramic.

A method of operation is provided. The operating method includes: operating an aircraft powerplant (20) with a powerplant performance characteristic at a first setting, the operating of the aircraft powerplant (20) including combusting a first fuel in the aircraft powerplant (20); receiving a command to operate the aircraft powerplant (20) with the powerplant performance characteristic at a second setting; modeling operation of the aircraft powerplant (20) using a first fuel digital twin of the aircraft powerplant (20), the operation of the aircraft powerplant (20) modeled to determine a first fuel value of a first operational parameter for the aircraft powerplant (20) when operating the aircraft powerplant (20) with the powerplant performance characteristic at the second setting and combusting the first fuel in the aircraft powerplant (20); and selecting the first fuel or a second fuel for operating the aircraft powerplant (20) with the powerplant performance characteristic at the second setting based on the first fuel value of the first operational parameter.

A bypass duct (22) of a turbofan gas turbine engine (10) includes a first shroud (26) extending at least partially around an axis (A) and a second shroud (24) extending at least partially around the axis (A), the second shroud (24) being radially spaced apart from the first shroud (26) to define a bypass passage between the first (26) and second (24) shrouds. The bypass duct (22) further includes a fairing (12,12',12") disposed in the bypass passage. The fairing (12,12',12") includes a first outer shell (32A,32A',32A") and a second outer shell (32B,32B',32B") defining therebetween an internal passage (30,30',30"). The first (32A,32A',32A") and second (32B,32B',32B") outer shells extend radially from the first shroud (26) to the second shroud (24). A plate (40,40',40") extends across the internal passage (30,30',30") between the first (32A,32A',32A") and second (32B,32B',32B") outer shells. A damper (42,42',42") is engaged with the plate (40,40',40") between the first (32A,32A',32A") and second (32B,32B',32B") outer shells.

A spur gear (42) of a gearbox (30) of a gas turbine engine system of an aircraft includes a gear body (50), and a plurality of gear teeth (52) extending radially outwardly from the gear body (50). A gear tooth of the plurality of gear teeth (52) includes two opposing tooth sides (66,68) extending to a tooth tip (62). The tooth tip (62) defines a radially outboard extent of the gear tooth (52). A first tooth side (66) of the opposing tooth sides (66,68) has a first pressure angle (72), and a second tooth side (68) of the opposing tooth sides (66,68) has a second pressure angle (76) different from the first pressure angle (72).

A rotor housing (46) for an aircraft rotary engine (12) includes a side housing body (82) and a rail (84). The side housing body (82) extends along an axis between and to an inner side (94) and an outer side (96). The side housing body (82) forms a fluid cooling passage (106) and a plurality of ribs (114). The fluid cooling passage (106) extends about the axis at the inner side (94). The plurality of ribs (114) are coincident with and extend into the fluid cooling passage (106). The plurality of ribs (114) are distributed about the fluid cooling passage (106) as an array of ribs (114). The rail (84) is disposed at the plurality of ribs (114). The rail (84) extends about the fluid cooling passage (106). The side housing body (82) and the rail (84) form a plurality of fluid cooling channels (118) connected in fluid communication with the fluid cooling passage (106).

A method for assembling a rotational equipment assembly having a rotational equipment module (74) and a rotational equipment case (76) includes axially supporting the rotational equipment module (74) with a first support assembly (156), positionally fixing a rotor assembly (86) and a stator assembly (88) of the rotational equipment module (74) together with a second support assembly (158), coupling a pulling rod assembly (160) with a tooling shaft (84) of the rotational equipment module (74), and assembling the rotational equipment assembly by positioning the rotational equipment case (76) axially abutting the rotational equipment module (74) and axially pulling the pulling rod assembly (160) to axially fit the rotational equipment module (74) with the rotational equipment case (76).

A gas turbine engine (20) includes a compressor section (28), a combustion section (30), and a turbine section (32). The turbine section has a rotor stage (40) disposed adjacent an annular stator vane stage (42). The annular stator vane stage includes annular inner (50) and outer radial shrouds (52), stator vanes (48), and a retaining ring (72). The stator vanes are circumferentially spaced apart from one another, extending between the shrouds. The inner radial shroud is mechanically engaged with a first engine support structure (54) at a first position axially forward of the plurality of stator vanes, and the outer radial shroud is mechanically engaged with a second engine support (64) structure at a second position axially forward of the plurality of stator vanes. The annular stator vane stage includes a flange (66) extending outward from the outer radial shroud that is contiguous with the retaining ring.

A compressor assembly (60) has: an impeller (61) rotatable about a central axis (A), the impeller (61) mounted on a shaft (14B); a bearing housing (30) extending around the central axis (A) and radially supporting the shaft (14B); and a diffuser (62) downstream of the impeller (61), the diffuser (62) having: an annular member (63) extending circumferentially around the central axis (A), vanes (64) circumferentially distributed around the central axis (A) and protruding from one of the annular member (63) and the bearing housing (30), the vanes (64) extending from bases (64A) at the one of the annular member (63) and the bearing housing (30) to tips (64B) at the other of the annular member (63) and the bearing housing (30), and lug and slot connections defined between the annular member (63) and the bearing housing (30) to secure the annular member (63) to the bearing housing (30).

A mounting arrangement includes a casing (17) of an aircraft engine (10) extending around an axis, an accessory (22), and a bracket (40) for mounting the accessory to the casing. The bracket (40) includes a first bracket segment (44) coupled to the casing at a first mounting point and a second bracket segment (48) coupled to the accessory at a second mounting point. The mounting arrangement further includes one or more shoulder bolts (56) fastening one or more of the first bracket segment (44) to the casing at the first mounting point and the second bracket segment (48) to the accessory at the second mounting point, and one or more wave springs (54) disposed about a shaft of the one or more shoulder bolts (56) and engaging one or more of the first bracket segment (44) at the first mounting point and the second bracket segment (48) at the second mounting point.

A combustor (100; 200) includes a liner (102) defining a combustion chamber (105) and receiving a fuel and air mixing body (104; 204; 300). The mixing body (104...300) has a central fuel supply (118, 120; 223,, 225; 302), and radial distribution passages (128, 138, 142; 224, 230, 239; 304, 314, 318, 322, 330, 338) communicating fuel from the central fuel supply (118..302) radially outwardly relative to a central axis (C) of the central fuel supply (118...302) and to mixing passages (110, 113, 136, 140; 210, 212, 220, 236; 310; 320, 328). The radial distribution passages (128...338) have injection ports (130) in the mixing passages (110...328). The mixing passages (110...328) extend from a rear face (106; 206) of the mixing body (104 ...300) to an inner face (116; 214) facing into the combustion chamber (105). Air inlets (108) in the mixing body (104...300) communicate air into the mixing passages (110...328), and there is cellular material (114; 213, 221) in the mixing passages (110...328) at a location at which the fuel is injected into the mixing passages (110...328). A gas turbine engine (20) is also disclosed.

A method is provided during which lubrication material is arranged with a fastening member. The lubrication material is at a first temperature during the arranging of the lubrication material such that the lubrication material is solid and forms a self-supporting lubrication ring. The arranging of the lubrication material includes one of: mounting the self-supporting lubrication ring onto a shank of the fastening member; or mounting the self-supporting lubrication ring onto a surface defining an aperture such that the lubrication ring extends around the aperture.

An aircraft propulsion system (20) includes a turbocompressor unit (40), an engine exhaust assembly (32), and a tube assembly (118). The turbocompressor unit (40) includes a compressor (44), a turbine (42), a turbocompressor rotational assembly (46), and a static structure (48). The static structure (48) forms a first cavity (72) and a second cavity (84). The engine exhaust assembly (32) includes an annular scroll (38) disposed between the first cavity (72) and the second cavity (84). The annular scroll (38) includes a scroll body (86) forming a flow channel (90), and a plurality of hollow stator vanes (94). The plurality of hollow stator vanes (94) is disposed within the flow channel (90). The tube assembly (118) includes a tubular body (120). The tubular body (120) extends through one of the plurality of hollow stator vanes (94) between and to the first cavity (72) and the second cavity (84). The tubular body (120) forms an internal passage (112) extending from the first axial tube end (130) to the second axial tube end (132). The internal passage (112) connects the first cavity (72) in fluid communication with the second cavity (84).