Pratt & Whitney Canada Patente

A system comprising a sensor (204) connected to an aircraft system (102) configured to monitor an aircraft system characteristic and generate sensor data responsive thereto. A diagnostic data module (DDM) (202) is configured to receive the sensor data from the sensor (104) and transmit the received sensor data from the DDM (202) to a remote monitoring unit. The DDM (202) operates independently of a preexisting data monitoring system (106) of the aircraft system (102).

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 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.

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).

An impeller rotor (30) includes a plurality of fluid circuits (106) arranged circumferentially about an axis (44) within the impeller rotor (30). Each of the fluid circuits (106) includes a first outlet passage (110), a second outlet passage (112) and an inlet passage (108) fluidly coupled to the first outlet passage (110) and the second outlet passage (112) in parallel. A seal runner (118) includes a bore (132), a sidewall (126) and a rim (128). The bore (132) is fluidly coupled to the first outlet passage (110) of each of the fluid circuits (106). The rim (128) projects radially out from the sidewall (126) to an outer end (164) of the rim (128). The rim (128) is seated in a recess (204) in the impeller rotor (30). The rim (128) includes a flat disposed at the outer end (164) of the rim (128). The flat of the rim (128) is configured to engage a flat of the impeller rotor (30) to rotationally fix the seal runner (118) to the impeller rotor (30).

A shaft assembly (42) for a gas turbine engine (20) including an outer shaft (46), an inner shaft (44), a lock nut (48), a locking shaft (50), and a washer (52). The outer shaft (46) includes a bore (56) extending between first and second ends. A portion (74, 78, 82) of the inner shaft (44) extends within the bore (56) of the outer shaft (46). The lock nut (48) is secured to the inner shaft (44) proximate a second axial end (74) of the inner shaft (44). The locking shaft (50) includes a body portion (100) extending between a coupling end (102) and a retaining end (104). The locking shaft (50) is disposed within the bore (56), with the coupling end (102) engageable with an end of the lock nut (48). A washer (52) is engageable with the retaining end (104) and the outer shaft second end. In the assembled state, the locking shaft (50) prevents rotation of the lock nut (48).

An apparatus (20) for an aircraft powerplant includes an impeller rotor (30) and a seal runner (118). The impeller rotor (30) includes a plurality of fluid circuits (106). Each of the fluid circuits (106) includes a first outlet passage (110), a second outlet passage (112) and an inlet passage (108) fluidly coupled to the first outlet passage (110) and the second outlet passage (112) in parallel. The seal runner (118) includes a bore (132), an inner guide surface (144) and an outer land surface (156). The seal runner (118) extends circumferentially about the axis (44) and radially between the inner guide surface (144) and the outer land surface (156). The bore (132) is fluidly coupled to the first outlet passage (110) of each fluid circuit (106). The inner guide surface (144) forms a radial outer peripheral boundary of the bore (132). The inner guide surface (144) radially diverges away from the axis (44) as the inner guide surface (144) extends axially away from the impeller rotor (30) and to a distal end (136) of the seal runner (118).

An aircraft engine (10), has: a turbine (15) including a turbine rotor (15C) rotatable about a central axis (A); a scroll case (20) having an inlet (22) fluidly connected to a source of combustion gases and an outlet (23) fluidly connected to the turbine (15), and a conduit (21) extending around the central axis (A) from the inlet (22) to the outlet (23); a bearing housing (30) extending around the central axis (A); and a flange (60) connecting the scroll case (20) to the bearing housing (30), the flange (60) having a first end (61) connected to the scroll case (20), a second end (62) connected to the bearing housing (30), and a deflecting wall section extending between the first end (61) and the second end (62), the deflecting wall section being spaced apart from the scroll case (20) by an annular gap (G0), the deflecting wall section of the flange (60) being deformable in a radial direction.

A hydrogen fuel nozzle (100) for a gas turbine engine has a plurality of flow circuits (110, 120, 130), each of which is configured to flow one of a fuel or an oxidizer to a gas turbine engine combustor positioned downstream of the hydrogen fuel nozzle (100) when the gas turbine engine is assembled. At least one of the plurality of flow circuits (110, 120, 130) is a hydrogen flow circuit (110) configured to flow hydrogen to the gas turbine engine combustor. The hydrogen fuel nozzle (100) is formed using an additive manufacturing (AM) technique. A non-hydrogen-exposed zone of the hydrogen fuel nozzle (100) is made with a first material which is a conventional high temperature capable alloy. A hydrogen-exposed zone of the hydrogen fuel nozzle (100) is made with a second material, which is a hydrogen-compatible alloy.

A baffle assembly (42) for interacting with a lubricating fluid (44) around a rotor in an aircraft power plant (10) is disclosed. The baffle assembly (42) may mitigate misting of lubricating fluid (44) inside a gearbox (12) and improve the effectiveness of the lubricating fluid (44). The baffle assembly (42) includes a first sidewall (48) configured to be fastened to a first structure (50) disposed to a first side of the rotor, a second sidewall (52) configured to be fastened to a second structure (54) disposed to a second side of the rotor axially opposite the first side of the rotor, and a baffle (56). The baffle (56) is supported by and extends between the first sidewall (48) and the second sidewall (52) so that the baffle (56) is disposed at a position adjacent to and radially outwardly of the rotor during use. A vibration damper (58A,58B) engaged with the baffle (56) to dampen a vibration of the baffle (56) during use.

A hydrogen fuel nozzle (500a) for a gas turbine engine includes a plurality of flow circuits (510a, 510b, 510c), each of which is configured to flow one of a fuel or an oxidizer to a gas turbine engine combustor positioned downstream of the hydrogen fuel nozzle (500a) when the gas turbine engine is assembled. At least one of the plurality of flow circuits (510a, 510b, 510c) is a hydrogen flow circuit configured to flow hydrogen to the gas turbine engine combustor and at least one of the plurality of flow circuits (510a, 510b, 510c) includes one or more swirlers (532a) positioned inside the at least one of the plurality of flow circuits (510a, 510b, 510c) to direct flow through the at least one of the plurality of flow circuits (510a, 510b, 510c). The hydrogen fuel nozzle (500a) is formed using an additive manufacturing (AM) technique from a hydrogen-compatible alloy and wherein an inside diameter of the hydrogen flow circuit has a surface finish selected to limit hydrogen embrittlement.

An aircraft propulsion system (20) includes a compressor section (26), a combustor section (28), and a fuel system assembly (36). The compressor section (26) includes a plurality of inlet guide vanes (42), an inlet guide vane actuator (IGVA) (120), and an IGVA servo (118). The IGVA (120) is operably coupled with the plurality of inlet guide vanes (42). The IGVA servo (118) is connected in fluid communication with the IGVA (120). The combustor section (28) includes a combustor (44). The fuel system assembly (36) includes a jet fuel system (86) and a hydrogen fuel system (88). The jet fuel system (86) is connected in fluid communication with the combustor (44) and the IGVA servo (118). The hydrogen fuel system (88) is connected in fluid communication with the combustor (44) and the IGVA servo (118).

An aircraft propulsion system (20) includes a combustor (44), an engine oil system (34), a fuel system assembly (36), and a control system (38). The engine oil system (34) includes a fuel-oil heat exchanger (80). The fuel system assembly (36) includes a jet fuel system (86) and a hydrogen fuel system (88). The jet fuel system (86) and the hydrogen fuel system (88) are connected in fluid communication with the combustor (44). The control system (38) is configured to operate the fuel system assembly (36) in one of a first fuel mode to direct a jet fuel from the jet fuel system (86) through the fuel-oil heat exchanger (80) or a second fuel mode to direct a hydrogen fuel from the hydrogen fuel system (88) through the fuel-oil heat exchanger (80).

An aircraft propulsion system (20) includes a compressor section (26), a combustor section (28), and a fuel system assembly (36). The compressor section (26) includes a plurality of inlet guide vanes (42), an inlet guide vane actuator (IGVA) (120), and an IGVA servo (118). The IGVA (120) is operably coupled with the plurality of inlet guide vanes (42). The IGVA servo (118) is connected in fluid communication with the IGVA (120). The combustor section (28) includes a combustor (44). The fuel system assembly (36) includes a jet fuel system (86) and a hydrogen fuel system (88). The jet fuel system (86) is connected in fluid communication with the combustor (44). The hydrogen fuel system (88) is connected in fluid communication with the combustor (44) and the IGVA servo (118). The hydrogen fuel system (88) is configured to direct the hydrogen fuel to the IGVA servo (118) to control a position of the IGVA (120).

An aircraft propulsion system (20) includes a combustor (44), an engine oil system (34), and a fuel system assembly (36). The engine oil system (34) includes a fuel-oil heat exchanger (80). The fuel system assembly (36) includes a jet fuel system (86), a hydrogen fuel system (88) and a switching valve (126). The jet fuel system (86) and the hydrogen fuel system (88) are connected in fluid communication with the combustor (44). The jet fuel system (86) is further selectively connected in fluid communication with the fuel-oil heat exchanger (80) by the switching valve (126). The hydrogen fuel system (88) includes a liquid hydrogen source (104). The hydrogen fuel system (88) is further selectively connected in fluid communication with the fuel-oil heat exchanger (80) by the switching valve (126). The switching valve (126) configured in a first circuit flow path directs the jet fuel through the fuel-oil heat exchanger (80). The switching valve (126) configured in a second circuit flow path directs the hydrogen fuel through the fuel-oil heat exchanger (80).

An aircraft propulsion system (20) includes a combustor section (28), an engine oil system (34), and a fuel system (36). The combustor section (28) includes a combustor (44) forming a first portion of a core flow path (68) through the aircraft propulsion system (20). The engine oil system (34) includes a fuel-oil heat exchanger (80). The fuel system assembly (36) includes a jet fuel system (86) and a hydrogen fuel system (88). The jet fuel system (86) is connected in fluid communication with the combustor (44) and configured to selectively direct a jet fuel to the combustor (44). The hydrogen fuel system (88) is connected in fluid communication with the combustor (44) and configured to selectively direct a hydrogen fuel to the combustor (44). The hydrogen fuel system (88) is further connected in fluid communication with the fuel-oil heat exchanger (80).

An aircraft engine (10), has: a turbine (15) rotatable about a central axis (A); a scroll case (20) having an inlet (22) and an outlet (23) connected to the turbine (15), and a conduit (21) extending around the central axis (A) from the inlet (22) to the outlet (23); a bearing housing (30) extending around the central axis (A) and including a support flange (41); and a turbine support case (40) secured to the bearing housing (30), the turbine support case (40) having spokes (44) distributed around the central axis (A) and extending along a direction having an axial component relative to the central axis (A), the spokes (44) extending through the scroll case (20) and radially supported by the bearing housing (30), one of the turbine support case (40) and the scroll case (20) defining lugs (26) circumferentially distributed around the central axis (A), the other of the turbine support case (40) and the scroll case (20) defining slots (48), the turbine support case (40) and the scroll case (20) circumferentially locked to one another via the lugs (26) engaging the slots (48).

An engine includes an engine oil system. The engine oil system includes a pump assembly (70) forming a portion of an oil flow path. The pump assembly (70) includes a drive shaft (72), a plurality of pumps (76), and a pump housing (74). The drive shaft (72) extends along a rotational axis (82). The plurality of pumps (76) are arranged axially along the rotational axis (82). Each of the plurality of pumps (76) includes a pump rotor (104) coupled with the drive shaft (72). The pump housing (74) extends along the rotational axis (82) between a first housing end (86) and a second housing end (88). The pump housing (74) circumscribes and houses the drive shaft (72) and the plurality of pumps (76) between the first housing end (86) and the second housing end (88). The pump housing (74) includes a plurality of housing body portions (90A-E) axially stacked along the rotational axis (82) to form the pump housing (74). The plurality of housing body portions (90A-E) collectively form an internal oil passage (92) through the pump housing (74).

A fan and low-pressure compressor assembly of a gas turbine engine includes a fan (42) operably connected to a first shaft (40) of the gas turbine engine, and a low-pressure compressor rotor (76) operably connected to the first shaft (40), the low-pressure compressor rotor (76) and the fan (42) sharing a common central longitudinal axis. A fan load path (88) of the fan (42) is separate from a low-pressure compressor rotor load path (116) of the low-pressure compressor rotor (76).

An aircraft propulsion system (20) includes a combustor (44), an engine oil system (34), and a fuel system assembly (36). The engine oil system (34) includes a fuel-oil heat exchanger (80). The fuel system assembly (36) includes a jet fuel system (86), a hydrogen fuel system (88), and a switching valve (126). The jet fuel system (86) and the hydrogen fuel system (88) are connected in fluid communication with the combustor (44). The jet fuel system (86) is selectively connected in fluid communication with the fuel-oil heat exchanger (80) by the switching valve (126). The hydrogen fuel system (88) is selectively connected in fluid communication with the fuel-oil heat exchanger (80) by the switching valve (126). The switching valve (126) is configurable between a first circuit flow path and a second circuit flow path. The switching valve (126) configured in the first circuit flow path directs the jet fuel through the fuel-oil heat exchanger (80). The switching valve (126) configured in the second circuit flow path directs the gaseous hydrogen fuel through the fuel-oil heat exchanger (80).

A computer-readable model (200) of a target speckle pattern (118) is constructed with computer-stimulated digital image correlation (DIC) unit that includes a camera (112) and a three-dimensional object model (116). The three-dimensional object model (116) is positioned within a view frustum (210) of the camera (112) and facing the camera (112). A nominal speckle pattern (114) is generated based on the parameters of the camera (112). The camera (112) projects the nominal speckle pattern (114) onto the three-dimensional object model (116) so that a target speckle pattern (118) can be constructed. The target speckle pattern (118) is then saved in a non-transitory computer-readable memory of the computer-readable model (200). The computer-readable model (200) may be executed by a controller of a manufacturing system, such as a CNC machining system or a 3D printing system, to manufacture a stencil to be used in performing digital image correlation (DIC) analyses of object(s).

An engine nacelle assembly includes a nacelle (30) that defines a flow path between an intake end (42) and an exhaust end (44), and a nozzle (38) that is disposed at the exhaust end (44) of the nacelle (30) about a central axis (A). The nozzle (38) includes opposing thrust reverser openings (50) and thrust reverser doors (52). A cross-sectional area transverse to the central axis (A) proximate a forward side (58) of each of the opposing thrust reverser openings (50) is oval shaped with an offset radius (64) within offset arc segments (104) that are aligned with the opposing thrust reverser openings (50) and a first radius (66) outside of the offset arc segment (104). The first radius (66) is less than the offset radius (64).

A method includes controlling a fuel pump (104) to pump fuel through an inlet of a fuel metering valve, FMV, (300), commanding the FMV (300) to open a metering window of the FMV (300) a controllable increment beyond an edge of the metering window, and checking for a presence of fuel downstream from a discharge of the FMV (300).

In response to exiting from a non-full non-filling state of a secondary manifold (320) of a gas turbine engine (102a; 102b), the method of filling the secondary fuel manifold (320) of the engine (102a; 102b) includes: setting at least one engine schedule limit based on a measured value of a scheduling parameter that is within a first range that corresponds to a non-full state of the secondary fuel manifold. The method includes determining whether the filling of the secondary fuel manifold is completed. The method includes adjusting the at least one engine schedule limit to control at least one operational parameter of the engine (102a; 102b), in response to a determination that the filling of the secondary fuel manifold (320) is completed.

A gas turbine engine (201) of an aircraft is provided. The gas turbine engine (201) includes a shaft (210), a housing (220) and an auxiliary ball bearing assembly (240) to rotatably support the shaft (210) within the housing (220). The auxiliary ball bearing assembly (240) includes inner and outer races (250, 260), ball bearings (270) between the inner and outer races (250, 260), an axial spring (281) anchored between the housing (220) and the outer race (260) to prevent axial movement of the shaft (210) during overhaul and an oil cavity (290) defined between the housing (220) and the outer race (260) for axial damping during the overhaul and for radial damping during operational conditions.

A method for increasing oxidation resistance of an alloy surface comprises scanning a high energy source onto the alloy surface to melt the alloy surface; after the melted alloy surface solidifies, re-melting at least a partial area of the alloy surface using the high energy source to thereby form a re-melt layer, such that the alloy surface comprises a first oxidation resistance, the re-melt layer comprises a second oxidation resistance, and the second oxidation resistance is greater than the first oxidation resistance.

An aircraft propulsion system (20) and a method of controlling a compressor (23) operating line in a propulsion system (20) for an aircraft is provided. The system (20) includes a gas turbine engine (22), an electric motor (34; 36; 41), a gearbox (28), and a propulsor (26). The gas turbine engine (22) has a compressor (23), a combustor (25), a turbine (27), and an engine shaft (29) extending between the compressor (23) and the turbine (27). The gas turbine engine (22) is in drive engagement with the gearbox (28), and the propulsor (26) is driven by the gearbox (28). The method includes: controlling the gas turbine engine (22) to provide motive force to the propulsor (26); determining an operational line of the compressor (23); and controlling the operational line of the compressor (23) by controlling the electric motor (34; 36; 41) to change an amount of load on the compressor (23) attributable to the electric motor (34; 36; 41).

A hydrogen fuel nozzle (100) for a gas turbine engine includes a plurality of flow circuits (110, 120, 130), each of which is configured to flow one of a fuel or an oxidizer to a gas turbine engine combustor positioned downstream of the hydrogen fuel nozzle (100) when the gas turbine engine is assembled. At least one of the plurality of flow circuits (110, 120, 130) is a hydrogen flow circuit configured to flow hydrogen to the gas turbine engine combustor. The hydrogen fuel nozzle (100) is formed using an additive manufacturing (AM) technique from a hydrogen-compatible alloy and wherein an inside diameter of the hydrogen flow circuit has a surface finish selected to limit hydrogen embrittlement.

An anti-rotation tool (100) includes a shaft (102), a head (104), a roller band (106), and a locking device (108). The head (104) is disposed at the shaft (102). The head (104) forms a cradle (130). The cradle (130) forms a component mating surface (132). The roller band (106) extends between and to a proximal band end (138) and a distal band end (140). The proximal band end (138) is fixedly mounted to the head (104). The roller band (106) includes a first lateral chain (146), a second lateral chain (148), a plurality of pins (150), and a plurality of rollers (152). Each of the plurality of pins (150) extends between and to the first lateral chain (146) and the second lateral chain (148). Each of the plurality of rollers (152) is rotatably mounted on a respective one of the plurality of pins (150). The plurality of rollers (152) form a component mating interface (162). The locking device (108) is positionable in a locked condition or an unlocked condition with the roller band (106).

An integrally bladed rotor, IBR, (201) is provided. The IBR (201) includes a hub (210) and blades (220). Each blade (220) includes an airfoil section (230) to aerodynamically interact with a flow of air for compressing the air and a root fillet (240) integrally formed with the hub (210) and from which the airfoil section (230) integrally extends. The root fillet (240) includes a curved exterior surface (250) having points of tangency with the airfoil section (230) to define an outboard extent (241) of the root fillet (240) and a root fillet body (260) between the hub (210) and the outboard extent (241) of the root fillet (240) and defining a cutout (270) for arresting a crack that is most likely to form and propagate.

An electric assembly (24) for an aircraft (18) is provided that includes a battery string (60), a switch (94), a battery charging system (56), and a charge monitoring device (50). The battery string includes a plurality of battery modules (62) and a contactor (72). The plurality of battery modules are electrically connected in series. The contactor is disposable in an open configuration or a closed configuration. The battery charging system is in communication with the battery string, and configured to perform a charging operation of the battery string. The charge monitoring device includes a processor and non-transitory memory. The non-transitory memory stores instructions executable by the processor, causing the processor to determine a charging parameter of the contactor, identify the presence of a fault condition for the contactor, and signal the switch to actuate the contactor to the open configuration. The presence of the fault condition indicates the charging parameter exceeds the fault condition threshold.

A propulsion system (20) includes a battery (64), a plurality of temperature sensors (110), and a controller (92). The battery (64) includes a plurality of battery cells (112). Each temperature sensor of the plurality of temperature sensors (110) is disposed at a respective battery cell of the plurality of battery cells (112). The controller (92) includes a first control channel (94) and a second control channel (96). The first control channel (94) is connected in signal communication with the plurality of temperature sensors (110). The second control channel (96) is connected in signal communication with the plurality of temperature sensors (110). Each of the first control channel (94) and the second control channel (96) is configured to monitor a temperature of each battery cell of the plurality of battery cells (112) measured using the plurality of temperature sensors (110).

A propulsion system (20) includes a battery (64), a plurality of temperature sensors (110), and a controller (92). The battery (64) includes a plurality of battery cells (112). The plurality of temperature sensors (110) includes a plurality of first temperature sensors (110A; 110C; 110E; 110F) and at least one second temperature sensor (110B). Each temperature sensor of the plurality of first temperature sensors (110A; 110C; 110E; 110F) is disposed at a respective battery cell of the plurality of battery cells (112). The at least one second temperature sensor (110B) is disposed at the plurality of battery cells (112). The controller (92) includes a first control channel (94) and a second control channel (96). The first control channel (94) is configured to monitor a T1 temperature of each battery cell of the plurality of battery cells (112) measured using the plurality of first temperature sensors (110A; 110C; 110E; 110F). The second control channel (96) is configured to monitor at least one T2 temperature of the plurality of battery cells (112) measured using the at least one second temperature sensor (110B).

A validation tool (100) for determining conformity of leading edges (56) of vanes (32) on an impeller (22) of an aircraft engine (10) includes a support (102) and a probe element (104). The support (102) has a base (103) and a spindle (105) extending away from the base (103), the base (103) being configured to be received on a reference surface (60) of the impeller (22) and the spindle (105) defining a spindle axis (106). The probe element (104) is rotatably mounted on the support (102) and includes an annular body (110) defining a central bore (111) matingly receiving the spindle (105) of the support (102) therein. The probe element (104) is rotatable about the spindle axis (106) and includes a conformal probe (120) thereon extending axially toward the base (103) of the support (102). The conformal probe (120) is configured to have a clearance fit with multiple of the leading edges (56) at a same span-wise position of each, when the probe element (104) is rotated on the spindle (105) of the support (102).

A propulsion system (20) includes a battery (64), a plurality of temperature sensors (110), and a controller (92). The battery includes a plurality of battery cells (112). Each temperature sensor of the plurality of temperature sensors is disposed at a respective battery cell of the plurality of battery cells. The controller includes a first control channel (94). The first control channel is configured to monitor a temperature of each battery cell of the plurality of battery cells measured, execute a model (136) trained to identify a faulted condition or an unfaulted condition of each temperature sensor of the plurality of temperature sensors using one or more operating parameters of the propulsion system, and identify the faulted condition or the unfaulted condition of a first temperature sensor of the plurality of temperature sensors using the model to determine an expected temperature range for the first temperature sensor based on the one or more operating parameters.

An aircraft engine (10), has: a turbine (15) including a turbine rotor (15C) rotatable about a central axis (A); a scroll case (20) having an inlet (22) connected to a source of combustion gases (19) and an outlet (23) fluidly connected to the turbine (15), and a conduit (21) extending around the central axis (A) from the inlet (22) to the outlet (23); a bearing housing (30); a turbine support case (40) secured to the bearing housing (30), the scroll case (20) disposed axially between the turbine support case (40) and the bearing housing (30), the turbine support case (40) having spokes (44) extending through the scroll case (20) and radially supported by the bearing housing (30); and heat shields (60) distributed around the central axis (A) in circumferential alignment with the spokes (44) and extending axially through the scroll case (20) for thermally shielding the spokes (44), a heat shield of the heat shields (60) disposed around a spoke of the spokes (44), the heat shield (60) defining an air gap (G1) surrounding the spoke (44) inside the scroll case (20).

A powerplant (20) assembly includes first (64) and second (66) powerplant components, a plurality of fasteners (154) and an insert (70). The first powerplant component includes a first component mount (78) and a plurality of first component apertures. The first component apertures are arranged circumferentially about an axis (72) and extend axially through the first component mount. The first component apertures include a plurality of first component fastener apertures (88) and a first component insert aperture (90). The second powerplant component includes a second component mount (110) and a plurality of second component fastener apertures (118). The fasteners attach first and second component mounts together. Each of the fasteners is mated with a respective one of the first component fastener apertures and a respective one of the second component fastener apertures. An insert (70) is threaded into the first component insert aperture. The insert includes a threaded bore (144).

A powerplant (20) is provided for an aircraft. This aircraft powerplant (20) includes a compressor section (40), a fuel-air heat exchanger (86), a recuperator (88), a combustor section (41) comprising a combustor (68), a turbine section (42), a flowpath (50) and a fuel delivery system (28). The flowpath (50) extends from an airflow inlet (64) into the flowpath (50), through the compressor section (40), the fuel-air heat exchanger (86), the recuperator (88), the combustor section (41) and the turbine section (42), to a combustion products exhaust (66) from the flowpath (50). The fuel delivery system (28) includes a fuel source (74), a fuel circuit (76) and a fuel injector (72). The fuel circuit (76) fluidly couples the fuel source (74) to the fuel injector (72) and extends through the fuel-air heat exchanger (86). The fuel injector (72) is configured to introduce fuel into the flowpath (50) for combustion within the combustor (68).

An engine (22) includes a bearing assembly (58). The bearing assembly (58) includes a bearing compartment housing (60), a first bearing (62), and a second bearing (64). The bearing compartment housing (60) extends circumferentially about a rotational axis (40). The bearing compartment housing (60) includes an outer housing body (74), an inner housing body (72), and a plurality of struts (76). The outer housing body (74) surrounds and forms a bearing compartment (70). The inner housing body (72) extends between and to an inner radial side (84) and an outer radial side (82). The inner housing body (72) forms a plurality of transverse slots (92) extending through the inner housing body (72) from the inner radial side (84) to the outer radial side (82). The plurality of struts (76) connect the inner housing body (72) and the outer housing body (74). The first bearing (62) is mounted on the inner housing body (72) at a first axial position. The second bearing (64) is mounted on the inner housing body (72) at a second axial position.

An aircraft engine (10), has: a turbine (15) including a turbine rotor (15C) rotatable about a central axis (A); a scroll case (20) having an inlet (22) connected to a source of combustion gases (19) and an outlet (23) connected to the turbine (15), and a conduit (21) extending around the central axis (A) from the inlet (22) to the outlet (23); a bearing housing (30) including a support flange (41); a turbine support case (40) secured to the bearing housing (30), the scroll case (20) disposed between the turbine support case (40) and the bearing housing (30), the turbine support case (40) having spokes (44) extending through the scroll case (20) and radially supported by the bearing housing (30), a spoke of the spokes (44) having a distal end (44B) secured to the support flange (41) via: one or more fasteners (47), and a bracket (60) secured to the support flange (41), the bracket (60) engaged in a slot (44D) having an axially facing surface (44E) trapped between the bracket (60) and the support flange (41) of the bearing housing (30).