Pratt & Whitney Canada Patente

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 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 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 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).

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).

A mounting arrangement includes a casing (17) of an aircraft engine (10) extending around an axis (11), an accessory (22), and a bracket (40) for mounting the accessory to the casing (17). The bracket (40) includes a first bracket segment (44') coupled to the casing (17) at a first mounting point, a second bracket segment (48') coupled to the accessory at a second mounting point, and a wire rope isolator (54') including a looped wire (56') coupling the first bracket segment (44') to the second bracket segment (48'). The first bracket segment (44') at the first mounting point and the second bracket segment (48') at the second mounting point extend in intersecting planes.

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 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 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).

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 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 fuel injector (32) for a gas turbine engine combustor (26) includes a swirler (34), a mounting stage (38), and a distributor (36). The swirler (34) has a shaft (52), a collar (54), a throat section (56), and first and second axial ends (48, 50). The throat section (56) includes an inner radial surface (86) that defines a central passage (88) that extends between the swirler inner bore (160) and the collar (54). The collar (54) includes a plurality of apertures (78, 80) extending therethrough disposed radially outside of the central passage (88). The distributor (36) is selectively positionable relative to the throat section (56).

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 (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.

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 (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 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).

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).

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 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).

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 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).

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 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 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.

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.

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 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).

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 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).

A hybrid aircraft power plant has a gas turbine engine (10) has a high-pressure spool (25) including a high-pressure compressor (25a), a high-pressure turbine (25b), and a high-pressure shaft (25c) drivingly engaging the high-pressure turbine (25b) to the high-pressure compressor (25a), a low-pressure spool (24) including a low-pressure compressor (24a), a low-pressure turbine (24b), and a low-pressure shaft (24c) drivingly engaging the low-pressure turbine (24b) to the low-pressure compressor (24a), an electric motor (32) drivingly engaged to the low-pressure shaft (24c), and a torque-transmitting device operatively connected to the high-pressure shaft (25c) and having an engaged configuration in which the torque-transmitting device drivingly engages the electric motor (32) to the high-pressure shaft (25c) and a disengaged configuration in which the torque-transmitting device disconnects the electric motor (32) from the high-pressure shaft (25c).

A fluid cooler (24) for installation in a bypass duct (22) of a turbofan gas turbine engine and associated methods are provided. The fluid cooler (24) includes an inlet duct (62), a heat exchanger (40) and an outlet duct (64). The inlet duct (62) includes an inlet (32) protruding into the bypass duct (22) to receive a portion of the bypass air (b) into the inlet duct (62). The heat exchanger (40) is in fluid communication with the inlet duct (62). The heat exchanger (40) facilitates heat transfer between a fluid (F) and the portion of bypass air (b) received into the inlet duct (62). The heat exchanger (40) defines a general flow direction for the portion of bypass air (b) that is different from the main flow direction (D1) of bypass air (B) inside the bypass duct (22). The outlet duct (64) conveys the portion of bypass air (b) from the heat exchanger (40) back to the bypass duct (40).

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).

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.