Pratt & Whitney Canada Patente

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 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 method for operating an aircraft (22) having a first propulsion unit (28A) and a second propulsion unit (28B). During this method a propulsion unit-inoperative event is identified where the first propulsion unit (28A) is inoperative and the second propulsion unit (28B) is operative. A trimmed parameter for the second propulsion unit (28B) is converted to an actual parameter for the second propulsion unit (28B). The actual parameter is monitored over a period of time the second propulsion unit (28B) operates during the propulsion unit-inoperative event. An action is performed in response to determining that the second propulsion unit (28B) has operated for at least a time threshold with the actual parameter being equal to or greater than a parameter threshold.

Systems and methods disclosed herein account for the crosswind effects when assessing the engine field margins based on a run in an outdoor facility where no thrust measurement is available. The impact of crosswinds on engine performance may be quantified by observing the differences in several engine performance parameters available in the field against a performance cycle model reference stored in data tables. The differences are compared to Average New Engine (ANE) margins to result in available margins.

An aircraft engine (10), has: a turbine (15); 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 from the inlet (22) to the outlet (23); a bearing housing (30) including a support flange; an exhaust case (15B) downstream of the turbine (15); and a turbine support case (40) secured to the bearing housing (30) and to the exhaust case (15B), the turbine support case (40) having spokes (44) extending along a direction having an axial component, the spokes (44) extending through the scroll case (20) and radially supported by the bearing housing (30), a spoke having a distal end (44B) secured to the support flange via: one or more fasteners (47), and a retaining member (48) at the distal end (44B), the retaining member (48) defining an abutment face (48A) facing an axial direction and circumferentially overlapping the support flange, a portion of the support flange located axially between the distal end (44B) and the abutment face (48A).

An aircraft powerplant (20) includes a core (28) of a gas turbine engine (24) and an exhaust section (26). The core includes a core flowpath (40), a compressor section (36), a combustor section (37) and a turbine section (38). The core flowpath extends longitudinally through the compressor section, the combustor section and the turbine section. The exhaust section includes an exhaust flowpath (60) and a flow diverter (66). The flow diverter includes a diverter panel (70) configured to pivot between a first position and a second position. The flow diverter is configured to fluidly couple the diffuser section (72) to the first nozzle section (76) and fluidly decouple the diffuser section from the second nozzle section (82) when the diverter panel is in the first position. The flow diverter is configured to fluidly couple the diffuser section to the second nozzle section and fluidly decouple the diffuser section from the first nozzle section when the diverter panel is in the second position.

A combustion liner v-band includes a louver segment (44) having an arcuate member (52), an array of scoops (54), and outlet openings (56). The arcuate member (52) includes an interior wall (60) and side walls (62) extending partially circumferentially about a centerline axis (C) of the combustor (28) to bound a channel (58) that is open at a radial outer boundary of the arcuate member (52). The scoops (54) extend from the interior wall (60). The outlet openings (56) extend through the interior wall (60) and the scoops (54) to place the channel (58) in fluid communication with a region exterior to the louver segment (44). Each outlet opening (56) extends along a discharge axis (A) that has a nonzero axial angle component (α) and a nonzero circumferential angle component (β).

An assembly (20) includes a case (22), a housing (24) and a component (26). The case (22) includes an aperture (54), an exterior side (48) and an interior side (46). The aperture (54) extends axially along an axis (40) through the case (22) from the exterior side (48) to the interior side (46). The interior side (46) forms a peripheral boundary of an internal volume (32) of the aircraft engine. A cavity (72) is formed by and extends axially between the housing (24) and the case (22). The component (26) includes a projection (94) and a mount (92) that is disposed within the cavity (72). The projection (94) projects axially out from the mount (92), through the aperture (54), and into the internal volume (32) to a distal end of the component (26). The mount (92) includes a rim and a plurality of spring beams. The spring beams bias the rim axially towards the case (22) to provide a sealed engagement axially between the mount (92) and the case (22).

An apparatus is provided for a turbine engine. This apparatus includes a rotor blade (72) configured to rotate about an axis (24). The rotor blade (72) includes a base (78) and an airfoil (76). The base (78) includes a platform (82), an attachment (86), a pocket (88) and a contoured side surface. The pocket (88) is located radially between the platform (82) and the attachment (86). The pocket (88) projects radially outward into the base (78) and along the contoured side surface to the platform (82). The pocket (88) extends axially within the base (78) and along the contoured side surface between opposing axial ends of the pocket (88). The pocket (88) projects laterally into the base (78) to the contoured side surface. The contoured side surface includes a plurality of convex sections arranged axially along the axis (24) between the opposing axial ends of the pocket (88). The airfoil (76) projects radially outward from the platform (82) to a tip (96) of the airfoil (76).

A compressor assembly (44) of a gas turbine engine includes: a case (100) surrounding a core; a row of stator blades (112) within the case, wherein the row of stator blades has first shrouds (114); a row of rotor blades (132) within the case and adjacent to the row of stator blades, wherein the row of rotor blades has second shrouds (134). The first shrouds define first mate faces (118) that axially face the second shrouds, and the second shrouds define second mate faces (138) that axially face the first shrouds. A bleed port groove (150) is defined between the first mate faces and the second mate faces. A bleed port (155) is defined along the bleed port groove, a bleed cavity (160) is defined in the case. The bleed cavity is exterior to the first shrouds and the second shrouds, and fluidly coupled to the core via the bleed port. A circumferentially extending first wire (170) is secured within the bleed port groove.

An apparatus (50) for stabilizing a fan assembly of a gas turbine engine during transport includes a ring (52) configured to engage a nose cone (54) of the fan assembly; an outer bracket (56) connected to the ring (52) and configured to mount to a fan casing of the fan assembly; and a tensioning assembly between the ring (52) and the outer bracket (56) to move the ring (52) along an axis of a shaft of the fan assembly. An assembly with the gas turbine engine and a method are also disclosed.

A gas turbine engine (20) is provided that includes a compressor, a turbine, an engine shaft (30), a gearbox (38), and a tower shaft assembly (40). The engine shaft (30) is engaged with the compressor and the turbine. The tower shaft assembly (40) is in drive engagement with the engine shaft (30) and the gearbox (38). The tower shaft assembly (40) includes a tower shaft (32), a first bearing (48), and a first gear (44). The tower shaft (32) extends axially along a rotational axis (54) between first and second ends (50, 52). The tower shaft (32) has a support shoulder (60) having a non-circular configuration. The first bearing (48) is engaged with the support shoulder (60) of the tower shaft (32). The first bearing (48) has roller elements (100) and an inner ring (96) engaged with the roller elements (100). The first gear (44) is engaged with the tower shaft (32) and the first bearing (48), and is configured for engagement with the gearbox (38).

A gearbox (301) of a gas turbine engine of an aircraft with a two-bearing arrangement includes a coupling shaft (310) rotatable about a rotational axis (A) thereof, an output gear (320) including a first gear element (330), a gear web (34) and a bearing assembly (350). The gear web (340) is rotatable with the coupling shaft (310) and includes a shaft (341) connected with the coupling shaft (310), a rim (342) including a second gear element (3421) disposed to form a gear mesh (360) with the first gear element (330) and a web (343) extending from the shaft (341) to the rim (342). The bearing assembly (350) includes a fixed bearing support (351) and a bearing element (352) supportively disposable between the fixed bearing support (351) and the rim (342) of the gear web (340).

An engine includes an engine rotational assembly and an engine oil system (34). The engine rotational assembly includes an engine shaft (50) and a rotor. The engine rotational assembly is rotatable about a first rotational axis (42). The engine oil system (34) includes a pump assembly (70) including a drive shaft (72), a pump (76), and a pump housing (74). The drive shaft (72) is coupled with the engine shaft (50). The drive shaft (72) extends along a second rotational axis (82) between and to a first shaft end (78) and a second shaft end (80). The drive shaft (72) includes a hand crank pad (86) at the second shaft end (80). The hand crank pad (86) includes a rotational drive member (88). The pump (76) includes a pump rotor (104) coupled with the drive shaft (72). The pump housing (74) extends along the second rotational axis (82) between a first housing end (90) and a second housing end (92). The second housing end (92) forms an opening (94) at the hand crank pad (86).

A powder bed fusion additive manufacturing system (100') includes a build powder bed (106') positioned on a build plate (108') in a build chamber (110'), an energy source scanning system (115') configured to scan an energy source (116') across a top layer (118') of the build powder bed (106') to consolidate selected portions of the top layer (118') to build a single layer of a desired part (120'), a build piston (112') configured to adjust the height of the build plate (108') after the layer of the desired part (120') is built, and a fresh build powder distributor (126') configured to distribute a layer fresh build powder (104') over a build powder bed (106') such that the layer of fresh build powder (104') is level and smooth. The fresh build powder distributor (126') comprises a distributor excitation device (126a') configured to cause the fresh build powder distributor (126') to vibrate at a predetermined vibrational frequency while distributing the layer of fresh build powder (104') over the build powder bed (106').

A method is provided for assembling a structure (20) of an aircraft powerplant. This method includes mounting a seal runner (32) onto a rotor (30) using an installation tool (108), the installation tool (108) including an inner locating surface (118) and an outer locating surface (120), the inner locating surface (118) axially recessed from and radially inboard of the outer locating surface (120). The mounting of the seal runner (32) includes axially pressing a runner mount (72) onto the rotor (30) by moving the installation tool (108) axially towards the rotor (30), with the outer locating surface (120) axially pressed against a distal end (68) of the seal runner (32), until the inner locating surface (118) axially engages the distal end (46) of the rotor (30). After the seal runner (32) is mounted onto the rotor (30) using the installation tool (108), the runner mount (72) is attached to the rotor (30) through a radial interference fit between the runner mount (72) and the rotor (30) and the runner mount (72) is axially disengaged from the rotor (30)°.

A method for detecting relative movement of a position of beta sensors (112) comprises detecting a characteristic associated with an output waveform from at least two beta sensors (112) generated responsive to passage of an outlier indicator (102) of a beta ring (104) past the at least two beta sensors (112), comparing the detected characteristic with a reference point for the characteristic associated with the output waveform for the at least two beta sensors (112) within a controller (115), determining if the detected characteristic associated with the output waveform is different from the reference point for the characteristic associated with the output waveform by a first amount within the controller (115) and generating a fault code indicating that the at least one beta sensor (112) has moved at least the first amount. The first amount indicates that the at least one beta sensor (112) has moved the first amount from a previous position.

A method includes detecting a characteristic associated with an output waveform from a beta sensor (112) generated responsive to passage of an indicator (102) of a beta ring (104) past the beta sensor (112), comparing the detected characteristic with a reference point for the characteristic associated with the output waveform for the beta sensor (112) within a controller (115), determining if the detected characteristic associated with the output waveform is different from the reference point for the characteristic associated with the output waveform by a first amount within the controller (115), wherein the first amount indicates that the beta sensor (112) has moved the first amount from a previous position with respect to the beta ring (104) and in response to the determining that the detected characteristic is different from the reference point by the first amount, generating a fault code indicating that the beta sensor (112) has moved at least the first amount with respect to the beta ring (104).

A method of and tool for positionally aligning a gearbox (34) of a gas turbine engine (20) is provided. The method includes: providing an alignment tool (48) that includes a plurality of probe pins (56) and a support flange (52); attaching the alignment tool (48) to an engine case (40) of the gas turbine engine (20); engaging the gearbox (34) with the engine case (40); positionally aligning the gearbox (34) by contacting a first surface of the gearbox (34) with a first probe pin of the plurality of probe pins (56), and contacting a second surface of the gearbox (34) with a second probe pin of the plurality of probe pins (56); and securing the gearbox (34) to the engine case (40) while the first probe pin is in contact with the first surface of the gearbox (34) and the second probe pin is in contact with the second surface of the gearbox (34).

A gas turbine engine (20) is provided that includes a compressor, a turbine, and engine shaft (30), a gearbox (38), and a tower shaft assembly (40). The engine shaft (30) is engaged with the compressor and the turbine. The tower shaft assembly (40) is in drive engagement between the engine shaft (30) and the gearbox (38). The tower shaft assembly (40) includes a tower shaft (32), a first bearing (48), and a first gear (44). The tower shaft (32) extends axially along a rotational axis (54). The first bearing (48) is engaged with the tower shaft (32). The first bearing (48) has a plurality of roller elements (100) and an inner ring (96) engaged with the roller elements (100). The first gear (44) is engaged with the tower shaft (32) and extends axially between first gear first and second axial end surfaces (114, 116). A first interference fit exists between the first gear (44) and the inner ring (96) of the first bearing (48). The first gear (44) is configured for engagement with the gearbox (38).

An engine oil system (34) includes an oil tank (68) and a scavenge pump assembly (70). The scavenge pump assembly (70) includes a drive shaft (72), a plurality of scavenge pumps (76), and a scavenge return manifold (78). Each of the plurality of scavenge pumps (76) includes a pump rotor (88) coupled with the drive shaft (72). The scavenge return manifold (78) includes a manifold body (90) forming a common outlet passage (92) and a plurality of inlet passages (94) of the scavenge return manifold (78). The common outlet passage (92) includes a manifold outlet (98) of the scavenge return manifold (78) connected in fluid communication with the oil tank (68). Each of the plurality of inlet passages (94) includes a manifold inlet (96) of the scavenge return manifold (78) connected in fluid communication with a respective one of the plurality of scavenge pumps (76). Each of the plurality of inlet passages (94) extends in an oblique direction from the manifold inlet (96) to the common outlet passage (92).

A first fuel injector (64) includes a first injector base (76), a first injector flange (78) and a first gaseous fuel passage (86). The first injector flange is connected to the first injector base at a distal end (74) of the first fuel injector. The first injector flange projects radially out from the first injector base to an outer rim (92) of the first fuel injector. The first gaseous fuel passage includes a first passage upstream section (130) and a first passage downstream (132) section fluidly coupled to the first passage upstream section. The first passage upstream section extends axially within the first injector base. The first passage downstream section extends radially through the first injector flange to an outlet (136) in the outer rim. A trajectory (140) of the first gaseous fuel passage at the outlet is angularly offset from the first axis by a first acute offset angle.

An assembly (20) includes a rotating seal land (30), a seal element (68), a stationary seal support (34) and a mating plate (82). The seal element (68) axially overlaps, circumscribes and radially sealingly engages the rotating seal land (30). The stationary seal support (34) axially overlaps and circumscribes the rotating seal land (30) and the seal element (68). The stationary seal support (34) includes a support notch (48), a support shoulder (50), a support inner interior surface (52) and a support outer interior surface (54). The support notch (48) projects axially into the stationary seal support (34) from a distal end (36) of the stationary seal support (34) to the support shoulder (50). The support notch (48) projects radially outward away from an axis (28) into the stationary seal support (34) from the support inner interior surface (52) to the support outer interior surface (54). The mating plate (82) is seated in the support notch (40) and mounted to the stationary seal support (34). The mating plate (82) axially contacts the support shoulder (50) and the seal element (68).

Centrifugal compressors (300; 400; 500; 600; 700) for gas turbine engines include a casing (402; 506; 604; 702) and a rotor arranged within the casing (402). The rotor includes a hub (306; 402) and blades (304; 504; 602). Each blade (304; 504) extends from an inlet (308; 412; 512) to an outlet (310; 514) of the compressor (300...700). A casing treatment (318; 404; 502; 610; 704; 800) is applied to the casing (402) that includes a set of recirculators (318a,b; 408a,b; 522; 612, 614; 706; 802, 804; 902) that each define a recirculator passage (414a,b; 524; 618; 710) that extends from a passage inlet (416a,b; 526; 616, 622; 708; 806; 904, 906) to a passage outlet (418a,b; 528; 620, 626; 712; 808; 908). A recirculation flow through the recirculators (318a,b...902) is in a recirculation flow direction counter to a main flow direction. An inlet of each recirculator (318a,b...902) is at a position downstream relative to a respective outlet of the same recirculator (318a,b...902) and each inlet and outlet has a dimension in a main flow direction defined based on a relationship relative to a leading edge span of the blades (304; 504; 602).

An engine oil system (34) includes an oil tank (68) and an access port cover (74). The oil tank (68) includes a wall forming an internal oil cavity (76). The wall forms an access port (80) extending along an access port axis (82) through the wall. The access port cover (74) is mounted on the wall and seals the access port (80). The access port cover (74) includes a cover body (98) and a packing ring (100). The cover body (98) extends between and to an inner axial end (102) and an outer axial end (104). The cover body (98) forms a cover lip (114) at the outer axial end (104). The cover lip (114) extends circumferentially about the access port axis (82). The cover lip (114) extends radially outward and axially outward, relative to the access port axis (82), along a lip axis (124) to the outer axial end (104). The packing ring (100) forms a fluid seal between the cover body (98) and the wall within the access port (80).

An engine oil system (34) for an aircraft engine includes an oil tank (68) and an oil tank return (78). The oil tank (68) forms an internal oil cavity (74). The oil tank return (78) forms a portion of an oil flow path (64) through the engine oil system (34). The oil tank return (78) includes a return body (80) forming a return passage (88) extending between and to a passage inlet (90) and a passage outlet (92). The return body (80) includes an inner diameter wall (106) and an outer diameter wall (108). The inner diameter wall (106) and the outer diameter wall (108) form a turn (104) of the return passage (88) between the passage inlet (90) and the passage outlet (92). The return body (80) forms a top side (114), a corner (116), and an oblique side (118) at the passage outlet (92). The corner (116) and the oblique side (118) are disposed at the outer diameter wall (108). The corner (116) extends between the top side (114) and the oblique side (118).

An engine oil system (34) includes a conduit (72) and a strainer assembly (74). The conduit (72) forms a portion of an oil flow path (64). The conduit (72) includes a first conduit body (500) and a second conduit body (502). The first conduit body (500) is mounted to the second conduit body (502) at a flange interface (512). The first conduit body (500) and the second conduit body (502) form an internal passage (516) of the conduit (72). The strainer assembly (74) is disposed within the conduit (72) at the flange interface (512). The strainer assembly (74) includes a tubular body (518), a first packing ring (520), a second packing ring (522), and a strainer member (524). The first packing ring (520) forms a first fluid seal between the tubular body (518) and the first conduit body (500). The second packing ring (522) forms a second fluid seal between the tubular body (518) and the second conduit body (502). The strainer member (524) is mounted to the tubular body (518) within the oil flow path (64).

A method for correcting extrinsic loading effects on kinematic accuracy of a machining assembly (20) includes determining, for a positioning system (32) of the machining assembly (20) including a cantilevered arm (44) and a rotary table (34) mounted on the cantilevered arm (44), a first position and a first orientation of a first axis (54) and a second position and a second orientation of a second axis (56), filling a machining tank (46) with a dielectric fluid, determining a third position and a third orientation of the first axis (54) and a fourth position and a fourth orientation of the second axis (56), and determining position and orientation correction factors for the first axis (54) and the second axis (56) using a position and orientation deviation of the first axis (54) and the second axis (56) in the first condition from the first axis (54) and the second axis (56) in the second condition.

A rotor blade (72) includes an airfoil (86), a root (90), a cooling air passage (122) and a cooling air deflector (94). The root (90) extends axially along an axis (24) between a first end (114) and a second end (116). The root (90) extends laterally between a first side (118A) and a second side (118B). The root (90) projects radially inward and away from the airfoil (86) to an inner end (112) of the root (90). The cooling air passage (122) includes a passage inlet (124) disposed at the inner end (112). The cooling air passage (122) projects radially into the rotor blade (72) from the passage inlet (124). The cooling air deflector (94) projects radially inward from the inner end (112) of the root (90) to an inner end (130) of the cooling air deflector (94). The cooling air deflector (94) is disposed at the second side (118B) and is spaced laterally from the first side (118A). The cooling air deflector (94) includes a cooling air aperture (150).

A gear train (24) includes an output shaft (160), a clutch assembly (70), and an output gear roller bearing (82). The output shaft (160) is configured for rotation about a rotational axis (76). The clutch assembly (70) includes an inner ring-shaft (108), an output gear (78), a clutch (80), and a clutch ball bearing (84). The inner ring-shaft (108) extends between and to a first outer radial side (112) and a first inner radial side (114). The inner ring-shaft (108) is mounted on the output shaft (160) at the first inner radial side (114). The output gear (78) includes an outer ring (98). The outer ring (98) extends between and to a second outer radial side (100) and a second inner radial side (102). The clutch (80) is disposed between the outer ring (98) and the inner ring-shaft (108). The clutch ball bearing (84) is disposed between the first outer radial side (112) and the second inner radial side (114). The output gear roller bearing (82) is in rolling engagement with the outer ring (98) on the second outer radial side (112).

A labyrinth seal for an aircraft engine includes a base plate (12) for mounting to a housing (10), a first annular insert (14) with a stationary seal element for attaching to the base plate (12), an annular base (22) for mounting to a shaft (20), and second annular insert (24) with a rotating seal element for attaching to the annular base (22). Each of the base plate (12), annular base (22), and the first (14) and second (24) inserts with the seal elements, such as fins, may be made of different materials, which allows clearances (including axial clearances) of the seal elements to be controlled more precisely and may allow reduced wear. Base elements (base plate (12) and/or annular base (22)) and seal elements (first (14) and/or second (24) annular inserts) may be respectively attached to each other by a friction fit (50), an adhesive (70), a weld (60), a plurality of bolts (80), or a plurality of rivets (90).

An assembly is provided for an aircraft. This aircraft assembly includes a first mount (28), a second mount (30), a mounting assembly (20) and a fluid reservoir structure. The mounting assembly (20) connects the first mount (28) to the second mount (30) along an axis (26) while providing controlled intercomponent movement between the first mount (28) and the second mount (30) to reduce internal stress and/or wear. The mounting assembly (20) includes an outer bushing (32), an inner bushing (34) and a fastener (36). The fluid reservoir structure includes one of the first mount (28) or the second mount (30).

A system for an aircraft (1) comprises an electric energy module (108) and a connector module (114c) operatively connected to the electric energy module (108). The connector module (114c) is structured to attach the electric energy module (108) to the aircraft (1) when the connector module (114c) is in an engaged mode while in use, and operable between the engaged mode and a disengaged mode, the connector module (114c) being in the disengaged mode while in use disengaging the electric energy module (108) from the aircraft (1).

A protective cap (10) for an electrical connector of a wire harness has a body (12) made of an electrically conductive material. The body (12) has an end wall (12a) and a sleeve (12b) projecting from a first side of the end wall (12a) around a central axis (14). The sleeve (12b) has threads for threaded engagement with corresponding threads of the electrical connector. The sleeve (12b) and the end wall (12a) defines a receptacle (12f). A grounding/bonding feature (12c) is integrated on an outer surface of the sleeve (12b) or the end wall (12a). The grounding/bonding feature is electrically connectable to an electrically conductive component to electrically protect the unmated connector.

A propulsion system (20) for an aircraft includes a propulsion rotor, and at least one engine (22) mechanically coupled to the propulsion rotor via a geartrain (60). The propulsion system (20) also includes a first propulsor module (24) having a first housing, a first electric motor (112), and a first bladed propulsor. The first bladed propulsor is coupled with the first electric motor (112) within the first housing. The propulsion system (20) further includes a second propulsor module having a second housing, a second electric motor, and a second bladed propulsor. The second bladed propulsor is coupled with the second electric motor within the second housing. The propulsion system (20) also includes a first generator (82) driven by the at least one engine (22) and electrically connected with the first electric motor (112). Additionally, the propulsion system (20) includes a second generator driven by the at least one engine (22) and electrically connected with the second electric motor.

An assembly is provided for an aircraft. This aircraft assembly includes a fluid reservoir structure, an aircraft engine housing structure and a mounting assembly (20). The fluid reservoir structure includes a first mount (28). The aircraft engine housing structure includes a second mount (30). The mounting assembly (20) connects the first mount (28) to the second mount (30) while providing controlled intercomponent movement between the first mount (28) and the second mount (30) to reduce internal stress and/or wear. The mounting assembly (20) includes a first bushing (32), a second bushing (34) and a fastener (36).

A propulsion system (20) for an aircraft includes a propulsion rotor, an engine (22), a propulsor module (24), and a blended exhaust system. The engine (220) has an axis (48), is mechanically coupled to the propulsion rotor, and includes an exhaust section (38). The propulsor module (24) includes a housing forming a propulsor duct with a duct inlet and duct outlet, and a propulsor assembly (94) with an electric motor (112) coupled to a bladed propulsor disposed within the propulsor duct. The blended exhaust system includes an exhaust duct assembly having an engine exhaust inlet at the exhaust section, a propulsor exhaust inlet at the duct outlet, and a blended exhaust outlet. The exhaust duct assembly is configured to receive an engine exhaust stream and a propulsor exhaust stream, form a blended exhaust stream including both, and direct the blended exhaust stream out of the blended exhaust system through the blended exhaust outlet.

An assembly of a gas turbine engine, having: a shaft (50) having a shaft outer surface (120A) and a shaft inner surface (120B), the shaft (50) axially extends from a shaft forward end (50A) to a shaft aft end (50B), the shaft forward end (50A) defines axially extending shaft slots (130) that are circumferentially spaced apart from each other; a bearing (110) having a first race (150), the first race (150) is ring shaped with a first race outer surface (160) and a first race inner surface (170), the first race (150) is sized so the first race inner surface (170) fits around the shaft outer surface (120A), the first race (150) extends axially from a first race forward end (180) to a first race aft end (190), the first race forward end (180) defines radially inwardly extending bosses (200) that are circumferentially spaced apart from each other and aligned with ones of the axially extending slots (130) of the shaft forward end (50A), rotationally fixing the first race (150) against the shaft (50).

A method of manufacturing a component of a housing defining a rotor cavity (20), the method includes: obtaining a body (214C,218F) of the component of the housing, the body (214C,218F) having a first face (211A,218A) and a second face opposite to the first face (211A,218A) , the second face facing away from the rotor cavity (20), the body (214C,218F) made of a first material; performing a treatment to the body (214C,218F) to increase a wear-resistance of the first material; depositing a bonding layer (214D,218H) on the second face of the body (214C,218F), the bonding layer (214D,218H) including a second material being dissimilar from the first material, the first material being more wear-resistant than the second material, the second material defining an exposed face of the bonding layer (214D,218H) facing away from the body (214C,218F); and completing a shape of the body (214C,218F) by bonding a quantity of a third material to the exposed face of the bonding layer (214D,218H).

An aircraft engine has a high pressure spool including a high pressure turbine drivingly connected to a high pressure compressor. A low pressure spool including a low pressure compressor is fluidly connected to the high pressure compressor. A low pressure turbine is drivingly connected to the low pressure compressor to drive the low pressure compressor. A load is drivingly connected to the low pressure turbine, the load consisting of one of a propeller and a helicopter rotor. A method of creating classes of an aircraft engine from an engine platform is disclosed.