Pratt & Whitney Canada Patente

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).

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 aircraft engine (10), has: a turbine (15); a scroll case (20) connected to a source of combustion gases (19) and to the turbine (15), and a conduit (21) extending around a central axis (A); a bearing housing (30) including a support flange (41); a turbine support case (40) secured to 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 ribs (41A,60) secured to the support flange (41) and distributed around the central axis (A), a rib (60) of the ribs (41A,60) circumferentially and axially overlapping the distal end (44B) of the spoke (44), the rib (60) having a rib face (61) in radially opposed facing relationship to a spoke face (44D) of the spoke (44), a projection of the rib face (61) in a circumferential direction relative to the central axis (A) intersecting the spoke face (44D).

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).

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 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.

A method of evaluating a battery cell (44) within a battery module (42) of an aircraft hybrid-electric propulsion (HEP) system (20) is provided. The HEP system (20) includes a thermal engine (22), an electric motor (24), and a propulsion unit (28). The HEP system (20) is configured to control at least one of the thermal engine (22) or the electric motor (24) to be in drive communication with propulsion unit (28). The method includes: controlling the HEP system (20) to have the battery module (42) and the battery cell (44) disposed in a rest mode; determining an open-circuit voltage (OCV) of the battery cell (44) after a settling period of time; determining an internal resistance value of the battery cell (44) using the determined OCV; and evaluating the battery cell (44) based on the determined OCV and the determined internal resistance value.

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 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 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).

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 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.

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 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 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).

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).

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).

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).

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.

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).

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 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).

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').

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 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 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.

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 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).