Pratt & Whitney Canada Patente

A method of operating a multi-engine system of an helicopter includes the multi-engine system having a first turboshaft engine having a first shaft, a second turboshaft engine having a second shaft, a gearbox having a clutch system (152), and a range of rotation speeds defined as a placarded zone. The method includes: rotating the first and second shafts at a first idle rotation speed (NL<LI>) below the placarded zone when clutched to a load; increasing a rotation speed of the first shaft from the first idle rotation speed (NL<LI>) to a flight rotation speed (NL<F>) above the placarded zone; unclutching the second shaft from the load during the increasing; and increasing a rotation speed of the second shaft to a second idle rotation speed (NL<HI>) when the second shaft is unclutched from the load, the second idle rotation speed (NL<HI>) above the placarded zone and below the flight rotation speed (NL<F>).

A turbofan engine (10) has a bypass passage for channelling a bypass flow and a core passage for channelling a core flow around a central axis (30). An exhaust mixer arrangement (32; 32') comprises a mixer body (33) mounted for rotation about the central axis (30) and having an annular wall extending around the central axis (30). The annular wall defines a plurality of circumferentially distributed alternating inner and outer lobes (42, 44), with each inner lobe (42) protruding into the core passage, and each outer lobe (44) protruding into the annular bypass passage (24). A driving unit (50) is operatively connected to the mixer body (33) for selectively driving the mixer body (33) in rotation about the central axis (30). A controller (62) is operatively connected to the driving unit (50) for controlling a rotational speed of the mixer body (33) as a function of a flight operating condition.

A rotor stage (26A) for a gas turbine engine (20) is provided that includes first and second rotor blades (32A 32B), a disk (34), and a seal member (46). The rotor blades (32A, 32B) each have an airfoil (42), an attachment section (36), a neck section (38), and a platform (40). The platform (40) extends laterally outward from the neck section (38). The platform (40) has platform inner and outer radial surfaces (40C, 40D). The disk (34) has forward and aft axial end surfaces (62, 64) and disk slots (44). A first rotor blade (32A) includes a first rib (72) extending outwardly from a platform inner radial surface (40C). In the assembled state, the blade attachment sections (36) are received in the disk slots (44). In the assembled state, the rotor blade platforms (40) and an outer radial surface of the disk (34) form a tab slot (70). In the assembled state, the first rib (72) locates the seal member (46) relative to platform lateral edge surfaces (40A, 40B).

A torque measurement system (201) is provided and includes a first shaft (210) to rotate by input torque, a second shaft (220) to rotate by torque received from the first shaft, the second shaft including first lugs (221), a reference shaft (230) including second lugs (231) and being connected to the first and second shafts via spline and helical spline connections (232, 233), respectively, an elastic element (240) to oppose a reaction force of the helical spline connection to axially displace the reference shaft resulting in reference shaft twisting due to the helical spline connection and a measurement system (250) to measure angular displacement between the first and second lugs due to the reference shaft twisting.

A rotor stage (26A) for a gas turbine engine (20) is provided that includes first and second rotor blades (32A, 32B), a disk (34), and a seal member (46). The rotor blades (32A, 32B) each have an airfoil (42), an attachment section (36), a neck section (38), and a platform (40). The disk (34) has forward and aft axial end surfaces (62, 64), an outer radial surface that extends between the forward and aft axial end surfaces (63, 64), and disk slots (44) configured to receive rotor blade attachment sections (36). In an assembled state, the attachment sections (36) are fully received in the disk slots (44), and adjacent rotor blade platforms (40), and the outer radial surface of the disk (34) form a tab slot (70). In the assembled state, a central segment (46B) of the seal member (46) is engaged with a platform channel (52) disposed at a platform lateral edge surface (40A), and an axial locator tab (46A, 46E) of the seal member (46) is disposed in the at least one tab slot (70).

A rotor stage (26A) for a gas turbine engine (20) is provided that includes first and second rotor blades (32A, 32B), a disk (34), and a seal member (46). The rotor blades (32A, 32B) each have an airfoil (42), an attachment section (36), a neck section (38), and a platform (40). The platform (40) has platform inner and outer radial surfaces (40C, 40D). The disk (34) has first and second disk slots (44) that are adjacent one another. In an assembled state, the attachment sections (36) are received in the slots (44). In the assembled state, the platforms (40) of the first and second rotor blades (32A, 32B) and an outer radial surface of the disk (34) form at least one slot (68). The seal member (46) includes a central segment (46B) and an axial locator tab (46A, 46E). In the assembled state, the central segment (46B) is disposed for engagement with the first and second rotor blade platforms (40), and the axial locator tab (46A, 46E) is disposed in the slot (68).

Gas turbine engines (20) include a rotating component (304) having a disc and defining an axis therethrough. A static component (302) is arranged axially adjacent to the rotating component (304) and includes a primary rub flange (320) that extends axially from the disc of the rotating component (304) toward the static component (302). The static component (302) includes an interfacing rub flange (324) extending axially from the static component (302) toward the rotating component (304) and is arranged radially adjacent to the primary rub flange (320). The primary rub flange (320) includes a primary rubbing surface (322) and the interfacing rub flange (324) includes an interfacing rubbing surface (326) with the primary rubbing surface (322) spaced from the interfacing rubbing surface (326) by a separation gap (328). The primary rub flange (320) is configured to deflect in the presence of a crack in the disc to close the separation gap (328) such that the primary rubbing surface (322) contacts the interfacing rubbing surface (326).

A propulsion system (20) for an aircraft (1000) includes a battery, an electrical distribution system (66), a ground-based charger (86), and an engine controller (28). The battery includes a plurality of battery strings each including a plurality of battery cells (72). The electrical distribution system includes a battery string switch assembly (98) operable to electrically interconnect each of the plurality of battery strings together in parallel. The ground-based charger is electrically connected to the electrical distribution system. The ground-based charger includes a charger controller (132) including a simple electronic hardware (SEH) system. The engine controller is connected in signal communication with the SEH system. The engine controller is configured to determine a battery charging profile (138) specific to the battery and charge the battery by controlling the ground-based charger, through the SEH system, to supply electrical power to the electrical distribution system at a set voltage and a set current defined by the battery charging profile.

A propulsion system (20) for an aircraft (1000) includes a battery (64), an electrical distribution system (66), a ground-based charger, and a battery management system (68). The battery includes a plurality of battery strings (72) each including a plurality of battery cells (76). The electrical distribution system includes a battery string switch assembly (98) operable to electrically interconnect each of the plurality of battery strings together in parallel. The ground-based charger includes a charger controller (132) including a simple electronic hardware (SEH) system. The battery management system includes a battery management system (BMS) controller (108) connected in signal communication with the SEH system. The BMS controller is configured to determine a battery charging profile specific to the battery and charge the battery by controlling the ground-based charger, through the SEH system, to supply electrical power to the electrical distribution system at a set voltage and a set current defined by the battery charging profile.

A propulsion system (20) for an aircraft includes a battery, an electrical distribution system (66), and a ground-based charger (86). The battery (64) includes a plurality of battery strings (72). Each of the plurality of battery strings (72) includes a plurality of battery cells (76). The electrical distribution system (66) includes a battery string switch assembly (98). The battery string switch assembly (98) is operable to electrically interconnect each of the plurality of battery strings (72) together in parallel. The ground-based charger (86) is electrically connected to the electrical distribution system (66). The ground-based charger (86) includes a charger controller (132). The charger controller (132) is configured to determine a battery charging profile (138) specific to the battery (64) and charge the battery (64) by supplying electrical power to the electrical distribution system (66) at a set voltage and a set current defined by the battery charging profile (138).

A propulsion system (20) includes a propulsor (26), an engine (22), an electrical assembly (24), and a protection circuit. The engine (22) is coupled with the propulsor (26). The electrical assembly (24) includes a battery (66), a motor control unit (64), an electric motor (62), and an electrical distribution system (68). The electrical distribution system (68) is configured to electrically interconnect the battery (66) and the motor control unit (64). The electrical distribution system (68) includes a contactor between the battery (66) and the motor control unit (64). The motor control unit (64) is electrically connected with the electric motor (62). The electric motor (62) is coupled with the propulsor (26). The protection circuit includes a current sensor configured to measure an electrical current between the battery (66) and the motor control unit (64). The protection circuit is operable to identify a reverse current flow using the measured electrical current and transmit a protection open control signal to the contactor to control the contactor to switch to an open state.

A propulsion system (20) includes a propulsor (26), an engine (22), an electrical assembly (24), and an emergency energy protection assembly. The engine (22) includes a fuel system (38). The fuel system (38) includes a fuel control unit. The electrical assembly (24) includes a battery (66), a motor control unit (64), an electric motor (62), and an electrical distribution system (68). The electrical distribution system (68) includes at least one contactor electrically connected between the battery (66) and the motor control unit (64). The at least one contactor is switchable between a closed state and an open state. The engine (22) and the electric motor (62) are coupled with the propulsor (26). The emergency energy protection assembly includes an emergency actuator. The emergency actuator is actuable in an actuated state to control the at least one contactor to switch to the open state, control the motor control unit (64) to deenergize the electric motor (62), and control the fuel control unit to stop fuel flow for the engine (22).

An assembly for a propulsion system (20) of an aircraft (1000) includes a gearbox module (32), at least one accessory load assembly (24), a propulsor (34), and an electric motor (38). The gearbox module (32) includes a gear assembly (88) and an output shaft (90). The gear assembly (88) is connected to the output shaft (90). The at least one accessory load assembly (24) includes at least one accessory load (124) coupled to the output shaft (90). The propulsor (34) is coupled to the output shaft (90). The electric motor (38) includes a rotor (42). The rotor (42) is coupled to the gear assembly (88) to drive rotation of the output shaft (90) about a rotational axis (104). The rotation of the output shaft (90) drives rotation of the propulsor (34) and the at least one accessory load (124) for each accessory load assembly (24) of the at least one accessory load assembly (24).

A propulsion system (20) for an aircraft (1000) includes a battery (64) and a battery sensor assembly (92). The battery (64) includes a plurality of battery strings (74). A first battery string of the plurality of battery strings (74) includes a plurality of substring assemblies (94). Each of the substring assemblies (94) includes a battery substring (96), a first substring contactor (98), and a first substring controller (102). The battery substring (96) includes a plurality of battery cells (106) electrically connected together. The first substring contactor (98) is electrically connected with the battery substring (96). The first substring contactor (98) is positionable in a closed state or an open state. The first substring controller (102) is configured to control a position of the first substring contactor (98) in the closed state or the open state. The battery sensor assembly (92) includes a cell temperature sensor (92A) for each of the plurality of battery cells (106). The cell temperature sensor (92A) is connected in signal communication with the first substring controller (102).

An apparatus for a gas turbine engine (20), such as an apparatus for a turbine exhaust section (49). The apparatus includes an airfoil (70) extending spanwise along a span line (72) from a base (74) to a tip (76). The airfoil (70) extends laterally between a pressure side (78) and a suction side (80), and the airfoil (70) extends along a chord line (82) from a leading edge (84) to a trailing edge (86). The airfoil (70) includes a slot (88) extending through the airfoil (70) from a slot inlet (94) to a slot outlet (96). The slot inlet (94) is disposed on the pressure side (78) and the slot outlet (96) is disposed on the suction side (80) The slot (88) defines a channel (98) for directing fluid flow from the pressure side (78) to the suction side (80) through the airfoil (70), such that the slot (88) will bypass the associated pressure losses from a blunt facing body.

An exhaust system (15) for an aircraft engine, has: a turbine exhaust duct, TED, (30) having an annular inlet conduit (33) extending around a central axis (17) for directing combustion gases generally in an axial direction, and outlet conduits (34; 35) fluidly communicating with the annular inlet conduit (33) and extending generally radially outward relative to the annular inlet conduit (33); and an exhaust case (40) surrounding the TED (30), the exhaust case (40) having: a frustoconical section (42) extending from a fore end (42A) at an intersection with a remainder of the exhaust case (40) to an aft end (42B) securable to a turbine case (18C) of the aircraft engine, the frustoconical section (42) converging towards the central axis (17) from the fore end (42A) to the aft end (42B); and a connecting flange (42C) at the aft end (42B), the connecting flange (42C) protruding inwardly towards the central axis (17).

A gas turbine engine (22) includes a turbine section (36), a turbine case (70), an inner case (72), and a damping seal baffle (74). The turbine section (36) extends along a rotational axis (50) of the gas turbine engine (22). The turbine case (70) includes a cantilevered inner wall (86). The cantilevered inner wall (86) extends circumferentially about the rotational axis (50). The cantilevered inner wall (86) extends between and to an upstream axial end (88) and a downstream axial end (90). The upstream axial end (88) is disposed at and downstream of the turbine section (36). The cantilevered inner wall (86) forms a seal platform surface (122). The inner case (72) extends circumferentially about the rotational axis (50). The inner case (72) is disposed radially inward of the upstream axial end (88). The damping seal baffle (74) extends circumferentially about the rotational axis (50). The damping seal baffle (74) is mounted to the inner case (72). The damping seal baffle (74) is disposed in contact with the seal platform surface (122).

A combustor heat shield (14) includes a wall (40, 42) bound by an inner surface (50) and an outer surface (52) spaced from the inner surface (50). The inner surface (50) bounds at least a portion of a combustion chamber (46). Effusion holes (68A, 68B) extend through the wall (40, 42) from the outer surface (52) to the inner surface (50) within a first zone (62) of the wall (40, 42) and a second zone (64) of the wall (40, 42) discrete from the first zone (62). The diameters (D1) of effusion holes (68A) within the first zone (62) are larger than diameters (D2) of effusion holes (68B) within the second zone (64).

An aircraft engine (10), has: an upstream stator (30) having upstream stator vanes (31) circumferentially distributed about a central axis (11); and a downstream stator (40) having downstream stator vanes (41) circumferentially distributed about the central axis (11), the downstream stator (40) located downstream of the upstream stator (30) relative to an airflow flowing within a core gaspath (24) of the aircraft engine (10), a number of the upstream stator vanes (31) being different than a number of the downstream stator vanes (41), major portions of leading edges (41A) of the downstream stator vanes (41) circumferentially overlapped by the upstream stator vanes (31), the downstream stator vanes (41) including: a first pair (44) of circumferentially adjacent vanes of the downstream stator vanes (41) spaced apart by a first pitch (P1), and a second pair (45) of circumferentially adjacent vanes of the downstream stator vanes (41) spaced apart by a second pitch (P2) different than the first pitch (P1).

A gas turbine engine includes a turbine section (36), a turbine case, an inner case (72), and a damping seal baffle (74). The turbine section (36) extends along a rotational axis (50) of the gas turbine engine. The turbine case (70) includes a cantilevered inner wall (86). The cantilevered inner wall (86) extends circumferentially about the rotational axis (50). The cantilevered inner wall (86) extends between and to an upstream axial end (88) and a downstream axial end. The upstream axial end (88) is disposed at and downstream of the turbine section (36). The cantilevered inner wall (86) forms a first sealing surface (108) and a second sealing surface (112). The inner case (72) and the damping seal baffle (74) extend circumferentially about the rotational axis (50). The damping seal baffle (74) is mounted to the inner case (72). The damping seal baffle (74) is disposed in contact with one or both of the first sealing surface (108) or the second sealing surface (112).

A de-aerator (54) includes a de-aerator housing (68) having a fluid inlet (54a) through which a mixture of oil and air enters, an air-oil outlet (54b) through which mixture of oil and air of lower oil concentration exits, and an oil outlet (54c) through which a flow of oil exits. An impeller (66) is positioned in the de-aerator housing (68) and is driven about an impeller axis (70) to separate oil from the mixture of air and oil and direct the oil toward the oil outlet (54c), and a seal (74) is positioned in the de-aerator housing (68) downstream of the impeller (66). The seal (74) includes a plurality of seal fins (76) extending radially outwardly from an impeller hub (78) of the impeller (66) toward an inner housing surface (80) of the de-aerator housing (68). The plurality of seal fins (76) define a labyrinth seal (82) with the inner housing surface (80) to minimize an oil outlet (54c) oil flow from exiting the de-aerator housing (68) via the air-oil outlet (54b).

An exhaust system (15), has: a turbine exhaust duct, TED, (30) having an annular inlet conduit (33) extending around a central axis (17), and outlet conduits (34; 35) fluidly communicating with the annular inlet conduit (33) and extending generally radially outward relative to the annular inlet conduit (33); and an exhaust case (18D) surrounding the TED (30), the exhaust case (18D) having openings (18E), each of the outlet conduits (34; 35) received through a respective one of the openings (18E); and exhaust conduits (40) secured to the exhaust case (18D) and each extending around a respective one of the openings (18E) of the exhaust case (18D), an exhaust conduit (40) of the exhaust conduits (40) including: a duct (41) protruding away from the exhaust case (18D) and fluidly communicating with a respective one of the outlet conduits (34; 35); a flange (42) flaring away from the duct (41), the exhaust conduit (40) secured to the exhaust case (18D) via the flange (42); and a stiffener (43) protruding transversally from the flange (42), the flange (42) located between the duct (41) and the stiffener.

A gear pump arrangement (201) of an aircraft gas turbine engine is provided. The gear pump arrangement (201) includes an accessory (210), a quill shaft (220), two gear pumps (240, 250) in series, axial retention members (271, 272) at the accessory (210) and one of the two gear pumps (250) to axially retain the quill shaft (220) and torque transmitting members (TTMs). The TTMs include a first TTM (281) between the one of the two gear pumps (250) and the quill shaft (220) and a second TTM (282) between the quill shaft (220) and the accessory (210) whereby the quill shaft (220) directly links the one of the two gear pumps (250) to the accessory (210).

A system (100) for measuring a gear backlash, having a connecting member (140) with: a lower portion (150), defining an engaging plate (160) with an aperture (165), positioned against a front end (120A) of a bevel gear (120); and an upper portion (170) for installing the engaging plate (160); a gauge body (190) positioned against an accessory pad (200) of a gearbox housing (130); a shaft system (210) extending through the gauge body (190), into the gearbox housing (130), having aft (220) and front (230) ends, and a sidewall (222): the aft end (220) extends aft of the gauge body (190) and rotates the shaft system (210); the front end (230) has: a first portion (240) that engages the plate aperture (165) when the shaft system (210) is rotated, biasing the engaging plate (160) against the bevel gear front end (120A); a second portion (250) engages a bevel gear shaft (120B) when the engaging plate (160) is biased; and a sensor (260), connected to the system aft end (220), that measures rotation of the shaft system (210) relative to the gauge body (190).

A method for managing mismatches in feedback from guide vane sensors (161a, 161b) of a turbine engine (151) with variable guide vanes (157) includes selectively muting individual feedback channels (165a, 165b) and obtaining time series turbine performance data of the turbine engine (151) associated with individual feedback channels (165a, 165b). The feedback channel (165a, 165b) of the guide vane sensor (161a, 161b) associated with the best performing time series performance data can be selected as sole feedback channel (165a, 165b) following detection of a mismatch condition.

An epicyclic gear train (12) of an aircraft powerplant includes a sun gear (32) having a rotation axis (RA), a ring gear (34), a plurality of intermediate gears (36) circumscribed by the ring gear (34) and meshed with the sun gear (32) and the ring gear (34), a carrier (38) supporting the intermediate gears (36), and a torque frame (40) attaching the carrier (38) to a stationary structure (42) of the aircraft powerplant to prevent rotation of the carrier (38) and of the intermediate gears (36) about the rotation axis (RA) of the sun gear (32). The torque frame (40) is attached to the carrier (38) at a connection. The torque frame (40) and the carrier (38) cooperatively define a lubricant passage (78, 82, 84) extending through the connection and supplying lubricant to at least one of the intermediate gears (36).

A pump assembly (10) of an aircraft engine includes a pump housing (36), a first pump stage (12) positioned in the housing (36) and including two first stage rotating elements (20, 22) meshed at a first mesh location (24), and a second pump stage (14) positioned in the housing (36) and including two second stage rotating elements (26, 28) meshed at a second mesh location (30). The first stage rotating elements (20, 22) are retained at respective shafts (16, 18) by one or more woodruff keys (44), and at least one second stage rotating element (26, 28) of the two second stage rotating elements (26, 28) is secured at its respective shaft (16, 18) via a spline connection (70) of the at least one second stage rotating element (26, 28) to the respective shaft (16, 18).

A turbine exhaust duct, TED, (30) for an aircraft engine, has: an annular inlet conduit (33) extending around a central axis (17) for directing combustion gases generally in an axial direction; outlet conduits (34; 35) communicating with the annular inlet conduit (33) and extending generally radially outward relative to the annular inlet conduit (33), the outlet conduits (34; 35) extending from inlet ends (34A; 35A) at intersections with the annular inlet conduit (33) to outlet ends (34B; 35B), an outlet conduit of the outlet conduits (34; 35) having: a forward section (34F; 35F) facing an axially forward direction and a rearward section (35R; 35R) opposite the forward section (34F; 35F), the forward section (34F; 35F) and the rearward section (35R; 35R) conjointly defining an outlet end of the outlet ends (34B; 35B); and a reinforcement plate (40) secured to the forward section (34F; 35F), the reinforcement plate (40) having a thickness (T1) greater than a baseline thickness (T0) of the outlet conduit outside the reinforcement plate (40).

Gas turbine engines include an engine core having compressor, combustor, and turbine sections arranged within an engine casing. A transfer tube (404) is arranged on an exterior of the casing and configured to define a fluid conduit for fluid (410) to be transferred about the engine. A fitting (402) is connected to the transfer tube (404) and a seal assembly (406) is arranged between the fitting (402) and the transfer tube (404). The seal assembly (406) provides a fluid seal between the fluid conduit and an ambient environment (412) at a connection between the fitting (402) and the transfer tube (404). The seal assembly (406) includes a primary seal (414) arranged in a gap (413) between the fitting (402) and the transfer tube (404) and a secondary seal (416) arranged in the gap (413) between the primary seal (414) and an ambient environment (412). The secondary seal (416) is configured to expand into engagement with the fitting (402) and the transfer tube (404) when exposed to temperatures of at least 800 °F.

A method and a tool (18) for performing a dimensional inspection on a vane ring (10), comprises referencing an inspection tool (18) relative to a vane (16) of the vane ring (10) to aim a measuring tool, such as a caliper (C), at predetermined spanwise and chordwise coordinates along the vane (16). Once the inspection has been properly positioned on the vane (16), the measuring tool (18) is used to measure a dimensional parameter (H), such as a vane thickness, at a location corresponding to the predetermined spanwise and chordwise coordinates. Then, the operator determines if the measured dimensional parameter is within acceptable tolerance limits.

An exhaust system (15) for an aircraft engine of an aircraft having a direction of travel, has: a turbine exhaust duct, TED, (30) having an annular inlet conduit (33) extending around a central axis (17) for directing combustion gases generally in an axial direction and outlet conduits (34, 35) communicating with the annular inlet conduit (33) and extending generally radially outward relative to the annular inlet conduit (33); and an exhaust case (18D) enclosing the TED (30), the exhaust case (18D) extending from a forward end (18F) to a rearward end (18R) relative to the direction of travel, the exhaust case (18D) having openings (18E), each of the outlet conduits (34, 35) received through a respective one of the openings (18E); and reinforcing members (40) secured to the exhaust cases between the forward end (18F) and the rearward end (18R), the reinforcing members (40) being located opposite circumferential sides of the openings (18E) and locally increasing a thickness of the exhaust case.

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

A propulsion system (20) for an aircraft (1000) includes a battery (64), at least one charger contactor (106, 132, 142), a first control channel (110, 114), and a second control channel (112, 116). The battery (64) includes a plurality of battery strings (72). The at least one charger contactor (106, 132, 142) includes a first gate (138, 148) and a second gate (140, 150). The at least one charger contactor (106, 132, 142) is operable to a closed position with a first control signal present at the first gate (138, 148) and a second control signal present at the second gate (140, 150). The at least one charger contactor (106, 132, 142) is operable to an open position with one or both of the first control signal absent at the first gate (138, 148) or the second control signal absent at the second gate (140, 150). The first control channel (110, 114) is operable to selectively apply the first control signal to the first gate (138, 148). The second control channel (112, 116) is operable to selectively apply the second control signal to the second gate (140, 150).

A propulsion system (20) for an aircraft (1000) includes a battery (64), an electrical distribution system (66), a charger, and a battery management system (68). The battery includes a plurality of battery strings including a first battery string. Each of the plurality of battery strings includes a plurality of battery cells. The battery management system includes a battery sensor assembly and a battery management system (BMS) controller. The battery sensor assembly includes a plurality of battery cell voltage sensors. The BMS controller is configured to control the charger to charge the battery, including at least the first battery string, by executing a battery charging profile including a target charging voltage, measure a cell voltage of each of the plurality of battery cells of the first battery string, and identify correlation or non-correlation of a charging voltage applied by the charger with the target charging voltage.

A pressure regulating valve (22) is provided that includes a housing (24), a sleeve (30), a piston (32), an adjustment nut (34), and a piston spring (36). The housing (24) includes a sleeve bore (26). The sleeve (30) has a side wall (42) that extends axially between a base end (44) and a distal end (46), an end wall (48) disposed at the base end (44), an interior cavity (50) that extends axially between the base end (44) and the distal end (46), and a plurality of ports (28). The sleeve (30) is disposed within the sleeve bore (26). The piston (32) extends axially between a land end (56) and a spring end (58). The adjustment nut (34) is engaged with the sleeve (30), and is configured for axial translation relative to the sleeve (30). The piston spring (36) is disposed to provide a spring force on the adjustment nut (34) and the piston (32). The sleeve (30) includes an asymmetric feature (68, 68A, 68C, 68D) that engages with the housing (24) to prevent rotation of the sleeve (30) relative to the housing (24).

A valve assembly (20) for an aircraft engine (10) includes a base (21) defining an air opening (19) configured to receive airflow therethrough in a flow direction from an upstream side (15) to a downstream side (17), a stopper (24) extending from the downstream side (17) of the base (21) adjacent the air opening (19), and a flapper plate (22) pivotably mounted to the base (21) about a pivot axis (25). The flapper plate (22) is pivotable relative to the base (21) and the stopper (24) between a closed position (22B) and a fully open position (22A). In the fully open position (22A), a downstream surface (28) of the flapper plate (22) abuts the stopper (24). In the closed position (22B), the flapper plate (22) obstructs the air opening (19). A deflector (41) extends from an upstream surface (26) of the flapper plate (22) and projects into the airflow when the flapper plate (22) is located in the fully open position.

A heat management system (100,200) for a thermal engine (10) includes a source of hydrogen (104,204), a first hydrogen conduit (102,202) flowing a first flow of hydrogen therethrough, the first hydrogen conduit (102,202) fluidly coupling the source of hydrogen (104,204) to a fuel system (106,206) of the thermal engine (10), and a second hydrogen conduit (110,216) flowing a second flow of hydrogen therethrough. The second hydrogen conduit (110,216) is in thermal communication with an engine exhaust conduit (118,220) to transfer heat from the engine exhaust conduit (118,220) to the second hydrogen conduit (110,216). The second hydrogen conduit (110,216) is also in thermal communication with the first hydrogen conduit (102,202) to transfer heat from the second hydrogen conduit (110,216) to the first hydrogen conduit (102,202).

A method of operating a hybrid electric powertrain (HEP) (100) may include determining, based on a HEP torque demand and a maximum torque limit for a gas turbine engine (108) having a mechanical output coupled to a power shaft (104) of the HEP (100), a target torque for the gas turbine engine (108). The method may also include controlling the gas turbine engine (108) to operate according to a target speed or the target torque. The method may also include, while operating the gas turbine engine (108) according to the target speed or the target torque, determining, based on the HEP torque demand and the target torque, a total torque compensation demand for an electric machine (110) having a mechanical output controllably couplable to the power shaft (104) of the HEP (100), and operating the electric machine (110) according to the total torque compensation demand.

An integrally bladed rotor (IBR) (201) includes a hub (210) and blades (220), each blade (220) includes an airfoil section (230) extending from a blade root (266) at the hub (210) to a blade tip defining a radially outboard extent of the rotor blade (220) to aerodynamically interact with a flow of air for compressing the air. A root fillet (240) is integrally formed with the hub (210) the airfoil section (230) integrally extends therefrom. The blade (220) includes a localized increased blade root thickness portion (256) defined based on at least one of an axial thickness (250) of a web portion of the hub (210) and a chord length (258) of the blade (220).

An aircraft system (20) includes a first propulsor rotor (28) rotatable about an axis (38), and a second propulsor rotor (30) rotatable about the axis (38). An electric machine (32) includes a first electric machine rotor (42), a second electric machine rotor (44) and an electric machine stator (46) radially between the first electric machine rotor (42) and the second electric machine rotor (44). The first electric machine rotor (42) is rotatable about the axis (38) and rotatably connected to the first propulsor rotor (28). The first electric machine rotor (42) and the electric machine stator (46) form a first motor (72) configured to drive rotation of the first propulsor rotor (28). The second electric machine rotor (44) is rotatable about the axis (38) and rotatably connected to the second propulsor rotor (30). The second electric machine rotor (44) and the electric machine stator (46) form a second motor (74) configured to drive rotation of the second propulsor rotor (30).