Pratt & Whitney Canada Patente

A modular floor stand (20) for supporting an aircraft engine (30) includes a first cart (40) having a first base (41) equipped with a first set of wheels (42), and first engine support members (43) extending upwardly from the first base (41). The modular floor stand (20) further includes a second cart (50) having a second base (51) equipped with a second set of wheels (52), and second engine support members (53) extending upwardly from the second base (51), the second cart (50) sized and shaped to selectively engage the first cart (40) in a nested arrangement. A linking bar (60) has a first end (60a) detachably connectable to the first cart (40) and a second end (60b) detachably connectable to the second cart (50). The linking bar (60) is configured to rigidly retain the first cart (40) and the second cart (50) in a fixed opposed facing relationship.

A system for selecting an active engine and a standby engine in an aircraft having at least two engines (102, 104) includes an engine selection controller (110) configured to select a first engine of the at least two engines (102, 104) to operate in a standby mode and a second engine of the at least two engines (102, 104) to operate in an active mode responsive to at least one engine selection parameter and generate engine mode selection data indicating the first engine to operate in standby mode. The engine selection controller (110) is further configured to output the engine mode selection data to an Idle Cruise Regimen (ICR) controller (304). A memory (106; 302) is configured to store the engine mode selection data indicating the selected first engine to operate in the standby mode from the engine selection controller (110) and sensor data from the at least two engines (102, 104).

A system for controlling operation of engines (102, 104) of a multi engine aircraft between standby mode and active mode comprising an idle cruise regimen controller (304) configured to actuate entry of a selected engine (102, 104) into the standby mode of operation in response to receipt of an indication of the selected engine (102, 104) as a standby engine. The idle cruise regimen controller (304) is further configured to decrease a torque of the selected engine (102, 104) until the torque of the selected engine (102, 104) reaches a low setting, reduce a gas generator speed of the selected engine (102, 104) to a first target speed responsive to the torque of the selected engine (102, 104) reaching the low setting, maintain the gas generator speed of the selected engine (102, 104) at the first target speed for a predetermined cool down period and reduce the gas generator speed of the selected engine (102, 104) from the first target speed to a second target speed after expiration of the predetermined cool down period.

A system for controlling operation of engines of a multi engine aircraft between standby mode and active mode includes an idle cruise regimen controller configured to actuate abnormal exit of a selected engine from the standby mode of operation. The idle cruise regimen controller is further configured to increase a maximum gas generator speed limit (NHMAX) of the selected engine from a standby mode level (604) to a predetermined exit limit (612), increase a speed of the selected engine to a re-engagement speed responsive to increasing of the maximum gas generator speed limit (NHMAX) and maintain a maximum torque limit (QMAX) of the selected engine at a maximum torque level (600) during increase of a gas generator speed (NH).

An apparatus for protecting signals from electromagnetic interference comprising a first conducting wire (302) configured to carry the signals from a first end of the first conducting wire (302) to a second end of the first conducting wire (302). A shield (306) made of a conducting material surrounds the first conducting wire (302) from the first end of the first conducting wire (302) to the second end of the first conducting wire (302). A second conducting wire (300) is associated with the shield (306). A ground connection connects the second conducting wire (300) to ground at a single location.

A system for controlling operation of engines of a multi engine aircraft includes an idle cruise regimen controller configured to actuate exit of a selected engine from a standby mode of operation. The idle cruise regimen controller is further configured to increase a maximum gas generator speed limit (NHMAX) of the selected engine from a standby mode level (604) to a predetermined exit limit (612), decrease a maximum torque limit (QMAX) of the selected engine to 20% of a maximum torque level (614), maintain the maximum torque limit (QMAX) of the selected engine at the 20% of the maximum torque level (614) for a predetermined warm up period, increase a gas generator speed (NH) of the selected engine to a first gas generator speed level (616), increase the maximum torque limit (QMAX) to a maximum torque level (600) and increase a speed of the selected engine to a re-engagement speed responsive to increasing of the maximum torque limit (QMAX).

A gearbox de-oiler comprises a housing (102) mounted to a shaft (104) for rotation therewith. The housing (102) defines a chamber (102e) containing a porous separator matrix (114). The chamber (102e) has an air-oil mixture inlet (106), an oil outlet (108) and an air outlet (110). A separation plate (120) separates the porous separator matrix (114) into a primary flow path segment (F1) and a secondary flow path segment (F2). The primary flow path segment (F1) extends axially from and fluidly connects the air-oil mixture inlet (106) to a radial passage. The secondary flow path segment (F2) extends axially from and fluidly connects the radial passage to the air outlet (110), ) axially overlaps the primary flow path segment (F1) and is disposed radially inwardly with respect thereto. Oil collector tubes (140) projects radially outwardly from the separation plate (120) across at least a portion of the primary flow path segment (F1) for fluidly connecting the secondary flow path segment (F2) to the oil outlet (108).

A method is provided for operating an aircraft engine (10) in a fire-induced emergency state. The engine (10) is housed in a nacelle (41) defining a nacelle bypass duct (43) having a first outlet (45a) fluidly connected to an inlet (11) of the engine (10) and a second outlet (45b) fluidly connected to an environment external to the nacelle (41). An indication of a fire of the aircraft engine (10) is received. In response to receiving the indication, an emergency shutdown sequence of the engine (10) is initiated. In response to the initiating the emergency shutdown sequence, a volume of air entering the inlet (11) of the aircraft engine (10) is increased by restricting air flow through the second outlet (45b) of the nacelle bypass duct (43). The volume of air entering the inlet (11) is used to drive a rotor (14) of the engine (10). The rotor (14) is used to drive an oil pump (32) to circulate oil through a main oil conduit (35) of the engine (10).

A rotating seal assembly (24) of a gas turbine engine (14) includes a rotating seal element (34), and a rotationally stationary seal support (36; 64) positioned radially outboard of the rotating seal element (34). The rotating seal element (34) and the seal support (36; 64) are configured to define a seal interface therebetween. The seal element (34) includes an axial pump portion (54) defined by one or more axially and circumferentially extending grooves (56) to urge an airflow (50) therethrough. The airflow (50) is configured to prevent a fluid from entering the seal interface. The axial pump portion (54) defines a radial running clearance (62) with a floating seal (60).

The air oil separator (130) can have a Venturi conduit (134) having, in sequential fluid flow communication along a fluid path, an inlet, a converging section (140), a throat section (142), a diverging section (144) and an air outlet, a motive flow nozzle (132) coaxial to the Venturi conduit (134) and directed to the inlet, a secondary flow passage between the motive flow nozzle (132) and the inlet, and an annular groove (160) defined around the fluid path, between the throat section (142) and the diverging section (144), the annular groove (160) fluidly connecting an oil outlet.

A method (300) for manufacturing a component of an aircraft engine (10) includes positioning two sub-components of the component in a fixture of a CNC machine, and performing, with a 3D scanner, a first scan of the two sub-components on the fixture. A predetermined tool pathway for combining the two sub-components is updated based on the first scan to establish a modified tool pathway for combining the two sub-components. A welding procedure is performed to combine the two sub-components based on the modified tool pathway to form the component. Using the 3D scanner, and with the component retained in place in the fixture, a second scan of the component is performed to identify the presence of one or more surface defects on the component. In response to identifying the presence of the surface defect(s), a corrective procedure is performed to correct the one or more surface defects on the component.

A method (300) includes obtaining, from a first sensor (211), a first pressure, the first pressure exerted through a volume of oil (207) in an oil tank (205) of an aircraft upon the first sensor (211), the first sensor (211) disposed near a sensor port (209) proximate to a bottom of the oil tank (205). The method includes obtaining, from a second sensor (215), a second pressure, the second pressure comprising an atmospheric pressure of the oil tank (205) or of an environment external to the oil tank (205), determining a pressure differential between the first pressure and the second pressure, and determining an oil level of the oil tank (205) based on the pressure differential and an expected oil density.

A turboshaft gearbox output gear arrangement (100) of a gas turbine engine, having: a power turbine shaft (110); a torque shaft (120), driven by the power turbine shaft (110), extending from a forward end (120F) to an aft end (120A); a gear shaft (122), driven by the torque shaft (120); a mating gear (125), driven by the gear shaft (122), wherein the mating gear (125) and gear shaft (122) define a gear mesh plane (127); a reference tube (130), coupled to the torque shaft (120), extending from a forward end (130F) to an aft end (130A), wherein the forward and aft ends of the reference tube (130) and torque shaft (120) extend along each side of the gear mesh plane (127); and a sensor (140D) that engages the reference tube (130) and the torque shaft (120) to measure twist in the torque shaft (120), wherein angular movement of the reference tube (130) relative to the torque shaft (120) is indicative of torque reacted in the torque shaft (120).

A motorized gear system (100) of a gas turbine engine (20), having: a reduction gear (23) including: an input shaft (131), driven by an engine output shaft (161); a first branch (140) having a first shaft (141) and a first gearset (151), driven by the input shaft (131); a second branch (150) having a second shaft (151) and a second gearset (152), driven by the input shaft (131); an output shaft (161), driven by the first branch (140) and the second branch (150); and wherein the motorized gear system (100) includes a first motor (120A), coupled to the first branch (140).

An aircraft engine, has: a low-pressure assembly having first components having first inlets and first scavenge outlets; a high-pressure spool having second components having second inlets and second scavenge outlets; a lubricant system for lubricating the first components and the second components, the lubricant system having: a lubricant tank containing a volume of lubricant; a main pump fluidly connected to the lubricant tank and to the components; and a scavenge pump fluidly connected to the scavenge outlets and to the lubricant tank; and an electric motor drivingly engageable to the main pump, the lubricant system having: a standard configuration in which the electric motor is powered off and in which the main pump is drivingly engaged by a shaft of the aircraft engine; and a windmilling configuration in which the main pump is drivingly engaged by the electric motor.

A rotary internal combustion engine, has: a housing (20) having walls (24) defining a rotor cavity (21), the walls (24) having inner faces (24l) exposed to the rotor cavity (21) and outer faces (24O) opposed to the inner faces (24l), the walls (24) defining passages (24P) between the inner faces (24l) and the outer faces (24O), a wall of the walls (24) defining a pump-receiving cavity (24C) between an outer face of the outer faces (24O) and an inner face of the inner faces (24I) that corresponds to that outer face; a rotor (12) received within the rotor cavity (21) and drivingly engaged to an output shaft (13); and a pump system (103) received into the pump-receiving cavity (24C) and including a first inlet (103A) connectable to a fluid source (101), a first outlet (103B) and a second inlet (103C) both connected to the passages (24P), and a second outlet (103D) connectable to the fluid source (101), the pump system (103) operable to generate a fluid flow between the inlets (103A,103C) and the outlets (103B,103D).

An aircraft engine, has: a low-pressure assembly having first components (C1) having first inlets (I1) and first scavenge outlets (S1), one of the first components having a lubricant-receiving volume (V1); a high-pressure spool having second components (C2) having second inlets (I2) and second scavenge outlets (S2); a lubricant system (400A) having: a lubricant tank (301); a main pump (302) and a scavenge pump (303) driving drive a flow of the lubricant; a secondary pump (310) connectable to the volume (V1) and an outlet connectable to the first components (C1); valves (312,313); the lubricant system (400A) having: a standard configuration in which the valves (312,313) define a standard flow path extending from the lubricant tank (301) to the first inlets (I1) and the second inlets (I2) via the main pump (302); and a windmilling configuration in which the valves (312,313) define a windmilling flow path extending from the lubricant-receiving volume (V1) to the first inlets (I1) via the secondary pump (310) while bypassing the main pump (302), and from the first scavenge outlets (S1) to the lubricant-receiving volume (V1).

A rotating seal assembly (24) of a gas turbine engine (14) includes a rotating seal element (34) and a rotationally stationary seal support (36) positioned radially outboard of the rotating seal element (34). The rotating seal element (34) and the seal support (36) are configured to define a seal interface therebetween. The seal element (34) includes an axial pump portion (54) defined by one or more axially and circumferentially extending grooves (56) to urge an airflow (50) therethrough. The airflow (50) is configured to prevent a fluid from entering the seal interface.

An assembly for an aircraft propulsion system (20) includes a battery (64), an electrical distribution system (66), a battery cell voltage sensor (132A), and a controller (28, 108). The battery (64) includes a plurality of battery cells (76). The electrical distribution system (66) is operable to electrically interconnect the battery (64) with a charging power source (62, 86). The battery cell voltage sensor (132A) is operable to measure a cell voltage of a first battery cell of the plurality of battery cells (76). The controller (28, 108) is connected in signal communication with the battery cell voltage sensor (132A) and the HVPDU (88). The controller (28, 108) configured to execute a charging sequence for the battery (64), determine a change in cell voltage of the first battery cell (76), determine a test cell resistance of the first battery cell (76), determine a change in resistance of the first battery cell (76), and identify a presence or an absence of an end-of-life condition of the first battery cell (76).

A hybrid-electric propulsion (HEP) system (20) for powering an aircraft is provided. The HEP system (20) includes an inlet case (28), an engine output shaft (48), a gearbox (26), a gas turbine engine (22), an electric motor (24; 24A), and a gear system (54; 54A, 54B). The inlet case (28) forms a cavity (58) of the HEP system (20). The engine output shaft (48) is disposed within the cavity (58) and extends axially along a first rotational axis (62). The gearbox (26) is in drive communication with the engine output shaft (48), and the gas turbine engine (22) in drive communication with the engine output shaft (48). The electric motor (24; 24A) is in drive communication with the engine output shaft (48). At least a portion of the electric motor (24; 24A; 24B) is disposed within the cavity (58). The gear system (54; 54A, 54B) is configured to rotatably engage the electric motor (24; 24A) with the engine output shaft (48).

Impeller assemblies (300) for gas turbine engines include a rotating component (302) with a plurality of impeller blades (304) extending from a hub (306) and defining an axis (334). A split shroud (308) is arranged about the rotating component (302) and defines a flow path between an external surface of the hub (306) and an internal surface of the split shroud (308). The impeller blades (304) are configured to be rotationally driven relative to the split shroud (308). The split shroud (308) includes a first shroud element and a second shroud element configured to fixedly attach to the first shroud element. Each of the first shroud element and the second shroud element have a radially inward tapering portion (330) that extends from a shroud leading edge (324) of the respective shroud element to an inflection point (326) and a radially outward tapering portion (332) that extends from the inflection point (326) to a shroud aft edge (328).

A linear guide system for use in the assembly or disassembly of a gas turbine engine (20) is provided. The gas turbine engine (20) includes an engine shaft (34) and an engine casing (29). The engine shaft is engaged with a compressor section and a turbine section and extends along an axial centerline (32). The engine casing has first and second engine casing segments (29A, 29B) that are axially separable from one another. The linear guide system includes a plurality of rail and guide block assemblies (104). Each rail and guide block assembly includes a rail head (106), a rail (108), and a guide block (110). The rail head is configured for attachment to the first engine casing (29A). The guide block is configured for attachment to the second engine casing (29B), and is configured to receive the rail in a manner that permits linear traverse of the rail relative to the guide block.

An aircraft propulsion system includes a thermal engine, a gearbox (36), and a multiplate clutch. The gearbox (36) includes an input shaft (46), an output shaft (48), an epicyclic gear arrangement (42), a drive gear (44), and a first layshaft assembly (50). The multiplate clutch is engaged with the first layshaft assembly (50). The epicyclic gear arrangement includes a sun gear (54), a plurality of planet gears (56), and a ring gear (60). The first layshaft assembly (50) includes a first layshaft, a first layshaft input gear (70), and a first layshaft ring engagement gear (72). The first layshaft ring engagement gear (72) is engaged with the ring gear (60), and the first layshaft input gear (70) is engaged with the drive gear (44) and the input shaft (46). The multiplate clutch is disposable in a disengaged configuration and in an engaged configuration.

A method of monitoring a fuel nozzle (16) of an aircraft engine (10) includes supplying fuel to a combustor (22) of the aircraft engine (10) using the fuel nozzle (16) during operation of the aircraft engine (10), and sensing an operating condition of the aircraft engine (10) applicable to a duration during which the fuel nozzle (16) is supplying fuel to the combustor (22). A machine learning algorithm is executed to associate the operating condition of the aircraft engine (10) and the duration with a health degradation of the fuel nozzle (16). A servicing status for the fuel nozzle (16) is annunciated based on the health degradation of the fuel nozzle (16).

An aircraft engine has: a low-pressure assembly having first components (C1) having first inlets (I1) and first scavenge outlets (S1); a high-pressure spool having second components (C2) having second inlets (I2) and second scavenge outlets (S2); a lubricant system (300A) having: a lubricant tank (301); a main pump (302) and a scavenge pump (303) operable to drive a flow of the lubricant; a secondary pump (310); and valves (312,313); the lubricant system (300A) having: a standard configuration in which the valves (312,313) define a standard flow path extending from the lubricant tank (301) to the first inlets (I1) and the second inlets (I2) via the main pump (302); and a windmilling configuration in which the valves (312,313) define a windmilling flow path extending from a lubricant source to the first inlets (I1) via the secondary pump (310) while bypassing the main pump (302), and from the first scavenge outlets (S1) to the lubricant source while bypassing the scavenge pump (303).

A gas turbine engine (20) is provided that includes a compressor section (24), a combustor (26), a turbine section (28), an engine shaft (34), a shaft retention system (54), and an axial centerline (32). The engine shaft (34) is engaged with the compressor section (24) and the turbine section (28). The engine shaft (34) has an aft end (34A) and axially extends along the axial centerline (32) of the engine (20). The shaft retention system (54) includes a shaft nut (150) and a shaft retainer (154). The shaft nut (150) is attached to the aft end (34A) of the engine shaft (34). The shaft retainer (154) is mounted for axial translation normal to the axial centerline (32). The shaft retainer (154) is disposable in an engaged configuration wherein the shaft retainer (154) is coupled with the shaft nut (150), and in a disengaged configuration wherein the shaft retainer (154) is disengaged with the shaft nut (150).

An aircraft engine (10) includes an annular gas path extending across a compressor (14) and a turbine (18) with a casing (26), a shroud (32) and blades (28), a clearance (C) defined between blade tips and the shroud (32). An active tip clearance control device (36) directing a cooling airflow (F1, F2) to the casing (26) includes a first air source (40) at a first temperature, a second air source (42) at a second temperature greater than the first, a valve (46) having first and second inlets connected to the sources (40, 42) and an outlet selectively connectible with the inlets. The outlet is fluidly connected to the casing (26) to selectively direct the flow of cooling air (F1, F2) from the first or second source. A sensor (60) measures the clearance (C). A controller (48) connected to the sensor (60) receives a signal therefrom of a measured clearance (C), and generates commands causing the valve (46) to open to the first or second source (40, 42) when the clearance (C) is above or below a predetermined value.

A method may include receiving (402) a generator power demand (GPD) signal for a gas generator of a gas turbine engine (GTE). The method may also include applying (404) a smoothing filter to the GPD signal to create a smoothed GPD signal. The method may also include determining (406), using the smoothed GPD signal, a GPD anticipation parameter for controlling fuel flow rate to the GTE. The method may also include outputting (420) a fuel flow rate request to the GTE based on the GPD anticipation parameter and a power turbine speed (NPT) controlled by an NPT governing feedback loop.

A pump (200) for an aircraft engine (10), has: a housing (210) including a static wall (214), a peripheral wall (215) extending around an axis (A1), and a movable wall (216) movable along the axis (A1) in sealing contact with the peripheral wall (215), the static wall (214), the peripheral wall (215) and the movable wall (216) defining therebetween a rotor cavity (221) communicating with an inlet (211) and an outlet (212) of the housing (210); an input shaft (219) extending through the housing (210); a rotor (220) drivingly engaged by the input shaft (219) and movably contained within the rotor cavity (221), the rotor (220) having two sections axially telescopically connected to one another and both extending between two apexes (225); an actuator (244) operatively connected to the movable wall (216) and operable to displace the movable wall (216) relative to the static wall (214); and a biasing member (240) operatively connected to the two sections of the rotor (220) to bias the two sections away from one another to follow a movement of the movable wall (216).

A method for washing an aircraft engine (10) including a compressor (14) having variable guide vanes (141) pivotable about respective spanwise axes (S0) and having a secondary air system (100) in fluid communication with the compressor (14) via a bleed outlet (23) located downstream of the variable guide vanes (141), the method includes: injecting a washing fluid into the compressor (14) of the aircraft engine (10); while the compressor (14) is rotating at a rotational speed, pivoting the variable guide vanes (141) about their respective spanwise axes (S0) to a purging position at which an air pressure at the bleed outlet (23) is maximal for the rotational speed of the compressor (14); and after the injecting of the washing fluid, expelling the washing fluid out of the secondary air system (100) by injecting air in the bleed outlet (23) with the compressor (14) being in rotation while the variable guide vanes (141) are in the purging position.

A propulsion system (20) includes an engine (22), a drivetrain (26), and a propulsor (24). The engine (22) includes an engine output shaft (52). The drivetrain (26) is coupled with the engine output shaft (52). The drivetrain (26) includes a propulsor output shaft (60), a gear assembly (64), and an electric motor assembly (62). The gear assembly (64) includes a propulsor output gear (70) mounted on the propulsor output shaft (60). The propulsor output gear (70) includes a ring (84). The ring (84) includes an outer gear face (86) and an inner gear face (88). The electric motor assembly (62) includes at least one electric motor (94). The at least one electric motor (94) includes a rotor (96). The rotor (96) includes an electric motor output shaft (98) and an electric motor output gear (100) mounted on the electric motor output shaft (98). The electric motor output gear (100) is engaged with the inner gear face (88). The propulsor (24) is interconnected with the propulsor output gear (70) by the propulsor output shaft (60).

A method may include determining, via a load sharing management operation (212) of a controller, a target steady state load share for an associated one of a thermal engine (108) or an electric motor system (110) of a hybrid electric powertrain (HEP). The method may also include operating, by the controller, the associated one of the thermal engine (108) or the electric motor system (110) according to the target steady state load share. The method may also include determining, by the controller, that a measured load for the associated one of the thermal engine (108) or the electric motor system (110) does not meet the target steady state load share. The method may also include adjusting, by the controller, a control setting of the associated one of the thermal engine (108) or the electric motor system (110), resulting in an adjusted load output of the associated one of the thermal engine (108) or the electric motor system (110).

An aircraft propulsion system is provided that includes a thermal engine, a gearbox (36), and a hydraulically actuated clutch system. The gearbox (36) includes an input shaft (46), an output shaft (48), an epicyclic gear arrangement (42), a drive gear (44), and a first layshaft assembly (50). The clutch system includes a hydraulic coupling device engaged with the first layshaft assembly (50). The epicyclic gear arrangement (42) includes a sun gear (54), planet gears (56), and a ring gear (60). The first layshaft assembly (50) includes a first layshaft, a first layshaft input gear (70), and a first layshaft ring engagement gear (72). The first layshaft ring engagement gear (72) is engaged with the ring gear (60), and the first layshaft input gear (70) is engaged with the drive gear (44) and the input shaft (46). The clutch system is disposable in an disengaged configuration and in an engaged configuration.

A propulsion system (20) includes an engine (22), a drivetrain (26), and a propulsor (24). The engine (22) includes an engine output shaft (52). The drivetrain (26) is coupled with the engine output shaft (52). The drivetrain (26) includes a propulsor output shaft (60), a gear assembly (64), and an electric motor assembly (62). The gear assembly (64) includes a propulsor output gear (70) mounted on the propulsor output shaft (60). The electric motor assembly (62) includes a bevel gear assembly (92) and an electric motor (90). The bevel gear assembly (92) is rotatable about a second rotational axis (112). The bevel gear assembly (92) is engaged with the propulsor output gear (70). The electric motor (90) includes a rotor (96) rotatable about a third rotational axis (104). The third rotational axis (104) is oriented at an oblique angle relative to the second rotational axis (112). The rotor (96) is engaged with the bevel gear assembly (92). The propulsor (24) is interconnected with the propulsor output gear (70).

A hybrid aircraft power plant (10) includes an air mover (24) for propelling an aircraft (12), a gas turbine engine (14) operable to drive the air mover (24), and an electric machine (46) operable as a motor to drive the air mover (24). The gas turbine engine (14) includes a turbine disk (30) for extracting energy from a stream of combustion gas, and an exhaust duct (38) defining: an exhaust passage for conveying the stream of combustion gas from the turbine disk (30); and a cavity (41) disposed radially inward of the exhaust passage. A shaft (32,50) drivingly connects the turbine disk (30) with the electric machine (46) and extends inside the cavity (41) defined by the exhaust duct (38). A fan (51) is drivingly connected to the shaft (32,50) and operable to circulate cooling air inside the cavity (41) of the exhaust duct (38).

A powerplant assembly (24) includes an engine (32), a turbocompressor (34), and an overspeed protection valve (74). The engine (32) includes an air intake (64) and an engine exhaust (68). The turbocompressor (34) includes a compressor section (46), a turbine section (48), and a rotational assembly (50). The rotational assembly (50) includes a turbocompressor shaft (54), a bladed compressor rotor (56) of the compressor section (46), and a turbine rotor (58) of the turbine section (48). The engine (32) and the turbocompressor (34) form a gas flow path (60). The gas flow path (60) extends from the compressor section (46) to the air intake (64), through the engine (32), and from the engine exhaust (68) to the turbine section (48). The overspeed protection valve (74) forms a portion of the gas flow path (60). The overspeed protection valve (74) is positionable in an open state and a closed state.

An aircraft propulsion system assembly includes a propulsion system (20) and a control assembly (26). The propulsion system (20) includes a propulsor (28), an engine (22), and an electrical assembly (24). The engine (22) includes a rotational assembly (44) coupled to the propulsor (28). The electrical assembly (24) includes an electric motor (62) coupled to the propulsor (28). The control assembly (26) includes a propulsor sensor assembly (68), an electric motor sensor (83), channel A control units (70), and a channel B control unit (72). The channel A control units (70) includes an engine control unit (86) for the engine (22) and an electric motor control unit (88) for the electric motor (62). The channel B control unit (72) is connected in signal communication with the propulsor sensor assembly (68), the electric motor sensor (83), the engine control unit (86), and the electric motor control unit (88).

A gas turbine engine (20) is provided that includes a compressor section (22), a combustor (24), and a mixed-flow turbine (26). The mixed-flow turbine (26) has a first turbine ("FT") rotor stage (32), an annular first vane stage (34), and a second turbine ("ST") rotor stage (36). The FT rotor stage (32) includes an FT disk (32A) and a plurality of FT rotor blades (32B) extending outwardly from the FT disk (32a). The ST rotor stage (36) includes an ST disk (36A) and a plurality of ST rotor blades extending outwardly from the ST disk (36A). The annular first vane stage (34) includes a plurality of stator vanes (42) that are distributed circumferentially, and disposed axially between the FT rotor stage (32) and the ST rotor stage (36). One or more of the stator vanes (42) includes a radially extending interior passage configured to permit a flow of cooling air to pass through the respective one or more stator vanes (42).

A gas turbine engine (26) is provided that includes a combustor (38), a rotational spool (50), a spool balancing system, and at least one access port (74). The rotational spool (48) includes turbine and compressor sections, and a shaft that connects the turbine and compressor sections. The spool balancing system includes a circumferentially extending track component (68) and a plurality of balance weights (70) that are engaged with the track component (68). The balance weights (70) are disposable in secured and unsecured configurations. The track component (68) and the balance weights (70) are disposed to rotate with the rotational spool (50). The access port (74) is disposed within a non-rotational member of the gas turbine engine (26). The access port (74) is configured to receive a tool for engagement with the balance weights (70).

A method includes obtaining at a first sampling time, a first set of n sampled values of n control parameters for one or more control systems of an aircraft. The method includes performing, in a higher power mode, a first sequential lookup and interpolation operation of a table maintained in the memory (110) to obtain a first output control value, and sending, the first output control value to the one or more control systems of the aircraft. Determining that at least one condition associated with a first steady state has been satisfied and switching to a first lower power mode of the device (100). In the first lower power mode, performing a second sequential lookup and interpolation operation to obtain a second output control value and sending the second output control value to the one or more control systems of the aircraft.