Abstract: An airfoil assembly extends along a radial direction between a root and a tip, the airfoil assembly comprising: a first blade segment positioned proximate the root of the airfoil assembly; a second blade segment positioned adjacent the first blade segment along the radial direction; and a tensioning assembly comprising a plurality of tensioning strings that extend between and mechanically couple the first blade segment and the second blade segment.

Abstract: A turbine blade includes a blade body, a division plate, a plurality of first protruding ridges, a plurality of second protruding ridges, a first jet hole group, and a second jet hole group. The blade body includes a suction-side wall surface and a pressure-side wall surface, which are provided with multiple film cooling holes. The stagnation region at the leading edge of the blade body is provided with multiple film cooling holes. The first and second jet hole groups are provided on the division plate. The first protruding ridges are provided on a suction-side inner wall surface, and the second ridges are provided on a pressure-side inner wall surface. An engine including the turbine blade is further provided.

Abstract: A gas turbine engine comprises a turbine, a combustor fluidly coupled to the turbine, and a cooling air system. The turbine includes including a turbine rotor having a shaft mounted for rotation about an axis of the gas turbine engine and a set of turbine blades coupled to the turbine rotor for rotation therewith. The combustor includes an outer combustor case and an inner combustor case that cooperate to define a combustion chamber. The cooling air system is configured to cool the turbine using air form the combustion chamber of the combustor.

Abstract: A stator vane assembly of a gas turbine engine includes a stator shroud having a plurality of shroud openings formed therein. The stator shroud defines a flowpath surface. The assembly includes a plurality of stator vanes, one stator vane of the plurality of stator vanes installed in each shroud opening of the stator shroud. A volume of potting material is located in the plurality of shroud openings to retain the stator vanes thereat. One or more retainers are located at a back surface of the stator shroud opposite the flowpath surface to restrict circumferential movement of the plurality of stator vanes relative to the stator shroud.

Abstract: A turbine vane in a gas turbine engine includes an inner platform, an outer platform, and a vane airfoil positioned therebetween. The vane airfoil includes a first cooling passage extending between the outer platform and the inner platform, and a second cooling passage extending between the outer platform and the inner platform. The second cooling passage is arranged downstream of the first cooling passage with respect to a flow direction. The turbine vane includes a jumper tube disposed between the second cooling passage and the inner platform. The jumper tube includes an inlet, an outlet, and a tube wall enclosing a hollow interior. The inlet is positioned a distance within the second cooling passage. The outlet is positioned at least partially through an aperture of the inner platform.

Abstract: Tip shrouds or other shrouds with multi-slope geometries are generally implemented using segments, since performing the necessary cut for an alternative split-ring design is difficult given conventional cutting processes. However, a segmented design generally results in more leakage, relative to a split-ring design. Accordingly, a split-ring multi-slope design is disclosed that can be more easily manufactured. In particular, a continuous ring may be cut along a linear path to produce a split ring, and then the ends of the split ring may be machined to form complementary shiplap portions. The split ring may then be compressed for installation by overlapping the shiplap portions, to form a seal against leakage through the shroud.

Abstract: A boundary-layer turbomachine coupled to a shaft for transmitting power, comprising a plurality of ducts, and a plurality of blades and/or one or more protrusions. The plurality of ducts are defined by duct walls configured to rotate about the longitudinal axis and are concentrically arranged thereabout to convey fluid between inlet and outlet ends. Flow inlets draw the fluid into the plurality of ducts at least partially azimuthally around the longitudinal axis towards the outlet end. The plurality of blades and/or one or more protrusions extend radially in the duct between opposing duct walls. The one or more protrusions may spirally extend at least partially along and around the longitudinal axis to induct fluid into the duct. Slots may be provided in duct walls for centrifugal separation of liquid phase. Systems and methods of generating power using a plurality of boundary-layer turbines, including without condensers, pumps, and/or compressors.

Abstract: A variable geometry turbine includes: a turbine rotor; a scroll passage forming part which forms a scroll passage; an exhaust gas passage forming part which forms an exhaust gas passage for introducing an exhaust gas from the scroll passage to the turbine rotor; and a variable nozzle unit including a plurality of nozzle vanes disposed in the exhaust gas passage and configured to be rotatable about respective rotation centers. The exhaust gas passage forming part includes: a first plate member having an annular first plate part; and a second plate member having an annular second plate part which defines the exhaust gas passage between the first plate part and the second plate part and is disposed closer to a turbine outlet than the first plate part in an axial direction of the turbine rotor.

Abstract: Hybrid composite components, such as gas turbine engine containment assemblies, a hybrid composite component having an annular composite shell and an annular metallic shell joined with the composite shell. A containment assembly including a containment case extending along an axial direction about a longitudinal centerline of the gas turbine engine. The containment case has an inner surface and an outer surface spaced apart along a radial direction and includes a first composite shell joined with a metallic shell. The metallic shell defining a first portion of the inner surface and the first composite shell defining a second portion of the inner surface.

Abstract: A turbine system includes a foam generating assembly having an in situ foam generating device at least partially positioned within the fluid passageway of the turbine engine, such that the in situ foam generating device is configured to generate foam within the fluid passageway of the turbine engine.

Abstract: A gas turbine engine includes a plurality of rotating components housed within a compressor section and a turbine section. A first tap is connected to the compressor section and configured to deliver air at a first pressure. A heat exchanger is connected downstream of the first tap. A flowpath is defined between a rotating surface and a non-rotating surface. The flowpath is connected downstream of the heat exchanger and is configured to deliver air to at least one of the plurality of rotating components. At least a portion of the non-rotating surface and the rotating surface includes a base metal. An insulation material is disposed on a surface along the flowpath.

Abstract: A method and system for aligning a component within a turbine casing, and a related turbine casing, are disclosed. In a top-on position, a location of a reference point and another, vertically spaced reference point on the lower casing are measured. After removing at least the upper casing, the reference points' locations are measured again, and the locations of a reference point on an upper surface of the HJ flange are measured. A prediction offset value is calculated for the component support position in the top-on position based on the locations. The prediction offset value may include a vertical adjustment based, in part, on a translation of a triangular spatial relationship of a number of the reference points and/or a tilt angle, a horizontal adjustment, and a HJ flange surface distortion adjustment. The component support position is adjusted by the prediction offset value to improve alignment.

Abstract: The disclosure concerns a gland condenser skid system for thermodynamic machine, namely a steam turbine, the gland condenser skid system comprising a shell and plates heat exchanger as gland condenser (10), said shell and plates heat exchanger being formed of gasket-free welded tube sheets.

Abstract: A variable valve device can change valve operations of a cylinder head in which a cam chain is disposed at a center in a predetermined direction where cylinders are arranged and spark plugs are arranged by recessing outer walls on both sides in the direction in a concave shape. The device includes, for each cylinder, a rocker shaft extending along the direction, rocker arms supported by the rocker shaft, a connecting pin disposed in a pin hole of one of the rocker arms, a return pin disposed in a pin hole of the other of the rocker arms, a pressing member that causes the connecting pin to push the return pin, and a repulsive member that causes the return pin to push back the connecting pin, and the pins and the members are separated from the plug in a direction orthogonal to the direction in a plan view.

Abstract: A variable valve device includes a pair of cam housings separated in a predetermined direction, a rocker shaft supported by the pair of cam housings, rocker arms supported by the rocker shaft, a connecting pin disposed in a pin hole of one of the rocker arms, a return pin disposed in a pin hole of another of the rocker arms, a pressing member that causes the connecting pin to push the return pin to the other side, a repulsive member that causes the return pin to push back the connecting pin to one side, and an upper housing supported at both ends by upper surfaces of the pair of cam housings. In the variable valve device, the upper housing is formed with a first accommodation hole in which the pressing member is disposed and a second accommodation hole in which the repulsive member is disposed.

Abstract: A phasing system for an internal combustion engine having a concentric camshaft includes an annular stator rotatable by the crankshaft and having inner and outer circumferences, and two groups of arcuate cavities. The first group interrupts the inner circumference. A first phaser having an output member as a hub is mounted within the stator in contact with the inner circumference and supported directly on the camshaft tube as a concentric stator bearing. The first phaser has vanes connected to the output member and extending radially into the first group to divide cavities into opposed working chambers. A second phaser comprises first and second end plates at opposite stator sides and fastened to one another to axially seal the cavities therein and, as second output member of the second phaser, connect to vanes extending axially through the second group of cavities to divide the cavities into opposed working chambers.

Abstract: A shifting stopper latch assembly for a rocker arm can comprise a latch assembly housing, comprising a latch assembly oil feed, a piston bore, a stopper bore comprising a stopper opening, and a spring seat. A piston can be in the piston bore. A shifting stopper can be in the stopper bore and the shifting stopper can comprise a projection protruding out of the stopper opening. A return spring can be configured in the spring seat to bias the shifting stopper and the plunger away from the spring seat and towards the latch assembly oil feed. A valvetrain can comprise a rocker arm configured to actuate against a cam on a cam rail, and the rocker arm can comprise a latch surface configured to selectively engage and disengage the projection.

Abstract: A lubrication system for an internal combustion engine includes a fluid flow control device at an outlet side of the pump that regulates pressure conditions at the outlet side of the pump upstream of the lubrication circuit in the engine. The fluid flow control device is located in a recirculation path, and opens in response to a fluid pressure at the outlet of the pump exceeding a first threshold to allow the fluid to pass from the outlet side of the pump back to an inlet side of the pump. The recirculation path includes a tuning device that maintains the fluid pressure at engine speeds above the first threshold. In response to engine speeds above a second threshold, the tuning device restricts fluid flow in the recirculation path so that the fluid pressure in the fluid flow path increases.

Abstract: The present disclosure relates to a detection system for detecting a failure in a fluid path of a fluid system, the detection system including: at least one pipe comprising a pipe wall which defines a fluid path for the transfer of fluid and at least one electrically conductive element extending along the at least one pipe and which forms an electrical signal path for a first electrical signal which indicates a state of the at least one pipe; and, an electrical terminal electrically connected to the at least one electrically conductive element so that the first electrical signal is transmitted to the electrical terminal via the electrical signal path.

Abstract: A bilayler metamaterial silencer allows substantial fluid through the apparatus, while mitigating the propagation of sound through the apparatus, and while providing a form factor that is significantly more compact than previously-known devices. Moreover, illustrative embodiments allow a designer to specify one or both of the frequency or frequencies at which the apparatus mitigates sound propagation, and/or the bandwidth around the frequency or frequencies at which the apparatus mitigates sound propagation.

Abstract: A system includes a controller for an exhaust aftertreatment system including a SCR catalyst in exhaust gas-receiving communication with an engine and at least one reductant dosing system structured to provide reductant to the exhaust gas. The controller is structured to determine a ratio of NO to NO2 at or proximate an inlet of the SCR catalyst. The controller is further structured to command the at least one reductant dosing system to increase, decrease, or maintain an amount of reductant provided to the exhaust gas based on comparing the ratio of NO to NO2 to a previous NO to NO2 ratio.

Abstract: Systems and methods for diagnosing at least one component in an exhaust aftertreatment system are provided. The system includes an exhaust aftertreatment system coupled to an engine system, at least one sensor, and at least one processing circuit structured to: receive initial sensor data; determine an initial parameter value based on the initial sensor data; determine that the initial parameter does not satisfy an initial threshold; perform operations to diagnose at least one component of the exhaust aftertreatment system comprising: causing the engine system to operate through a sequence of a plurality of engine outputs; receiving a plurality of sensor data, each of the plurality sensor data corresponding to at least one of the plurality of engine outputs; comparing each of the plurality of sensor data to a corresponding threshold; and diagnosing the at least one component based on the comparison.

Abstract: A removal determination device that accurately determines a removal of an exhaust gas purification device from an exhaust pipe. The removal determination device comprises: an input temperature sensor detecting a temperature at an upstream side of casing; an output temperature sensor detecting a temperature at a downstream side of the casing; an engine stop time collector measuring a time period from an inactivation of an engine to a startup of the engine; and a determiner determining a presence of the exhaust gas purifying device based on the input temperature and the output temperature, in a case that the time period is equal to or longer than a predetermined time.

Abstract: An exhaust gas treatment system comprising a first and a second structure being in fluid communication and defining a flow path of the exhaust gases, the second structure being supported by the first structure, the first structure comprising a first catalytic converter downstream a mixing chamber, the first catalytic converter having a fluid connection to an outlet of exhaust gases, the second structure including a filter unit housing removably connected to the first structure and to an inlet module by a filter unit housing fixing, the second structure defining a longitudinal axis, the filter unit housing including a filter unit, and the inlet module including a second catalytic converter and having a fluid connection to an inlet of exhaust gases, the filter unit housing and/or the inlet module being configured to allow its removal from the exhaust gas treatment system.

Abstract: A support member includes: a pair of first support members provided to sandwich and support an attachment portion, one of the first support members being in contact with a vehicle side; and a second support member provided between the pair of first support members, formed separately from the first support members, and configured to support an inner peripheral side of a through hole. An attachment portion is provided at the other end portion of a plate spring member having one end portion fixed to a catalyst converter, and vibration applied to the catalyst converter is absorbed by elastic deformation of the plate spring member. Vibration in a thickness direction of the plate spring member is absorbed by elastic deformation of the first support members, and vibration in a longitudinal direction of the plate spring member is absorbed by elastic deformation of the second support member.

Abstract: An electric outboard motor includes a screw, an electric motor, an outboard motor case, and a cooling portion. The outboard motor case accommodates the electric motor. The cooling portion cools a heat generation portion at an inside or an outside of the outboard motor case by a cooling liquid. At least a lower region of the outboard motor case that is submerged in outside water is constituted of a multiple case structure body that includes an outer case and an inner case. In the multiple case structure body of the lower region, at least part of a space portion that is surrounded by the outer case and the inner case is a cooling liquid passage through which the cooling liquid supplied to the cooling portion flows.

Abstract: The present invention relates to a reservoir tank for coolant that stores the coolant for cooling an internal combustion engine. The reservoir tank includes: a tank chamber that stores the coolant; an opening, which is disposed in an upper part of the tank chamber, through which the coolant flows in and flows out, and in a circumferential wall of which a breathing hole is formed; and a first communication passage, a second communication passage, and a third communication passage that respectively communicate with the opening and the tank chamber.

Abstract: An HHO gas second fuel is produced in a pressure-resistant container and distributed at a low volumetric rate at multiple locations about the internal combustion engine.

Abstract: There are disclosed systems and methods for generating electrical power from oxy-fuel combustion in a turbine system. The turbine system makes use of recycled steam as a components of the turbine working fluid. Also disclosed is an integrated recuperator and separator, which may be used with the turbine system, configured to separate water and carbon dioxide from exhaust fluids from the turbine system, heat from the exhaust fluids being used to generate steam from the separated water for recycling to the turbine system. Carbon dioxide separated from the exhaust fluids is condensed to a liquid or supercritical phase and sequestered in a subsurface natural gas reservoir from which natural gas fuel for the turbine system is extracted.

Abstract: An assembly is provided for an aircraft propulsion system. This assembly includes a nacelle inlet structure and a nozzle. The nacelle inlet structure extends circumferentially about an axial centerline. The nacelle inlet structure includes an inlet lip, a bulkhead and an internal cavity formed by and axially between the inlet lip and the bulkhead. The nozzle is configured to direct gas into the internal cavity axially towards the bulkhead.

Abstract: A thermal management system for an aircraft includes a first gas turbine engine, first thermal bus, first heat exchanger, one or more first ancillary systems, vapour compression system, one or more second ancillary systems and second heat exchanger. A waste heat energy generated by a first gas turbine engine, and a first ancillary system, transfers to the first heat transfer fluid. A waste heat energy generated by a second ancillary system transfers to a second heat transfer fluid, and the second heat exchanger transfers the waste heat energy from the second heat transfer fluid to the first heat transfer fluid. The waste heat energy generated by a second ancillary system transfers to the first heat transfer fluid, and the first heat exchanger transfers the waste heat energy to a dissipation medium. The waste heat energy transferred to the second heat transfer fluid ranges from 20 kW to 300 kW.

Abstract: A thermal management system for an aircraft comprises a first gas turbine engine, one or more first electric machines rotatably coupled to the first gas turbine engine, a first thermal bus, and a first heat exchanger module. The first thermal bus comprises a first heat transfer fluid, with the first heat transfer fluid being in fluid communication, in a closed loop flow sequence, between the first gas turbine engine, the or each first electric machine, and the first heat exchanger. Waste heat energy generated by at least one of the first gas turbine engine, and the or each first electric machine, is transferred to the first heat transfer fluid. The first heat exchanger module comprises a first flow path and a second flow path.

Abstract: A thermal management system for an aircraft includes a first gas turbine engine, one or more first electric machines rotatably coupled to the first gas turbine engine, a first thermal bus, and a first heat exchanger. Waste heat energy generated by at least one first gas turbine engine, and first electric machine, transfers to the first heat transfer fluid. The first heat exchanger directs a first proportion of the first heat transfer fluid through a first heat dissipation portion wherein a first proportion of the waste heat energy transfers to a first dissipation medium dependent on the first dissipation medium temperature and mass flow rate. The first heat exchanger directs a second proportion of the first heat transfer fluid through a second heat dissipation portion wherein the second proportion of waste heat energy transfers to a second dissipation medium dependent on the second dissipation medium temperature and mass flow rate.

Abstract: A thermal management system for an aircraft comprises a first gas turbine engine, one or more first electric machines rotatably coupled to the first gas turbine engine, a first thermal bus, a first heat exchanger, and one or more first ancillary systems. The first thermal bus comprises a first heat transfer fluid, with the first heat transfer fluid being in fluid communication, in a closed loop flow sequence, between the or each first electric machine, the first gas turbine engine, the first heat exchanger, and the or each first ancillary system. Waste heat energy generated by at least one of the first gas turbine engine, the or each first electric machine, and the or each first ancillary system, is transferred to the first heat transfer fluid. The first heat exchanger is configured to transfer the waste heat energy from the first heat transfer fluid to a dissipation medium.

Abstract: A thermal management system for an aircraft includes a first gas turbine engine, one or more first electric machines rotatably coupled to the first gas turbine engine, a first thermal bus, and a first heat exchanger module. The first thermal bus includes a first heat transfer fluid, with the first heat transfer fluid being in fluid communication, in a closed loop flow sequence, between the first gas turbine engine, the or each first electric machine, and the first heat exchanger. Waste heat energy generated by at least one of the first gas turbine engine, and the or each first electric machine, is transferred to the first heat transfer fluid. The first heat exchanger module includes a first flow path and a second flow path. The first flow path is configured to direct a flow of the heat transfer fluid to a first heat dissipation portion.

Abstract: A thermal management system for an aircraft includes a thermal bus including one or more first heat sources; a heat sink; a vapour compression system including a compressor, a condenser, a receiver, a first side of a recuperator, an expansion valve, an evaporator, a second side of the recuperator, and the compressor; and one or more second heat sources. A heat transfer fluid transfers waste heat energy generated by the first heat sources to the heat sink. A second flow of waste heat energy from the second heat source(s) is transferred to a refrigerant. A third flow of heat energy in the refrigerant is transferred to the heat transfer fluid. The compressor includes a supplementary refrigerant reservoir, and the volume of refrigerant in the vapour compression system is increased during operation when a temperature of the heat transfer fluid is at or below a temperature of the second heat source(s).

Abstract: A thermal management system for an aircraft comprises a first gas turbine engine, a first thermal bus, a first heat exchanger, and a chiller. The first thermal bus comprises a first heat transfer fluid, with the first heat transfer fluid being in fluid communication, in a closed loop flow sequence, between the first gas turbine engine, the first heat exchanger, and the chiller. Waste heat energy generated by the first gas turbine engine, is transferred to the first heat transfer fluid. The chiller is configured to lower a temperature of the first heat transfer fluid prior to the first heat transfer fluid being circulated through the gas turbine engine. The first heat is exchanger is configured to transfer the waste heat energy from the first heat transfer fluid to a dissipation medium.

Abstract: A cylinder for combustion comprises a barrel having a cylindrical shape, and an air supply pipe. An insertion opening and a plurality of cooling flow paths are formed in the barrel. Collision region flow paths of the plurality of cooling flow paths each have a collision region circumvention flow path part intersecting with a collision gas axis extending in the direction of the flow of combustion gas moving toward a pipe center axis of the air supply pipe. The collision region circumvention flow path parts have an upstream-side direction component from the collision gas axis along the edge of the insertion opening, and have a downstream-side direction component from the collision gas axis along the edge of the insertion opening. No exit is formed in a portion of the collision region circumvention flow path parts within the range of specified angles about the pipe center axis.

Abstract: A gas turbine engine is provided. The gas turbine engine includes a compressor section, a combustion section, and a turbine section in a serial flow arrangement; and an air-to-air heat exchanger having an air-to-air heat exchanger potential defined by a product raised to a half power, the product being an effectiveness associated with the air-to-air heat exchanger multiplied by an airflow conductance factor associated with the gas turbine engine, and wherein the air-to-air heat exchanger potential is between 0.028 and 0.067 for a bypass ratio associated with the gas turbine engine between 3 and 10 and the effectiveness being between 0.5 and 0.9 and is between 0.015 and 0.038 for a bypass ratio associated with the gas turbine engine between 10 and 20 and the effectiveness being between 0.3 and 0.9.

Abstract: A method for operating a rotating detonation engine, having a radially outer wall extending along an axis; a radially inner wall extending along the axis, wherein the radially inner wall is positioned within the radially outer wall to define an annular detonation chamber having an inlet and an outlet, wherein the method includes flowing liquid phase fuel along at least one wall of the radially inner wall and the radially outer wall in a direction from the outlet toward the inlet to cool the at least one wall and heat the liquid fuel to provide a heated liquid fuel; flowing the heated liquid fuel to a mixer at the inlet to reduce pressure of the heated liquid fuel, flash vaporize the heated liquid fuel and mix flash vaporized fuel with oxidant to produce a vaporized fuel-oxidant mixture; and detonating the mixture in the annular detonation chamber.

Abstract: An assembly is provided for a gas turbine engine. This engine assembly includes an exhaust nozzle extending circumferentially about an axial centerline. The exhaust nozzle extends axially along the axial centerline to a trailing edge. The exhaust nozzle extends radially between an exterior inner surface and an exterior outer surface. The exhaust nozzle includes an inner skin, an outer skin and an internal cavity. The inner skin forms the exterior inner surface and includes a plurality of perforations. The internal cavity is disposed radially between the inner skin and the outer skin. The internal cavity is fluidly coupled with the perforations.

Abstract: An engine system is provided that includes a gas turbine engine and an actuation system. The gas turbine engine includes a compressor section, a combustor section, a turbine section and a flowpath extending through the compressor section, the combustor section and the turbine section. The actuation system includes a fluid reservoir, a flow regulator and a flow circuit. The flow regulator is configured as or otherwise includes a barrier between the fluid reservoir and the flow circuit. The flow regulator is configured to fluidly decouple the fluid reservoir from the flow circuit when the barrier is intact. The flow regulator is configured to fluidly couple the fluid reservoir with the flow circuit when the barrier breaks. The flow circuit is configured to direct gas from the fluid reservoir into the flowpath when the barrier breaks.

Abstract: A split ring seal has: a first circumferential end and a second circumferential end; an inner diameter surface and an outer diameter surface; a first axial end face and a second axial end face. The outer diameter surface has a sealing surface. The first axial end face has: a first section; a second section outboard of the first section; and an axial protrusion between the first section and the second section The second axial end face has: a first sealing surface; a second sealing surface radially outboard of the first sealing surface; a circumferential channel between the first surface and the second sealing surface; and a plurality of channels extending radially outward from the circumferential channel.

Abstract: Methods and systems of operating a gas turbine engine in a low-power condition are provided. In one embodiment, the method includes supplying fuel to the combustor by supplying fuel to the first fuel manifold via a first flow divider valve and supplying fuel to the second fuel manifold via a second flow divider valve. While supplying fuel to the combustor by supplying fuel to the first fuel manifold, the method includes stopping supplying fuel to the second fuel manifold and supplying pressurized gas to the second fuel manifold via the second flow divider valve to flush fuel in the second fuel manifold into the combustor and hinder coking in the second fuel manifold and associated nozzles.

Abstract: An assembly is provided for an aircraft propulsion system. This assembly includes a compressor section, a combustor section, a turbine section and a flowpath extending sequentially through the compressor section, the combustor section and the turbine section. The assembly also includes a turbine rotor, a propulsor rotor and an auxiliary turbine. The turbine rotor is within the turbine section. The turbine rotor is configured to rotatably drive the propulsor rotor. The auxiliary turbine includes an auxiliary turbine rotor. The auxiliary turbine rotor is configured to rotatably drive the propulsor rotor with the turbine rotor. The auxiliary turbine is configured to receive bleed gas from the flowpath.

Abstract: An assembly is provided for an aircraft propulsion system. This assembly includes a carrier, a first gear system, a second gear system and a lock device. The first gear system includes a sun gear, a first ring gear and a plurality of first intermediate gears between and meshed with the sun gear and the first ring gear. Each of the first intermediate gears is rotatably mounted to the carrier. The second gear system includes a second ring gear and a plurality of second intermediate gears meshed with the second ring gear. Each of the second intermediate gears is rotatably mounted to the carrier. The lock device is configured to lock rotation of the carrier about an axis during a first mode. The lock device is configured to unlock rotation of the carrier about the axis during a second mode.

Abstract: Gearboxes for aircraft gas turbine engines, in particular arrangements for journal bearings such gearboxes, and related methods of operating such gearboxes and gas turbine engines. A gearbox for an aircraft gas turbine engine includes: a sun gear; a plurality of planet gears surrounding and engaged with the sun gear; and a ring gear surrounding and engaged with the plurality of planet gears, each of the plurality of planet gears being rotatably mounted around a pin of a planet gear carrier with a journal bearing having an internal sliding surface on the planet gear and an external sliding surface on the pin.

Abstract: A method of operating an aircraft including a gas turbine engine and a fuel tank arranged to provide fuel to the gas turbine engine. The method includes: determining at least one fuel characteristic of the fuel arranged to be provided to the gas turbine engine; and proposing or initiating a change to a flight profile of the aircraft based on the at least one fuel characteristic. For example, intended route and/or altitude of the aircraft may be changed based on the one or more fuel characteristics.

Abstract: Embodiments of a flow system are provided. The flow system includes a flow source device configured to receive fluid from a first flow line and output fluid on a second flow line. An output is in fluid communication with the flow source device on the second flow line. An actuation system is in fluid communication with the second flow line through a third flow line. A flow meter measures a flow rate in the second flow line without bypassing any fluid to the first flow line. A controller is in communication with the flow source device and the flow meter. The controller adjusts the flow source device to achieve a desired flow rate at the output based on a condition of the flow source device and the flow rate measured by the flow meter.

Abstract: A dual-fuel turbine engine is a type of turbine engine that can operate using two different fuel sources, such as liquid fuel (e.g., diesel) and natural gas. For liquid fuels, the turbine engine includes a primary line to facilitate engine startup and a secondary line to modify power output. Generally, the turbine engine only uses one type of fuel for combustion at a time. The fuel lines and passages in the fuel nozzles that carry the unused fuel source should preferably be purged and sealed by a purging system to reduce coking. However, conventional purge systems typically purge and seal only the secondary line for liquid fuels. As a result, the passages in the fuel nozzles connected to the primary line are prone to coking. To address this problem, a purge system is disclosed that purges and seals the primary line and, optionally, the secondary line.