Abstract: A containment assembly is provided for an aircraft engine having a rotor with a set of blades. The containment assembly comprises a containment casing annularly surrounding the rotor radially outward of the set of blades. The containment casing is made of a material having a density less than that of steel. A layer of thermal insulation is disposed radially inward of the containment casing. The layer of thermal insulation is radially disposed between the containment casing and the set of blades.

Abstract: There is provided a power generation system for recuperating waste heat in an aircraft power plant. The power generation system includes a source of ammonia; an ammonia cracker in fluid communication with the source and decomposing the ammonia into hydrogen and nitrogen; a separator in fluid communication with the ammonia cracker, the separator separating the hydrogen and nitrogen from the ammonia cracker, the separator having a first outlet for a first output including hydrogen and a second outlet for a second output including hydrogen and nitrogen; a fuel cell in fluid communication with the first outlet of the separator, the fuel cell using the hydrogen of the first output of the separator and an oxidizing agent to generate electrical energy; and a thermal engine in fluid communication with the second outlet of the separator, the thermal engine using the hydrogen of the second output of the separator to generate mechanical energy.

Abstract: An aircraft engine oil distribution system includes a system container, an oil reservoir disposed in the system container, the oil reservoir having a plurality of fins protruding from an inner-surface of a wall of the oil reservoir and extending into the oil reservoir, and a heat exchanger disposed in the system container. The heat exchanger includes a cooling fluid passage defined at least between the one or more walls of the system container and the sidewall of the oil reservoir, a matrix disposed in the cooling fluid passage in heat transfer relationship with the plurality of fins of the oil reservoirs.

Abstract: A rotary internal combustion engine is provided that includes a housing and a rotor. The housing includes first and second side housings and a center housing. The center housing is disposed between and attached to the first and second side housings. The rotor is disposed within a rotor chamber and is engaged with a rotor shaft that extends between the first and second side housings. The rotor has a peripheral side wall that extends between a pair of end face surfaces. At least one of the first side housing or the second side housing includes a side plate having a seal surface, an interior surface, and a core disposed between the seal surface and the interior surface. The core comprises a ceramic matrix composite (CMC) material. The seal surface of the side plate engages in a sealing arrangement with a respective rotor end face surface.

Abstract: A scroll case is configured as a structural component to carry structural loads from other engine components, such as an internal containment ring, an exhaust casing and an engine tailpipe.

Abstract: The compressor section can have a centrifugal impeller operable to rotate around an axis, the centrifugal impeller having blades, a compressor inlet oriented towards the front and axially relative the axis, a compressor outlet oriented radially outwardly relative the axis, a diffuser having a diffusion flow path, a diffuser inlet in fluid flow communication with the compressor outlet, a diffusion flow path between a rear wall and a front wall; a collector extending circumferentially around the axis, having a collector inlet in fluid communication with the diffuser outlet, and a collector outlet; and hollow structural members protruding rearwardly from the rear wall, the hollow structural members being circumferentially interspaced from one another, each hollow structural member having a length extending radially along the rear wall and having an internal conduit extending radially inwardly along the length.

Abstract: An aircraft engine, comprising a thermal engine, an axial turbine having a turbine inlet, and an exhaust assembly fluidly connecting the thermal engine to the axial turbine. The exhaust assembly includes a housing and a deflector removably mounted within the housing. The deflector has circumferentially distributed vanes. The deflector is a first deflector, having a first set of geometric characteristics, that is removable from the housing and replaceable by a second deflector having a second set of geometric characteristics different from the first set of geometric characteristics. The first deflector is one of a first class of deflectors and the second deflector is one of a second class of deflectors, the first and second class of deflectors respectively defining first and second exhaust flow profiles that differ from each other.

Abstract: An exhaust assembly for a gas turbine engine includes an outer exhaust case, an inner exhaust case, and a hollow strut. The outer exhaust case forms an outer cavity radially outward of the outer exhaust case. The inner exhaust case is positioned radially inward of the outer exhaust case. The outer exhaust case and the inner exhaust case form a core flow path. The inner exhaust case forms a centerbody. The hollow strut includes a strut body, an inlet, an outlet, and an internal passage. The strut body is connected to the outer exhaust case and the inner exhaust case. The internal passage extending through the strut body from the inlet to the outlet. The inlet is located at the outer radial end. The inlet is in fluid communication with the outer cavity. The internal passage is configured to direct gas from the outer cavity to the outlet.

Abstract: A gas turbine intake has a swirl housing having a tangential inlet fluidly connecting an exhaust conduit, an annular outlet defined around a central axis and fluidly connecting a turbine gas path, a swirl path extending circumferentially around the central axis from the tangential inlet, and vanes located in the swirl housing, the vanes circumferentially interspaced from one another relative the central axis, each vane having a twisted and flat body having a length extending from a leading end to a trailing end, the leading end being oriented mainly circumferentially and axially at the swirl path, the trailing end being oriented mainly axially and radially at the annular outlet, the twisted and flat body twisting between leading and trailing ends around the central axis, around a radial axis perpendicular to the central axis, and around a tangential axis perpendicular to both the central axis and the radial axis.

Abstract: The compressor section can have a centrifugal impeller operable to rotate around an axis, the centrifugal impeller having blades, a compressor inlet oriented towards the front and axially relative the axis, a compressor outlet oriented radially outwardly relative the axis, a diffuser having a diffusion flow path, a diffuser inlet in fluid flow communication with the compressor outlet, a diffusion flow path between a rear wall and a front wall; a collector extending circumferentially around the axis, having a collector inlet in fluid communication with the diffuser outlet, and a collector outlet; and hollow structural members protruding rearwardly from the rear wall, the hollow structural members being circumferentially interspaced from one another, each hollow structural member having a length extending radially along the rear wall and having an internal conduit extending radially inwardly along the length.

Abstract: A turbine assembly of an aircraft engine includes a cooling system for optimizing a tip clearance gap defined between an inner surface of a turbine housing and blade tips of the turbine blades. The cooling system includes a cooling airflow passage located radially outward from the turbine housing and being in heat-transfer communication with the turbine housing. The cooling airflow passage receives a flow of cooling air therethrough for cooling the turbine housing. A heat sink is disposed on the outer surface of the turbine housing within the cooling airflow passage, the heat sink including heat transfer elements projecting into the cooling airflow passage away from the outer surface of the turbine housing. The heat transfer elements are in convective heat transfer relationship with the flow of cooling air in the cooling airflow passage.

Abstract: A gas turbine intake has a swirl housing having a tangential inlet fluidly connecting an exhaust conduit, an annular outlet defined around a central axis and fluidly connecting a turbine gas path, a swirl path extending circumferentially around the central axis from the tangential inlet, and vanes located in the swirl housing, the vanes circumferentially interspaced from one another relative the central axis, each vane having a twisted and flat body having a length extending from a leading end to a trailing end, the leading end being oriented mainly circumferentially and axially at the swirl path, the trailing end being oriented mainly axially and radially at the annular outlet, the twisted and flat body twisting between leading and trailing ends around the central axis, around a radial axis perpendicular to the central axis, and around a tangential axis perpendicular to both the central axis and the radial axis.

Abstract: The gas turbine intake can have a swirl housing having an inlet portion fluidly connecting an exhaust conduit, an annular outlet defined around a central axis and fluidly connecting a turbine gas path, a swirl path extending circumferentially around the central axis from the inlet portion to a circumferential outlet, the circumferential outlet fluidly connected back into the inlet portion, and vanes located in the swirl housing, the vanes circumferentially interspaced from one another relative the central axis and located radially inwardly from the swirl path relative the central axis, the swirl path being free of the vanes.

Abstract: The gas turbine intake can have a swirl housing having an inlet portion fluidly connecting an exhaust conduit, an annular outlet defined around a central axis and fluidly connecting a turbine gas path, a swirl path extending circumferentially around the central axis from the inlet portion to a circumferential outlet, the circumferential outlet fluidly connected back into the inlet portion, and vanes located in the swirl housing, the vanes circumferentially interspaced from one another relative the central axis and located radially inwardly from the swirl path relative the central axis, the swirl path being free of the vanes.

Abstract: A containment assembly is provided for an aircraft engine having a rotor with a set of blades. The containment assembly comprises a containment casing annularly surrounding the rotor radially outward of the set of blades. The containment casing is made of a material having a density less than that of steel. A layer of thermal insulation is disposed radially inward of the containment casing. The layer of thermal insulation is radially disposed between the containment casing and the set of blades.

Abstract: A gas turbine engine assembly includes an engine core and an electric machine. The engine core includes a first rotating structure, a second rotating structure, a combustor and a flowpath. The first rotating structure includes a first structure turbine rotor. The second rotating structure includes a second structure compressor rotor, a second structure turbine rotor and a second structure shaft connecting the second structure compressor rotor to the second structure turbine rotor. The second structure compressor rotor, the combustor, the second structure turbine rotor and the first structure turbine rotor are arranged sequentially along the flowpath. The electric machine is arranged within the engine core. The electric machine includes an electric machine rotor and an electric machine stator adjacent the electric machine rotor. The electric machine rotor is rotatable with the second rotating structure and located between the second structure compressor rotor and the first structure turbine rotor.

Abstract: The gas turbine intake can have a swirl housing assembly with a tangential inlet fluidly connecting an exhaust conduit, an annular outlet defined around a central axis and fluidly connecting a turbine gas path, a swirl path extending around the central axis and between the tangential inlet and the annular outlet, the swirl housing assembly having a proximal portion defining a first portion of the swirl path, a distal portion defining a second portion of the swirl path, vanes located in the swirl housing assembly, the vanes circumferentially interspaced from one another relative the central axis and extending between the proximal portion and the distal portion, the proximal portion fastened to the distal portion via a plurality of fasteners, a gasket sandwiched between the proximal portion and the distal portion by the plurality of fasteners, the gasket extending in a radial plane relative the central axis.

Abstract: A tangential on-board injector (TOBI), comprising: a body defining an annular passageway to receive cooling air, the TOBI defining a plurality of discharge nozzles; a rotating component mounted for rotation relative to the body about an axis of rotation; a seal extending between the body and the rotating component; a plurality of vanes circumferentially distributed about the axis of rotation and located downstream of the plurality of discharge nozzles relative to a flow of the cooling air circulating toward the seal from the plurality of discharge nozzles and upstream of the seal; and flow passages defined between the plurality of vanes, a flow passage of the flow passages extending along a passage axis, the passage axis having a tangential component at an outlet of the flow passage that is different than a tangential component of an exit flow axis of a nozzle of the plurality of discharge nozzles.

Abstract: There is provided a load transfer interface in an aircraft engine for transferring a load from a bearing housing to an engine casing. The load transfer interface comprises a first component operatively coupled to and receiving the load from the bearing housing. The first component has a first annular body with a spigot extending axially from the first annular body. The interface comprises a second component operatively coupled to the first component and to the engine casing. The second component has a second annular body with a spigot-receiving cavity disposed therein. The spigot-receiving cavity is shaped and positioned to receive the spigot of the first component. The second component receives the load from the first component and transfers the load to the engine casing.

Abstract: A turbine assembly in a turbine section of an aircraft engine includes a rotor with blades having blade tips, and a turbine housing radially surrounding the blades. A distance between an inner surface of the housing and the blade tips defines a tip clearance gap. A passive cooling system for optimizing the tip clearance gap includes a cooling airflow passage located radially outward from, and in heat-transfer with, the turbine housing. The cooling airflow passage has an inlet opening located upstream of the rotor and an exit opening located downstream of the rotor. The inlet opening provides air flow into the cooling airflow passage. The exit opening provides air flow communication between the cooling airflow passage and a main gaspath of the turbine section. A flow of cooling air through the cooling airflow passage is induced, to cool the housing.