A gas turbine engine includes a fan rotor driven by a fan drive turbine about an axis through a gear reduction to reduce a speed of the fan rotor relative to a speed of the fan drive turbine. A fan case surrounds the fan rotor, and a core engine with a compressor section, including a low pressure compressor. The fan rotor delivers air into a bypass duct defined between the fan case and the core engine. A rigid connection is between the fan case and the core engine includes three triangular-frame connecting members rigidly connected to the fan case at a fan case connection point, and to the core engine at a core engine connection point. The triangular-frame connecting members each are defined by two rigid legs which extend between the fan case and to the core engine, along directions each have a component extending radially inwardly and a component in opposed circumferential directions to each other. A plurality of non-structural fan exit guide vanes and the non-structural fan exit guide vanes are provided with an acoustic feature to reduce noise. The non-structural fan exit guide vanes are rigidly mounted to at least one of the fan case and the core engine.

A turbine nozzle assembly for use in a turbine engine is provided. The assembly includes an inner barrel and a turbine nozzle support ring. The inner barrel has a forward end and an aft end. The turbine nozzle support ring includes an annular body that defines a forward end, an opposite aft end, an inner surface, and an opposite outer portion. The forward end of the annular body is coupled to the aft end of the inner barrel. The annular body includes a first arcuate segment and a second arcuate segment removably coupled to the first arcuate segment. The first arcuate segment has a first arcuate length and the second arcuate segment has a second arcuate length. The second arcuate length is shorter than the first arcuate length.

A gas turbine fan casing made of composite material with a fibrous reinforcement includes a plurality of superimposed turns of a strip-shaped fibrous texture having a three-dimensional weaving between a plurality of layers of warp yarns and a plurality of layers of weft yarns, the fibrous reinforcement being densified by a matrix. The fibrous texture includes at least one lateral section of variable thickness in which the weft yarns have a size or a count different from the size or the count of the weft yarns of the plurality of layers of weft yarns present in the remainder of the fibrous texture.

A pressure relief latch includes a housing configured to be mounted to a pressure relief door. A bolt is coupled to the housing about a first axis and is rotatable about the first axis between a first bolt position and a second bolt position. The bolt includes a bearing centered about a second axis. The bearing is configured to rotate about the second axis. A spring assembly includes a first assembly end mounted to the bearing and a second assembly end in rotational communication with the housing about a third axis. The spring assembly is rotatable about the second and third axes and further includes a spring extending between the first and second assembly ends. The bolt is configured to permit movement of the pressure relief door from the closed position to the open position as the bolt rotates from the first bolt position to the second bolt position.

A barrier device is provided for an inlet cowl for an aircraft engine including an inner barrel circumferentially surrounding an opening in the inlet cowl formed along an axis of rotation of the aircraft engine, through which air passes to the aircraft engine, the inner barrel including a face sheet disposed on a radially inward side of the inner barrel relative to the axis. The barrier device includes a containment doubler of the inner barrel, disposed on a radially outward side of the inner barrel, and a blade fragment barrier including one or more strips disposed between the containment doubler and the face sheet, so as to extend circumferentially at least partially around the opening and to occupy a radial distance between the containment doubler and the face sheet.

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.

A turbine housing has a scroll passage. The scroll passage includes: an outer peripheral surface extending along an axial direction of the turbine housing; an inner peripheral surface disposed inward of the outer peripheral surface in a radial direction of the turbine housing; a one-side surface, which is a side surface on one side in the axial direction of the turbine housing, extending along the radial direction of the turbine housing; an other-side surface, which is a side surface on another side in the axial direction of the turbine housing, disposed closer to an outlet of the turbine housing than the one-side surface and extending along the radial direction of the turbine housing; a one-side outer peripheral R portion connecting an outer peripheral end of the one-side surface and a one-side end of the outer peripheral surface; an other-side outer peripheral R portion connecting an outer peripheral end of the other-side surface and an other-side end of the outer peripheral surface; and an other-side inner peripheral R portion connecting an inner peripheral end of the other-side surface and an other-side end of the inner peripheral surface. In a cross-sectional view of the scroll passage, when a ratio of a width dimension of the scroll passage along the axial direction to an R dimension of each of the one-side outer peripheral R portion, the other-side outer peripheral R portion, and the other-side inner peripheral R portion is defined as a one-side outer peripheral R ratio, an other-side outer peripheral R ratio, and an other-side inner peripheral R ratio, respectively, the scroll passage has an R ratio increasing region where at least one of the one-side outer peripheral R ratio, the other-side outer peripheral R ratio, and the other-side inner peripheral R ratio increases from upstream to downstream in the scroll passage.

An intermediate casing includes an inner wall having an outer surface extending along a longitudinal axis, the outer surface having a first profile in a first plane, an outer wall having an inner surface a second profile in the first plane, and first through fourth arms each extending radially. The outer and inner surfaces and the first and second arms delimit a first cavity, the first cavity having a first area and the outer and inner surfaces being separated by a first distance. The outer and inner surfaces and the third and fourth arms delimit a second cavity, the second cavity having a second area and the outer and inner surfaces being separated by a second distance. The first and second profiles are such that the first area is substantially identical to the second area and the first distance is different from the second distance.

An aircraft turbomachine having a centrifugal compressor, an annular combustion chamber, an annular casing extending around the chamber and delimiting an annular space (E) in which the chamber is situated, and a heat exchanger. The heat exchanger can include a first circuit supplied with exhaust gas from the turbomachine, and a second circuit connected by first and second volutes respectively to an outlet of the compressor and to the annular space. The first and second volutes can be positioned at an axial distance from one another, and the second volute is can be connected to the annular space by a straightener which is situated at least in part outside the casing and which is integrated into an annular connecting pipe which connects the second volute to this casing.

A fan case assembly adapted for use with a gas turbine engine includes a fan track liner and an annular case. The fan track liner extends circumferentially at least partway about a central axis of the gas turbine engine. The annular case is configured to support the fan track liner at a radial position relative to the central axis. The fan case assembly further includes an air recirculation duct configured to redirect air around the fan track liner.