A connecting structure (10) for load transfer, in particular in a gas turbine (1), including a strut (20) and at least one wall element (30) is provided. The strut (20) at one end is integrally joined to the wall element (30), and the strut (20) and the wall element (30) are enclosed by a fillet (40), at least in areas, and integrally joined to same. An elastic deformation of the involved elements of the structure during the load transfer and/or load absorption is improved in that a root section (50) that is formed by a ridge (56) and that extends from the strut (20) to the wall element (30) is situated on the fillet (40).

An annular intermediate casing for a turbomachine through which an aerodynamic airstream circulates, including a radially internal shroud (31), a radially external shroud, at least one stator blade including a vane with a leading edge, and at least one aero-dynamic member which is formed of a structural arm extending radially at least in part between the radially internal shroud and the radially external shroud. The aerodynamic member includes a leading edge which is aligned with the leading edge of the stator blade in one plane and which is borne by a profiled portion positioned upstream of the structural arm in the direction of circulation of the aerodynamic airstream, the profiled portion, the structural arm, the radially internal shroud and the radially external shroud being monobloc.

A system for retaining stators and reducing air leakage in a gas turbine engine having an axis includes a stator having an inner platform, an outer platform, a low pressure side, a high pressure side, and at least one foot, and designed to turn air. The system also includes a case positioned radially outward from the stator and having at least one recess designed to interface with the at least one foot to resist movement of the stator relative to the case. The system also includes a bladder positioned between the outer platform of the stator and the case and designed to receive pressurized fluid having a greater pressure than ambient pressures experienced at the low pressure side of the stator and to further resist movement of the stator relative to the case in response to receiving the pressurized fluid.

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 blade and gas turbine include a stationary blade main body provided internally with cavities, and an inner shroud linked to an end portion, in the longitudinal direction, of the stationary blade main body, and which is internally provided with an inner shroud cooling passage with which a first cavity is in fluid communication. The inner shroud is provided, in front edge corner portions, with a first chamfered portion intersecting a front surface, a side surface, and an upper surface, and first cooling holes in fluid communication with the inner shroud cooling passage are provided in the first chamfered portion.

A turbine vane assembly adapted for use in a gas turbine engine includes a spar, a turbine vane, and load transfer pins. The spar comprises metallic materials and is configured to support other components of the turbine vane assembly relative to an associated turbine case. The turbine vane comprises ceramic matrix composite materials and is shaped to include an airfoil configured to direct the flow of hot gasses through a primary gas path of the turbine vane assembly.

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 variable guide vane assembly is provided for a turbine defining a core air flowpath. The variable guide vane assembly includes an airfoil band defining a flowpath surface and a cavity. The variable guide vane assembly further includes an airfoil including a first end extending at least partially into the cavity of the airfoil band and an opposite second end, the airfoil extending generally along an axis between the first end and the second end and being moveable generally about the axis relative to the airfoil band. The variable guide vane assembly further includes a sealing element operable to form a seal between the first end of the airfoil and the airfoil band.

A gas turbine engine includes a fan rotor driven by a fan drive turbine about an axis through a gear reduction. An inner core engine has an inner core engine housing surrounding a compressor section, including a low pressure compressor. A rigid connection between a fan case and the inner core engine includes A-frames rigidly connected at a connection point to the fan case. Fan exit guide vanes rigidly connect to the fan case, and to the inner core engine. A fan intermediate case is positioned forward of a first rotor stage in the low pressure compressor. A rigid structure is connected to the inner core engine and to the fan exit guide vanes. The rigid structure defines a structure moment stiffness. The fan intermediate case defines an intermediate case moment stiffness. A ratio of the structure moment stiffness to the intermediate case moment stiffness is between 5 and 15.

The invention proposes an aeronautical turbine engine assembly comprising an upstream casing (55) to which guide blading (48a) is fastened, and a downstream casing (58) to which a sealing element (62) provided with an abradable material for rotor blading is fastened. This assembly further comprises a shroud ring (66) placed between the upstream casing and the downstream casing and fastening means (68) for detachably fastening the shroud ring. In order to be fastened to the upstream casing, the guide blading (48a) of the turbine engine is mounted on a downstream hook (480b) of the upstream casing, without being hooked onto the shroud ring (66), and the downstream casing (58) has an upstream hook with which the sealing element (62) is engaged in order to be fastened to the downstream casing, or the shroud ring has an upstream hook on which the sealing element (62) is mounted so as to be fastened to the downstream casing.