A mandrel for a molding process includes a first portion and a second portion that form a mandrel having a channel between the first and second portions. At least one locking pin is configured to urge the first and second portions away from each other to establish a gap in response to inserting the locking pin into the channel along a respective end of the mandrel. A molding apparatus and method of forming a matrix composite component are also disclosed.

A turbine rotor in an embodiment includes: a rotor body portion; and a plurality of turbine disks provided on the rotor body portion in a center axis direction of the rotor body portion. The turbine rotor includes: a high-pressure cooling passage formed in the rotor body portion, the high-pressure cooling passage to which a high-pressure cooling medium is supplied, and the high-pressure cooling passage that discharges the high-pressure cooling medium to the high-pressure side turbine stage; and a low-pressure cooling passage formed in the rotor body portion, the low-pressure cooling passage to which a low-pressure cooling medium whose pressure is lower than the pressure of the high-pressure cooling medium is supplied, and the low-pressure cooling passage that discharges the low-pressure cooling medium to the low-pressure side turbine stage.

Systems and methods for delivering a buffer fluid to a shaft seal of a gas turbine engine are provided. An exemplary system includes, a buffer fluid source, one or more first conduits providing fluid communication between the buffer fluid source and the shaft seal along a first route, and one or more second conduits providing fluid communication between the buffer fluid source and the shaft seal along a second route different from the first route. A heat exchanger is also disposed along the first route to facilitate heat transfer between buffer fluid in the one or more first conduits and a cooling fluid.

An aircraft including a fuselage and a propulsion assembly. The propulsion assembly includes at least one fan rotor placed behind the fuselage as an extension thereof along a longitudinal axis, and a nacelle which forms a fairing of the at least one fan rotor through which at least one air flow passes. The aircraft comprises a plurality of stator radial arms mounted upstream of the at least one fan rotor and extending between the fuselage and the nacelle. The radial arms comprise at least one variable-pitch movable portion configured to axially divert the air flow.

An aircraft including a fuselage and a propulsion assembly. The propulsion assembly includes at least one fan rotor placed behind the fuselage as an extension thereof along a longitudinal axis, and a nacelle which forms a fairing of the at least one fan rotor through which at least one air flow passes. The aircraft comprises a plurality of stator radial arms mounted upstream of the at least one fan rotor and extending between the fuselage and the nacelle. The radial arms comprise at least one variable-pitch movable portion configured to axially divert the air flow.

A gas turbine engine includes a combustion section, a turbine section, and a compressor section. The combustion section includes a combustor casing, a combustor, a cooling duct, and an outer duct. The combustor casing defines at least in part a diffuser cavity and a fluid inlet. The combustor disposed is in the diffuser cavity. The cooling duct is in fluid communication with the fluid inlet in the combustor casing and is configured to transport a flow of cooled air. The outer duct surrounds at least a portion of the cooling duct and extends along a portion of an entire length of the cooling duct. The outer duct defines a gap with the cooling duct and is configured to transport a flow of buffer air. The turbine section is disposed downstream from the combustion section. The cooling duct is in fluid communication with the turbine section.

A nozzle blade for a turbine engine includes a blade body in which at least one through-cavity is provided extending between an inner end of the blade and an outer end of the blade and housing, respectively, a first and a second liner, the first liner extending along a trailing edge of the blade, the second liner extending along a leading edge of the blade, wherein each of the first and second liners includes an intertwining system suitable for linking the first liner to the second liner, independently of the blade body.

A fire resistance device is intended to be interposed between an upstream end of an aircraft turbine engine mounting structure and a cowling of the turbine engine delimiting an inter-flow compartment.

A modular casing manifold for cooling fluids of a gas turbine engine is presented. The modular casing manifold has an annular shape including an axial inner plate, an axial outer plate, a radial forward plate and a radial aft plate. The forward plate is attached to the inner and outer plates at forward end. At least a portion of the aft plate is attachable to and removable from the inner and outer plates at aft end for enabling cooling fluid to cool turbine blades of the gas turbine engine. The modular casing manifold includes preswirler segments. At least a number of the preswirler segments are attachable to and removable from the forward plate for enabling cooling fluid to cool turbine blades of the gas turbine engine. The modular casing manifold enables alternative cooling fluids to cool turbine blades of the gas turbine engine with minimal cost and assembly flexibility.

The present technique presents a turbomachine component having an airfoil e.g. a vane of a gas turbine. The airfoil wall defines an internal space which includes a first and a second cooling channels having a first and a second impingement inserts, that define a first main and a first peripheral flow channels in the first cooling channel and a second main and a second peripheral flow channels in the second cooling channel, respectively. Impingement jets ejected from the main flow channels via impingement holes of the corresponding impingement inserts are received in the corresponding peripheral flow channels. A channel connecting conduit conducts a flow of the cooling air from the first cooling channel to the second cooling channel. The channel connecting conduit includes an inlet connected to an outlet of the first cooling channel, and an outlet connected to an inlet of the second cooling channel.