A method of managing thermal energy in a propulsion system includes diverting a flow of bleed air from a compressor section of the propulsion system. An amount of the flow of bleed air diverted from the compressor section is at least 5% of an inlet flow at an inlet of a high pressure compressor of the compressor section. The flow of bleed air is provided to a thermal management system. The flow of bleed air is passed through an expansion turbine of the thermal management system. The flow of bleed air is provided to a thermal load.

An airfoil assembly includes first and second fiber-reinforced composite airfoil rings each having inner and outer platform sections. A sealing shelf extends from a suction side wall of one of the first and second fiber-reinforced composite airfoil rings, and towards the pressure side wall of the other of the rings. A first seal is captured radially inwardly of the sealing shelves on outer platforms of each of the first and second fiber-reinforced composite airfoil rings. A second seal is captured inward of the sealing shelves on the inner platforms of each of the first and second fiber-reinforced composite rings to provide a seal between the respective first and second fiber-reinforced composite airfoil rings, and a captured internal component. A gas turbine engine is also disclosed.

An assembly for a compartment of a gas turbine engine that includes a housing, a gear, and a baffle disposed about the gear. The baffle includes an upstream portion and a downstream portion. The upstream portion includes an upstream inner wall and an upstream outer wall separated from the gear. The upstream inner wall is positioned between the upstream outer wall and the gear. An upstream flow channel is formed between the upstream inner wall and the upstream outer wall. The downstream portion of the baffle includes a downstream inner wall and a downstream outer wall separated from the gear. The downstream inner wall is positioned between the downstream outer wall and the gear. A downstream flow channel is formed between the downstream outer wall and the downstream inner wall.

A turbine blade for a gas turbine engine has an airfoil including leading and trailing edges joined by spaced-apart pressure and suction sides to provide an external airfoil surface extending from a platform in a spanwise direction to a tip. The external airfoil surface is formed in substantial conformance with multiple cross-sectional profiles of the airfoil defined by a set of Cartesian coordinates set forth in Table 1, the Cartesian coordinates provided by an axial coordinate scaled by a local axial chord, a circumferential coordinate scaled by a local axial chord, and a span location.

A building envelope, in particular a wall, a floor, or a roof of a building with at least two shells spaced some distance apart from one another, which encloses an intermediate space, said space being essentially empty with the exception of weight-bearing and/or construction-engineering elements or being filled at least in sections with porous, open-celled material and sealed from the interior and exterior of the building, wherein controllable sealing means are provided for sealing the intermediate space from the interior and exterior and optionally separated building envelope sections from one other.

A turbomachine for an aircraft extending axially along an axis X comprising a primary flow path in which an air stream flows intended for the combustion chamber and a secondary flow path in which an air stream flows intended for propulsion, the compressor comprising a plurality of first variable bypass valves, a plurality of second variable bypass valves, the turbomachine comprising a first control system configured to control the movement of the plurality of first variable bypass valves and a second control system configured to control the movement of the plurality of second variable bypass valves, the first control system and the second control system being separate.

A blade pivot of adjustable orientation for a turbomachine fan hub includes: a block having a retaining device configured to retain a fan blade root and a coupling device for the transmission of a torque; a ball type rolling bearing for taking up centrifugal forces having an inner ring; a clamping nut screwed onto an external thread of the block in order to clamp the inner ring of the ball type rolling bearing for taking up centrifugal forces to the block; a rolling bearing for taking up transverse forces; a pitch setting transmission ring positioned inside the inner radial end of the block and equipped with a coupling device cooperating with the coupling device of the block, and a device for locking the pitch setting transmission ring on the block.

A fan case assembly is adapted to extend around blades of a fan rotor included in a gas turbine engine. The fan case assembly includes an annular case that extends around an axis, a fan track liner coupled to the annular case and extending circumferentially at least partway about the axis, and a bolting arrangement that couples the fan track liner to the annular case.

A turbine assembly for use with a gas turbine engine includes a bladed wheel assembly and a vane assembly. The bladed wheel assembly is adapted to interact with gases flowing through a gas path of the gas turbine engine. The vane assembly is located upstream of the bladed wheel assembly and adapted to direct the gases at the bladed wheel assembly.

A flowpath component for a gas turbine engine includes a body having a leading edge and a trailing edge. A first exterior wall connects the leading edge to the trailing edge and a second exterior wall connects the leading edge to the trailing edge. At least one first internal cooling passage has a first inlet at a first end of the body. At least one second internal cooling passage has a second inlet at a second end of the body. The at least one first internal cooling passage is isolated from the at least one second internal cooling passage.