A gas turbine engine includes at least one trap that absorbs or adheres to calcium-magnesium-alumino-silicate (CMAS) entrained in intake air entering the engine.

A coating on a substrate is disclosed having layers including yttrium aluminum garnet (YAG) and yttrium aluminum perovskite (YAP).

Flowpath assemblies, methods of forming flowpath assemblies, and hypersonic vehicles are provided. For example, a flowpath assembly for a combustor comprises a tube array comprising a plurality of tubes, a joining material disposed between adjacent tubes of the plurality of tubes to join together the adjacent tubes, a flowpath layer, and an outer layer. The plurality of tubes and the joining material are disposed between the flowpath layer and the outer layer. The flowpath layer defines a combustion flowpath. Each of the plurality of tubes, the joining material, the flowpath layer, and the outer layer are formed from a composite material. The combustor comprising the flowpath assembly may be included in a ramjet engine of a hypersonic vehicle. A fabrication method may include laying up composite plies to form a tube array including the plurality of tubes, the joining material, and the flowpath and outer layers.

The invention relates to a ceramic sealing system between a rotor blade (120) and a housing (I). A small porous zirconium oxide layer (II) on a turbine rotor blade, which zirconium oxide layer faces a ceramic layer system (15′, 15″) of higher porosity, achieves durable sealing systems. The housing (I″) has a metal substrate (7), a metal adhesion-promoting layer (10), and a thick, outer, ceramic layer (15′, 15″) based on zirconium oxide, in particular having a porosity 2≥14%.

A solid-propellant rocket engine (1) has a casing (2) and a thermal protection (15) internally coating the casing and delimiting a housing (17), which contains a mass of solid propellant (3); the thermal protection has a fixed portion (22) and at least one movable portion (23) that adheres to the mass of solid propellant (3) and can be moved from a back position to a forward position with respect to the fixed portion (22) through a thrust system obtained by pressuring a chamber 31 provided by installing a membrane 32 between the fixed portion 22 and the movable portion 23; the engine is tested by verifying the adhesion of the mass of solid propellant (3) to the movable portion (23) after having moved the movable portion (23) to the forward position by means of a thrust directed from the fixed portion towards the mass of solid propellant (3).

A system includes a layered structure. The layered structure includes first and second coalesced layers and an intermediate layer disposed between the first and second coalesced layers. The first and second coalesced layers have a higher degree of coalescence than the intermediate layer.

A property gradient coating to be applied by additive manufacture to an inner wall of a casing mounted on the periphery of moving blades of a turbomachine rotor, the coating including in superimposed layers from an outer surface of the coating to this inner wall of the casing, on the one hand, a first layer consisting of a three-dimensional scaffolding of filaments of an abradable material forming an ordered network of channels or microchannels whose pore sizes are between 50 and 250 microns and whose porosity is greater than 85%, and, on the other hand, a second layer having a function of dissipating energy from acoustic waves striking the outer surface of the coating and consisting of a three-dimensional scaffolding of filaments of a first thermosetting material forming an ordered network of channels or microchannels whose pore sizes are between 50 and 400 microns and whose porosity is greater than 60%.

An acoustic device includes: a perforated plate that has a plurality of holes penetrating in a plate thickness direction of the perforated plate and in which a main flow flows on one side of the perforated plate in the plate thickness direction; and a housing that is provided on the other side of the perforated plate in the plate thickness direction and partitions a space between the housing and the perforated plate, wherein part of the holes on the one side in the thickness direction is inclined to at least one of the one side and the other side of a flow direction of the main flow.

A solid-propellant rocket engine (1) has a casing (2) and a thermal protection (15) internally coating the casing and delimiting a housing (17), which contains a mass of solid propellant (3); the thermal protection has a fixed portion (22) and at least one movable portion (23) that adheres to the mass of solid propellant (3) and can be moved from a back position to a forward position with respect to the fixed portion (22) through a thrust system obtained by pressuring a chamber 31 provided by installing a membrane 32 between the fixed portion 22 and the movable portion 23; the engine is tested by verifying the adhesion of the mass of solid propellant (3) to the movable portion (23) after having moved the movable portion (23) to the forward position by means of a thrust directed from the fixed portion towards the mass of solid propellant (3).

An airfoil includes a root and a tip, which define a span of the airfoil therebetween. The airfoil also includes a leading edge and a trailing edge downstream of the leading edge along a flow direction. The leading edge and the trailing edge each extend across the span of the airfoil from the root to the tip. The airfoil further includes a pressure side surface and a suction side surface. The airfoil also includes a thermal barrier coating on the pressure side surface and the suction side surface. The thermal barrier coating includes a base layer and a top coat. A thickness of the base layer varies across each of the pressure side surface and the suction side surface with a maximum thickness of the base layer at the leading edge.