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.

Methods for depositing protective coatings on aerospace components are provided and include sequentially exposing the aerospace component to a chromium precursor and a reactant to form a chromium-containing layer on a surface of the aerospace component by an atomic layer deposition process. The chromium-containing layer contains metallic chromium, chromium oxide, chromium nitride, chromium carbide, chromium silicide, or any combination thereof.

A rocket nozzle is made from an optimized metal lattice structure, with a hardened material applied onto the metal lattice structure so as to coat the structure and fill voids in the lattice by chemical vapor deposition. A rocket nozzle is further provided having one or more bypass lines for taking expanding gas from a combustion chamber of the rocket nozzle and redirecting the expanding gas to a skirt of the rocket nozzle to thereby manipulate the shape of a plume of expanding gas exiting the rocket nozzle. A rocket nozzle is also provided having one or more main injectors extending into the combustion chamber for injecting fuel for combustion into the combustion chamber and one or more opposing injectors oriented to direct an opposing flow of energy and gas expansion towards the main injectors, or having one or more secondary injectors arranged around the combustion chamber proximal the throat.

Joining an airfoil with a platform by mechanical keying can provide advantages in applications of ceramic materials, such as ceramic matrix composites.

Disclosed is a ceramic matrix component having a fibrous core and a ceramic matrix composite shell surrounding at least a portion of the fibrous core. The ceramic matrix composite shell comprises a fibrous preform. The fibrous core has a greater porosity than the fibrous preform. A method of making the ceramic matrix component is also disclosed.

Disclosed is a ceramic matrix component having a fibrous core and a ceramic matrix composite shell surrounding at least a portion of the fibrous core. The fibrous core has a three dimensional braided structure and cooling passages. A method of making the ceramic matrix component is also disclosed.

Methods for depositing protective coatings on aerospace components are provided and include sequentially exposing the aerospace component to a chromium precursor and a reactant to form a chromium-containing layer on a surface the aerospace component by an atomic layer deposition process. The chromium-containing layer contains metallic chromium, chromium oxide, chromium nitride, chromium carbide, chromium silicide, or any combination thereof.

Embodiments of the present disclosure generally relate to protective coatings on turbine blades, turbine disks, and other aerospace components and methods for depositing the protective coatings. In one or more embodiments, a turbine blade includes a blade portion and a root coupled to the blade portion, where the root contains a protective coating disposed thereon. The protective coating is or contains one or more deposited crystalline film containing at least one of a metal oxide, a metal nitride, or a metal oxynitride and has a thickness of about 100 nm to about 10 μm. In some examples, a turbine blade assembly includes a disk and a plurality of the turbine blades coupled to the disk. The protective coating is disposed on the root on the turbine blade and/or a receiving surface on the turbine disk.

A blade and method for producing the blade for a gas turbine engine are described herein. The blade may include a composite airfoil. The airfoil may comprise a ceramic material, and a distal end. A tip may extend from the distal end of the airfoil. The tip of the airfoil may comprise a substantially porous structure and may comprise infiltrated material extending from an airfoil preform to a tip preform to join the airfoil preform and the tip preform.

A method for manufacturing a turbine engine vane a root connected to a blade extending in a longitudinal direction includes the steps of providing a root; and providing mold with a first cavity and a second cavity that together define a recess in which the vane is formed. The recess includes a first space in which the blade is formed and a second space in which the root (1), c) is formed. The method further includes the steps of providing aluminum strips; positioning a fibrous reinforcement; arranging the vane root in the second space; and injecting a foam comprising aluminum or injecting an aluminum alloy into the first space of the recess of the mold such that the foam impregnates the fibrous reinforcement.