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 system includes a combustor. The combustor has a combustor wall with a combustor dome at an upstream end of the combustor wall, and an outlet at a downstream end of the combustor wall opposite the upstream end. The combustor wall includes an inner wall portion and an outer wall portion defining an interior of the combustor therebetween. Each of the inner wall portion and outer wall portion extends from the combustor dome to the downstream end of the combustor wall. The combustor wall includes an air cooling passage embedded inside at least one of the inner wall portion and the outer wall portion. The air cooling passage extends from the upstream end of the combustor wall to the downstream end of the combustor wall.

A hybrid three-layer system is presented. The hybrid three-layer system includes a two-layer composite system and an additively manufactured third layer comprising a lattice structure. The composite layer system includes a metallic substrate, a structured surface, and a thermal protection system. The structured surface may be additively manufactured onto the metallic substrate and includes structured surface features formed to project above the metallic substrate. Each of the structured surface features are separated from adjacent structured surface features by grooves. The thermal protection coating may be thermally sprayed onto the structured surface and is bonded to each of the structured surface features. The lattice structure is in contact with a surface of the metallic substrate of the composite layer system.

A method of machining includes removing material from a target region of a ceramic component to form a blind opening in the ceramic component via removing a bulk of the material by a laser machining operation and then removing a remainder of the material by a mechanical machining operation.

A heat shielded assembly includes a fuel structure of a combustor of a gas turbine engine and a woven heat shield at least partially conformally surrounding the fuel structure and spaced from an exterior of the fuel structure by a distance where it surrounds the fuel structure. The fuel structure is configured to deliver fuel to the combustor. The woven heat shield comprises a first set of strands, a second set of strands interwoven with the first set of strands, and a weave pattern comprising the first set of strands and the second set of strands. Each strand of the first set of strands extends in a first direction, each strand of the second set of strands extends in a second direction transverse to the first direction, and the first set of strands and the second set of strands are not attached where they intersect in the weave pattern.

An assembly is provided for a gas turbine engine. This assembly includes a combustor and a fuel conduit. The combustor includes a combustor wall that forms a peripheral boundary of a combustion chamber within the combustor. The fuel conduit extends along and is formed integral with the combustor wall. The fuel conduit is disposed outside of the combustion chamber.

An assembly is provided for a gas turbine engine. This gas turbine engine assembly includes a stationary engine structure. The stationary engine structure includes a diffuser, a combustor, an engine case and a plenum. The combustor is disposed within the plenum. The engine case forms a peripheral boundary of the plenum. A gas path extends sequentially through the diffuser, the plenum and the combustor. A first section of the stationary engine structure is formed as a first monolithic body. The first section includes the diffuser and the combustor. A second section of the stationary structure is formed as a second monolithic body. The second section is configured as or otherwise includes the engine case.

A turbine engine part coated in at least a first ceramic layer forming a thermal barrier and including a ceramic material with first ceramic fibers dispersed in the first layer. The first layer may have a chemical composition gradient between a material for forming a thermal barrier and a material for providing protection against calcium and magnesium aluminosilicates, which is present at a greater content in an outer zone of the first layer, and/or the first layer may be porous and may present a porosity gradient such that an outer portion of the first layer presents lower porosity.

A support housing for use in distributing fuel in a gas turbine engine includes a main body defining an inlet aperture, a plurality of outlet apertures, and a substantially planar mounting surface. A first fuel channel has a wall that defines a first flow space and a support member extends across the first flow space and has a long axis oriented at an oblique angle with respect to the mounting surface.

A manufacturing method is provided. During this method, a preform component is provided for a turbine engine. The preform component includes a substrate. A meter section of a cooling aperture is formed in the substrate. An internal coating is applied onto a surface of the meter section. An external coating is applied over the substrate. A diffuser section of the cooling aperture is formed in the external coating and the substrate to provide the cooling aperture.