An additively manufactured impingement structure for a component is provided. The control structure includes an outer wall, an inner wall, and an impingement wall positioned between the outer wall and the inner wall. A fluid distribution passageway is defined between the inner wall and the impingement wall and an impingement gap is defined between the impingement wall and the outer wall. A plurality of impingement holes are defined in the impingement wall to provide fluid communication between the fluid distribution passageway and the impingement gap. A flow of cooling or heating fluid may be supplied to the fluid distribution passageway which distributes the flow and impinges it through the impingement holes onto the outer wall to cool or heat the outer wall, respectively.

A method for manufacturing a blade or vane assembly having at least one hollow airfoil (10) for a gas turbine, profile sections (A, B, C) of this airfoil being configured (S20) on the basis of a predetermined desired torsion of the airfoil, and the airfoil being manufactured (S30) on the basis of the configured profile sections using an additive manufacturing process (S30).

Disclosed is a blade for a turbomachine, comprising an outer shell and an inner core which is at least partially enclosed by the outer shell and has a higher porosity than the outer shell. The outer shell is formed by a ceramic body or a body made of a ceramic matrix composite material, and the inner core is formed by a fiber-reinforced ceramic or a fiber-reinforced ceramic matrix composite material.

A process for additively controlled surface features of a gas turbine engine casing. The process comprises forming the casing having an inner surface and an outer surface opposite the inner surface; forming a surface feature on the casing proximate the inner surface, wherein the surface feature comprises a structure on the inner surface configured to align or misalign with respect to a flow direction of a working fluid in a flow path of the casing.

A monopropellant thruster according to an exemplary aspect of the present disclosure includes, among other things, a first part having a catalyst bed, a thrust chamber, and a nozzle. The first part is integrally formed via a single additive manufacturing process. The thruster further includes a second part, which is a closeout. A method is also disclosed.

Aspects of the disclosure are directed to treating a substrate, the substrate including at least one of a refractory metal or a ceramic material, and depositing a media onto the treated substrate to generate a casting core. Embodiments include a fixture, a substrate located on the fixture, the substrate including at least one of a refractory metal or a ceramic material, and a delivery head that deposits media onto the substrate to generate a casting core. Aspects are directed to a core configured for casting a component, the core comprising: a substrate that includes at least one of a refractory metal or a ceramic material, and media deposited on the substrate, the media having a dimension within a range of between 0.5 and 100 micrometers.

A component for at least one of a combustion section or a turbine section of a gas turbine engine is provided. The combustion section and turbine section of the gas turbine engine at least partially define a core air flowpath, and the component includes a wall. The wall, in turn, includes a hot side and an opposite cold side. The hot side is exposed to and at least partially defines the core air flowpath when the component is installed in the gas turbine engine. The wall is manufactured to include surface contouring on the cold side of the wall to structurally accommodate a thermal management feature of the wall.

A component for a gas turbine engine includes a hot side wall, a plurality of connection walls, and a cold side wall. The hot side wall is exposed to a core air flowpath defined by the gas turbine engine. The cold side wall is spaced from the hot side wall and rigidly connected to the hot side wall through the plurality of connection walls. The hot side wall, connection walls, and cold side wall together define a cooling air cavity. The cold side wall defines a thermal stress relief slot for at least partially accommodating a relative thermal expansion between the hot side wall and the cold side wall to reduce an amount of thermal stress within the component during operation of the component.

A gas turbine engine includes a compressor section, combustion section, and turbine section. The turbine section includes a turbine component stage, the turbine component stage including a plurality of turbine components together including a flowpath surface along a circumferential direction of the gas turbine engine. The flowpath surface defines in part a core air flowpath of the gas turbine engine and further defines a contour along the circumferential direction. The contour repeats less frequently than once per turbine component to accommodate a hot gas streak through the turbine section.

A hybrid rocket motor includes a solid fuel element, and an oxidizer tank containing an oxidizer. The solid fuel element adjoins and at least partially defines a combustion chamber in which the solid fuel and the oxidizer are burned, to produce thrust from the hybrid rocket motor. The oxidizer tank is at least partially within the combustion chamber, and the entire oxidizer tank may be within the combustion chamber. The oxidizer tank may be protected by an insulating material, which may also serve as a structural material that contains the pressure of the oxidizer. The insulating material and the fuel material may both be polymer-based materials, although they may be different materials having different characteristics, for example including different additives to the same polymer material. The fuel element and the oxidizer tank may be made by additive manufacturing processes, for example by adding different materials in different locations.