An assembly for a gas turbine engine according to an example of the present disclosure includes at least one platform having a main body extending between a first mate face and a second mate face to establish a gas path surface. The main body has an internal passage extending circumferentially between a first opening along the first mate face and a second opening along the second mate face relative to an assembly axis. A perimeter of the first mate face establishes a first area, a perimeter of the second opening establishes a second area, and the second area is greater than the first area. An airfoil section extends radially from the at least one platform.

A fuel tank heat dissipation system for fuel cell (FC) cooling is disclosed. in one example, at least one FC is in thermal communication with an intermediary heat exchanger. A fuel tank is also in fluid communication with the intermediary heat exchanger. A fluid is used to receive heat from the intermediary heat exchanger and flow along a first fluid path to the fuel tank. A nozzle is used to spray the fluid about an interior surface of the fuel tank, where the spray of the fluid about the interior of the fuel tank allows the fluid to dissipate the heat. A second fluid path from the fuel tank to the intermediary heat exchanger, the second fluid path to return the fluid that has dissipated the heat to the intermediary heat exchanger.

A combustor liner for a gas turbine engine, the combustor liner including a panel configured to at least partially define a combustion chamber. The combustor liner further includes a shell configured to mount to the panel and form a gap between the panel and the shell. The panel includes a stud and a plurality of a stand-off pins proximate to the stud defining a cavity therebetween. The shell includes a plurality of angled impingement holes located away from the cavity but extending through the shell at an orientation such that cooling air passing through the angled impingement holes is directed towards the cavity between adjacent stand-off pins and at an acute angle relative to the stud.

An engine component for a turbine engine having a working airflow separated into a cooling airflow and a combustion airflow, the engine component comprising a wall defining an interior and having an outer surface over which flows the combustion airflow, the outer surface defining a first side and a second side. The engine component further comprising at least one cooling conduit provided in the interior and having conduit sidewalls and a set of cooling passages formed in the wall and fluidly coupling the at least one cooling conduit to the outer surface, at least one of the cooling passages in the set comprising a primary cooling passage portion and a secondary cooling passage portion. A diffusion slot located in the primary cooling passage portion and an impingement zone fluidly coupled to the diffusion slot.

An airfoil includes an airfoil section that has an airfoil wall that defines leading and trailing ends and first and second sides that join the leading and trailing ends. The first and second sides span in a longitudinal direction between first and second ends. The airfoil wall circumscribes an internal core cavity. A platform is attached with the first end of the airfoil wall. The platform includes an opening that opens into the internal core cavity and a land region that circumscribes the opening. A baffle is formed of a tube and an attachment portion. The tube extends in the internal core cavity and the attachment portion has a flange ring that is affixed to the platform at the land region.

Various embodiments include a heat transfer device, a turbomachine casing and a related storage medium. In some cases, the device includes: a body having an outer surface and an inner cavity within the outer surface; at least one aperture extending through the body, the at least one aperture positioned to direct fluid from the inner cavity through the body to the outer surface; a first lip proximate a first end of the body, and a second lip proximate a second end of the body, the first lip and the second lip each extending radially outward from the outer surface relative to a direction of flow of the fluid through the inner cavity; and a plug coupled with the body, the plug for obstructing an end of the inner cavity, the plug positioned to redirect flow of the fluid from a first direction to a second, distinct direction.

A turbine assembly in a turbine section of an aircraft engine includes a rotor with blades having blade tips, and a turbine housing radially surrounding the blades. A distance between an inner surface of the housing and the blade tips defines a tip clearance gap. A passive cooling system for optimizing the tip clearance gap includes a cooling airflow passage located radially outward from, and in heat-transfer with, the turbine housing. The cooling airflow passage has an inlet opening located upstream of the rotor and an exit opening located downstream of the rotor. The inlet opening provides air flow into the cooling airflow passage. The exit opening provides air flow communication between the cooling airflow passage and a main gaspath of the turbine section. A flow of cooling air through the cooling airflow passage is induced, to cool the housing.

Provided is a gas turbine vane and blade assembly in which lattice structures are installed between an impingement plate and an effusion plate. The gas turbine vane and blade assembly is capable of enhancing cooling efficiency in an impingement/effusion cooling technique.

In addition, the gas turbine vane and blade assembly can be manufactured using an additive manufacturing technique, and the lattice structures are capable of replacing supports that are used during an additive manufacturing process, and improving not only structural rigidity and stability but also cooling performance.

A gas turbine nozzle assembly of a gas turbine is provided. The turbine nozzle assembly may include a turbine nozzle extending from an inner diameter to an outer diameter and having an airfoil-shaped cross section having a leading edge and a trailing edge, and a pressure side and a suction side each of which extends from the leading edge to the trailing edge, wherein the turbine nozzle may include a hollow airfoil including a plurality of cavities positioned in the airfoil, an insert positioned in one or more of the plurality of cavities of the hollow airfoil, a plurality of cover plates, at least one of which is positioned at one of the inner diameter and at the outer diameter, and a plurality of impingement pans, at least one of which is positioned at one of the inner diameter and at the outer diameter.

A cooling device (11) for a turbine of a turbomachine extending along an axis includes at least one radially inner metal sheet (14) and one radially outer metal sheet (15) that are joined to one another and delimit, between them, cooling air circulation channels (17) extending circumferentially from a connection region (16). Each channel (17) includes at least one air inlet and air ejection orifices (19), that are designed to be oriented toward a region to be cooled. The cooling device also includes at least one cooling duct (21) intended for the circulation of cooling air, the duct (21) located radially outside said metal sheets (14, 15) and close to or in contact with the metal sheets so as to cool said metal sheets using the cooling air circulating in the duct (21), the cooling duct (21) extending axially and arranged toward the circumferential end regions of the channels (17).