A liner associated with an engine of an aircraft is described. The liner includes a panel and an array of projections configured to enhance a cooling of the panel and distributed on at least part of a first side of the panel that corresponds to a cold side of the panel.

Components for gas turbine engines are provided. The components include a hybrid skin core cooling cavity defined by a cold wall and a hot wall, wherein the hot wall is exposed to an exterior environment of the component, a hybrid resupply hole formed in the cold wall and fluidly connecting a cold cavity and the hybrid skin core cooling cavity, and a resupply cover located on the cold wall and within the hybrid skin core cooling cavity and positioned relative to the hybrid resupply hole to shield resupply air injected from the cold cavity into the hybrid skin core cooling cavity to minimize losses as the resupply air mixes with air flowing within the hybrid skin core cooling cavity.

A system effective for dual utilization of cooling fluid in a gas turbine engine is provided. A cooling annulus is subject to a hot-temperature combustion flow received from a combustor basket and includes a liner including a feed channel to receive cooling fluid. A feed manifold is in fluid communication with feed channel to feed cooling fluid to a plurality of conduits in fluid communication with a plurality of exit orifices that is in fluid communication with a plurality of resonators. A distributor manifold includes a plurality of manifold sectors in fluid communication with a plurality of conduits arranged to convey cooling fluid. Some of the plurality of resonators operates with different amounts of cooling fluid. A group of the plurality of exit orifices is configured to supply an amount of cooling fluid appropriate for a resonator in fluid communication with the group of the plurality of exit orifices.

A combustion chamber assembly for an engine includes a wall element fixed to a combustion chamber structure and a chamber between the wall element and the structure, the chamber being supplied with air through impingement-cooling openings in the structure and connected to the combustion space by film-cooling openings in the wall element. Two cooling-air holes formed in the structure generate a cooling-air flow toward the combustion space and past the wall element. The wall element has a flow guide device for generating at least one scavenging-air flow directed between the two cooling-air holes. A sum of flow cross sections of film-cooling openings in the wall element and of the flow guide device yields a larger area than a sum of flow cross sections of all wall element impingement-cooling openings via which air is guided through the structure into the chamber and to the rear side of the wall element.

In a method for protecting heat exchanger pipe in boiler systems having at least one heat exchanger pipe, which is surrounded by a ceramic component, into which flue gas is directed from at least two opposite sides, gas is introduced between the heat exchanger pipe and the ceramic component.

The invention refers to a transition piece of a combustor of a gas turbine including an impingement cooling zone, a sequential disposed liner having at least one cooling arrangement and a closing plate with respect to the sequential disposed liner. The sequential disposed liner has a cooling channel structure. The cooling channel structure forms a closed loop cooling scheme or a quasi-closed loop cooling scheme. The cooling channel structure is operatively connected to a cooling medium to cool at least one part of the sequential disposed liner.

A multi-walled structure is provided for a turbine engine. This structure includes a shell with a textured first surface, and a heat shield with a second surface. The heat shield is attached to the shell. The first and the second surfaces vertically define a cooling cavity between the shell and the heat shield. The cooling cavity fluidly couples a plurality of cooling apertures defined in the shell with a plurality of cooling apertures defined in the heat shield.

A combustion chamber inside a combustion box has disposed therein a burner positioning frame having: a rear frame part positioned on a rear side of a burner-disposed portion; and a burner positioning frame having side frame part positioned on each laterally outside of the burner-disposed portion. The rear frame part has formed therein a laterally elongated drawn part which is dented rearward. The secondary air flowing from the distribution holes positioned in such a portion of the partition plate as is closer to the rear end of the partition plate is restrained by the drawn part from flowing upward from between the partition plate and a lower edge of the rear frame part through a clearance between the rear frame part and a rear plate part of the combustion box.

An air swirler arrangement comprises a coaxial arrangement of an inner and an outer air swirler passage. Each air swirler passage comprises a radial flow swirler. Air swirler arrangement comprises a coaxial arrangement of first, second and third members. Second member has radially extending upstream portion spaced axially from first member and a convergent portion. Third member has a radially extending upstream portion spaced axially from the upstream portion of second member and a radially inner surface having convergent and divergent downstream portions and a radially outer surface having a divergent downstream portion. First, second and third members the vanes of the radial flow swirlers is a monolithic structure. Plurality of circumferentially spaced passages are provided within the third member and each passage has an inlet in the surface and an outlet arranged to direct fluid onto the divergent portion of the surface or the surface of the third member.

A cooling device includes a hot-part cooling system configured to guide air to a hot part of a gas turbine. The hot-part cooling system includes a compressor capable of operating independently of the gas turbine, and is configured to extract air inside a casing of the gas turbine to pressurize the air with the compressor and guide the air to the hot part of the gas turbine. The cooling device further includes a rotor cooling system configured to extract the air inside the casing of the gas turbine and guide the air to a turbine rotor, and a connecting system configured to guide the air pressurized by the compressor to the rotor cooling system while a fuel supply to the gas turbine is stopped.