A gas turbine includes a rotor that is rotatable about an axis, a casing configured to cover the rotor in a circumferential direction and having an annular space therein, a compressor configured to generate a high pressure of compressed air obtained by compressing external air and send the compressed air into the casing, a plurality of combustors disposed in the casing at equal intervals in the circumferential direction of the rotor and configured to combust the compressed air and fuel taken in from the casing to generate a combustion gas, a turbine driven by the combustion gas, and an air introduction passage defined by a partition plate configured to divide the space in the casing in the circumferential direction of the rotor and an inner circumferential surface of the casing and configured to introduce the compressed air in the casing into the combustor.

A method of reducing noise from a combustor of a gas turbine engine includes the steps of establishing a maximum noise limit that may be for a particular frequency range. A primary fuel flow percentage, which may be emitted from a fuel nozzle arrangement having various groupings of simplex and duplex nozzles, is then established. An immersion depth measured between an aft rim of a swirler and a distal tip of the fuel nozzles may then be reduced thereby reducing the noise amplitude.

A structure for damping at a fluid manifold assembly for an engine is generally provided. The fluid manifold assembly includes a first walled conduit defining a first fluid passage therewithin. A flow of fluid defining a first frequency is permitted through the first fluid passage. A second walled conduit includes a pair of first portions each coupled to the first walled conduit. A second portion is coupled to the pair of first portions. A second fluid passage is defined through the first portion and the second portion in fluid communication with the first fluid passage. The flow of fluid is permitted through the second fluid passage at a second frequency approximately 180 degrees out of phase from the first frequency.

A noise attenuation panel for a structure within a propulsion system includes a first plurality of unit cells; and a second plurality of unit cells, the second plurality of unit cells merged within the first plurality of unit cells, the first plurality of unit cells including a first periodic structure having a first unit cell, a second unit cell, a third unit cell and a fourth unit cell, each of the first unit cell, the second unit cell, the third unit cell and the fourth unit cell including a central body interconnected via a plurality of lateral tubes extending from the central body, the first periodic structure forming a first lateral layer of unit cells.

The present disclosure is drawn to a gas turbine whereby hydrogen is used as a primary fuel to generate the energy needed to drive the rotation of the turbine via a set of hydrogen and air nozzles.

Stacked cooling assemblies and combustor bead ends are provided. A stacked cooling assembly includes an inlet plate defining an inlet to a coolant circuit, an outlet plate defining an outlet of the coolant circuit, and an intermediate plate disposed between the inlet plate and die outlet plate. The intermediate plate defines an intermediate cavity. A downstream surface of the inlet plate, an upstream surface of the outlet plate, and the intermediate cavity collectively define a connecting channel that fluidly couples the inlet to the outlet.

A gas turbine engine may include a combustion section having a fuel nozzle, a swirler, and a ferrule configured to mount and center the fuel nozzle with the swirler. The combustion section may further include a damper on a cold side of the combustion section. The damper may have an acoustic cavity, a damper neck, and a cavity feed hole. The damper may operate as Helmholtz cavity to absorb a hydrodynamic or acoustic instability present in a region within the swirler.

The present disclosure generally relates to a system with a gas turbine engine including a first combustor and a second combustor. The first combustor includes a first end cover with a first geometry and the second combustor includes a second end cover with a second geometry. The first geometry has one or more geometric differences relative to the second geometry.

A system includes a gas turbine engine having a first combustor and a second combustor. The first combustor includes a first fuel conduit having a first plurality of injectors. The first plurality of injectors are disposed in a first configuration within the first combustor along a first fuel path, and the first plurality of injectors are configured to route a fuel to a first combustion chamber. The system further includes a second combustor having a second fuel conduit having a second plurality of injectors. The second plurality of injectors are disposed in a second configuration within the second combustor along a second fuel path, and the second plurality of injectors are configured to route the fuel to a second combustion chamber. The second configuration has at least one difference relative to the first configuration.

A turbine engine having a compressor section, a combustion section and a turbine section in serial flow arrangement. The turbine engine further having a combustor, provided within the combustion section, defining a combustion chamber and having at least one fuel injector. The fuel injector having a fuel channel, a first set of fuel orifices, and a flow restrictor located within the fuel channel and having a second set of fuel orifices.