A resonant sound absorbing device of a gas turbine combustor includes a plurality of resonance chambers independently disposed side by side in an axial direction of the gas turbine combustor so as to communicate with a gas passage of the gas turbine combustor via acoustic holes. The plurality of resonance chambers include n related resonance chambers each satisfying: $\begin{array}{cc}0.9\times {.Math.}_{i=1}^n\frac{F_i}{n}\le F_i\le 1.1\times {.Math.}_{i=1}^n\frac{F_i}{n}& \left(A\right)\end{array}$
where n is an integer of 2 or more, and Fi is a peak frequency corresponding to a maximum sound absorbing ratio of the ith related resonance chamber of the n related resonance chambers.

A reheat assembly for gas turbine engine including a jetpipe casing having a reheat core section configured to flow air from inlet to outlet; and reheat bypass section configured to bypass air from inlet to outlet, wherein the reheat core section and the reheat bypass section are radially separated by support duct within the jetpipe casing; reheat arrangement including a radially extending flameholder and a core fuel injection port, wherein: the flameholder, mounted to the jetpipe casing, extends through the reheat bypass section and partly into the reheat core section; the flameholder is configured to form a wake-stabilised region within the core flow of air and the bypass flow of air downstream of the flameholder; and the core fuel injection port is: circumferentially aligned with the flameholder upstream of the wake-stabilised region, and configured to discharge fuel into the reheat core section for mixing with the core flow of air.

A reheat assembly for a gas turbine engine includes: a jetpipe casing including: a reheat core section configured to convey a core flow of air from a reheat core inlet to a reheat core outlet; and a reheat bypass section configured to convey a bypass flow of air from a reheat bypass inlet to a reheat bypass outlet radially outward of the reheat core section, wherein the reheat core and reheat bypass sections are radially separated at the reheat core and reheat bypass inlets by a support duct within the jetpipe casing; a reheat arrangement including a radially extending flameholder and a plurality of fuel injection ports including a plurality of core fuel injection ports, wherein: the flameholder is configured to promote a formation of a core flow wake-stabilised region within the core flow of air downstream of the flameholder; and each of the plurality of core fuel injection ports are: circumferentially aligned with the flameholder upstream of the core flow wake-stabilised region, configured to discharge a respective flow of fuel into the reheat core section for mixing with the core flow of air, and offset with respect to one another along a radial direction of the jetpipe casing.

A combustor nozzle includes at least one cluster composed of a plurality of tubes through which air and fuel flow, the cluster including a main tube through which the air and fuel flows, and a plurality of sub-tubes disposed to surround the main tube, wherein a diameter of the main tube and a diameter of at least one of the plurality of sub-tubes are different.

A combustor has an axial fuel staging system to allow a fuel-air mixture to be injected from two axially spaced stages using an injector for injection of a secondary fuel-air mixture. The combustor includes a liner defining a combustion chamber; a transition piece coupled to a rear end of the liner; a flow sleeve defining an annular channel by surrounding the liner and the transition piece; and at least one injector disposed on a circumferential position of the flow sleeve to inject a mixture of fuel and air into the combustion chamber. Each of the at least one injector includes an injection pipe extending radially and passing through both the flow sleeve and either of the liner and the transition piece; a plate coupled to the injection pipe; and a plurality of mixing passages formed through the plate. The combustor improves fuel-air mixing and prevents flash back.

The present application provides a variable frequency Helmholtz damper system for use with a combustor of a gas turbine engine. The variable frequency Helmholtz damper system may include one or more Helmholtz dampers and a purge medium temperature control unit for providing a flow of purge medium to the Helmholtz dampers. The purge medium temperature control unit may be in communication with a temperature control fluid flow such that the purge medium temperature control unit may vary the temperature of the flow of purge medium.

A rotary machine includes at least one burner including a front panel having a front side and an opposing back side. The acoustic damper includes at least one wall, at least one cooling air inlet, at least one outlet, and at least one baffle. The wall extends from the back side of the front panel and defines a dampening chamber. The cooling air inlet is defined within the back side of the front panel and is configured to channel a flow of cooling air into the dampening chamber. The outlet is defined within the back side of the front panel and is configured to channel the flow of cooling air out of the dampening chamber. The baffle extends from the back side of the front panel and is configured to reduce a velocity of the flow of cooling air within the dampening chamber.

A combustor liner has an outer liner and an inner liner that define a combustion chamber therebetween, the combustion chamber having a dilution zone. The outer liner and the inner liner each have a converging-diverging section extending into the dilution zone of the combustion chamber and form a throat between them. A bridge member extends across the converging-diverging sections to form a damper cavity therebetween. Each of the converging-diverging sections includes at least one dilution opening defined through the respective converging-diverging section at the throat, and includes a damper inlet feed member on a downstream portion of the converging-diverging section, so that an acoustic damper is defined by the converting-diverging section, the bridge, and the damper inlet feed. The acoustic damper dampens acoustic characteristics of the combustor.

The systems and methods described herein relate to a dome of a gas turbine assembly configured to suppress pressure pulsations. The systems and methods provide a dome having an aperture configured to surround an injector assembly of a combustor. The dome having a front panel extending radially from the aperture. The systems and methods couple a first cavity to the front panel. The first cavity includes a series of ducts. A first duct of the series of ducts is configured to receive airflow into the first cavity from a compressor and a second set of ducts of the series of ducts and a third duct of the series of ducts are configured to direct airflow to the combustor from the first cavity, wherein the third duct has a larger diameter than the second set of ducts.

An annular gas turbine combustor for use in an aircraft includes a plurality of resonators having respective resonance chambers and arranged in a fuel injector arrangement space so as to be lined up with fuel injectors in a radial direction of the fuel injector. A bulkhead configured to separate a combustion chamber and the fuel injector arrangement space from each other includes a plurality of openings each of which is arranged between each resonance chamber and the combustion chamber. The resonance chambers communicate with the combustion chamber through the openings. At least part of each of the resonance chambers is arranged at a radially inner side of a largest outer diameter portion of an outer liner when viewed in an axial direction of the fuel injector.