Patent Description:
Combustion systems for gas turbine engines are generally tasked to reduce a plurality of emissions, such as carbon monoxide (CO), carbon dioxide (CO<NUM>), unburned hydrocarbons (UHC), smoke, and oxides of nitrogen (NOx), while also reducing combustion dynamics, such as pressure oscillations resulting in undesired vibrations and acoustics that deteriorate engine hardware, performance, and operability. At the same time, combustor assemblies must operate to desired overall fuel-air ratios to produce a desired output energy for the gas turbine engine. Still further, the combustor assembly must remain relatively simple for aero, industrial, or marine purposes.

However, reducing emissions, attenuating combustion dynamics, maintaining simplicity, while producing requisite energy output generally include conflicting design criteria. One known solution for reducing emissions includes staged combustion, in which the plurality of fuel nozzles of the combustor assembly defines several "on" fuel nozzles and several "off" fuel nozzles at various engine operating conditions, such as ignition/re-light, sub-idle, idle, and generally low power conditions. Although staged combustion provides some reduction in emissions such as CO and UHC, selective fuel injection to sectors of on/off fuel nozzles results in attenuation of the combustion process at outer edges of individual combustion zones created by the alternate sections of on/off fuel nozzles. As such, combustion efficiency is lowered, resulting in the formation of emissions elsewhere in the combustion process such as UHC, CO, or both. Still further, such arrangements may suffer from increased combustion dynamics. Furthermore, the on/off arrangement of fuel nozzles results in greater annular or circumferential temperature variations (i.e., hot spots) that adversely affect turbine efficiency. Still further, although such arrangements provide some lean blow-out (LBO) benefits at low power, known arrangements may not provide reductions in combustion dynamics, as well as reductions in emissions and mitigated hot spots at high power outputs. <CIT> discloses a method for operating a turbine engine having a plurality of fuel injectors arranged circumferentially in a combustor, with each fuel injector having a main fuel supply and a pilot fuel supply, the method creating a circumferential thermal gradient in the combustor.

Therefore, there is a need for a combustion system that provides attenuation of combustion dynamics, reduced emissions, and decreased variations in circumferential temperature profile.

The present disclosure is directed to a method according to claim <NUM>.

References to rotational speed of a shaft, rotor, blades, etc. includes mechanical speed and corrected speed, such as based on one or more of an inlet temperature, ambient temperature, or a proximate temperature of a fluid (e.g., air) within a gas path, unless otherwise specified.

Methods and structures for attenuating combustion dynamics are generally provided that may attenuate combustion dynamics, mitigate emissions, improve lean blowout margin, and attenuate circumferential temperature variations (e.g., hot spots). The methods include varying the flow rate of a first supply of fuel through at least two first fuel nozzles relative to a second supply of fuel through at least two second fuel nozzle to change the local stoichiometry and flame structure of the alternating arrangement of the first and second fuel nozzles. The alternating circumferential arrangement of the at least two first fuel nozzles and the at least two second fuel nozzles, such as to define an alternating arrangement of richer burning and leaner burning fuel nozzles, provides circumferential temperature non-uniformity while maintaining overall fuel-air ratio at the combustion chamber exit. As such, the non-uniform flame structure from the alternating circumferential arrangement of the first and second fuel nozzles suppresses combustion dynamics at part-power operating conditions of the gas turbine engine. Furthermore, temperature non-uniformities are then suppressed as the combustion gases flow downstream, such as through a dilution zone of the combustion chamber, thereby mitigating temperature non-uniformity at a turbine section, and associated adverse effects.

The alternating circumferential arrangement of richer burning and leaner burning fuel nozzles moves the fuel-air ratio of each fuel nozzle away from a critical stoichiometry with peak fuel nozzle swirler or mixer combustion dynamics. Furthermore, the alternating arrangement of richer burning and leaner burning fuel nozzles reduces NOx emissions by moving each fuel nozzle away from the stoichiometry producing maximum oxides of nitrogen. Still further, the aforementioned arrangement may further improve lean blow-out margin, thereby improving combustion stability and engine operability. Furthermore, the alternating arrangement of richer burning and leaner burning fuel nozzles maintains a desired overall fuel-air ratio of the combustion process while mitigating combustion dynamics.

Referring now to the drawings, <FIG> is a schematic partially cross-sectioned side view of an exemplary high bypass turbofan engine <NUM> herein referred to as "engine <NUM>" as may incorporate various embodiments of the present disclosure. Although further described below with reference to a turbofan engine, the present disclosure is also applicable to turbo machinery in general, including turbojet, turboprop, and turboshaft gas turbine engines, including marine and industrial turbine engines and auxiliary power units. As shown in <FIG>, the engine <NUM> has an axial or longitudinal centerline axis <NUM> that extends there through for reference purposes. The engine <NUM> defines a longitudinal direction L and an upstream end <NUM> and a downstream end <NUM> along the longitudinal direction L. The upstream end <NUM> generally corresponds to an end of the engine <NUM> along the longitudinal direction L from which air enters the engine <NUM> and the downstream end <NUM> generally corresponds to an end at which air exits the engine <NUM>, generally opposite of the upstream end <NUM> along the longitudinal direction L. In general, the engine <NUM> may include a fan assembly <NUM> and a core engine <NUM> disposed downstream from the fan assembly <NUM>.

The core engine <NUM> may generally include a substantially tubular outer casing <NUM> that defines an annular inlet <NUM>. The outer casing <NUM> encases or at least partially forms, in serial flow relationship, a compressor section having a booster or low pressure (LP) compressor <NUM>, a high pressure (HP) compressor <NUM>, a combustion system <NUM>, a turbine section including a high pressure (HP) turbine <NUM>, a low pressure (LP) turbine <NUM> and a jet exhaust nozzle section <NUM>. A high pressure (HP) rotor shaft <NUM> drivingly connects the HP turbine <NUM> to the HP compressor <NUM>. A low pressure (LP) rotor shaft <NUM> drivingly connects the LP turbine <NUM> to the LP compressor <NUM>. The LP rotor shaft <NUM> may also be connected to a fan shaft <NUM> of the fan assembly <NUM>. In particular embodiments, as shown in <FIG>, the LP rotor shaft <NUM> may be connected to the fan shaft <NUM> by way of a reduction gear <NUM> such as in an indirect-drive or geared-drive configuration. In other embodiments, the engine <NUM> may further include an intermediate pressure compressor and turbine rotatable with an intermediate pressure shaft altogether defining a three-spool gas turbine engine.

As shown in <FIG>, the fan assembly <NUM> includes a plurality of fan blades <NUM> that are coupled to and that extend radially outwardly from the fan shaft <NUM>. An annular fan casing or nacelle <NUM> circumferentially surrounds the fan assembly <NUM> and/or at least a portion of the core engine <NUM>. In one embodiment, the nacelle <NUM> may be supported relative to the core engine <NUM> by a plurality of circumferentially-spaced outlet guide vanes or struts <NUM>. Moreover, at least a portion of the nacelle <NUM> may extend over an outer portion of the core engine <NUM> so as to define a bypass airflow passage <NUM> therebetween.

<FIG> is a cross sectional side view of an exemplary combustion system <NUM> of the core engine <NUM> as shown in <FIG>. As shown in <FIG>, the combustion system <NUM> may generally include an annular type combustor <NUM> having an annular inner liner <NUM>, an annular outer liner <NUM> and a bulkhead <NUM> that extends radially between upstream ends <NUM>, <NUM> of the inner liner <NUM> and the outer liner <NUM> respectively. In other embodiments of the combustion system <NUM>, the combustion assembly <NUM> may be a can-annular type. The combustor <NUM> further includes a dome assembly <NUM> extended radially between the inner liner <NUM> and the outer liner <NUM> downstream of the bulkhead <NUM>. As shown in <FIG>, the inner liner <NUM> is radially spaced from the outer liner <NUM> with respect to engine longitudinal centerline <NUM> (<FIG>) and defines a generally annular combustion chamber <NUM> therebetween. In particular embodiments, the inner liner <NUM>, the outer liner <NUM>, and/or the dome assembly <NUM> may be at least partially or entirely formed from metal alloys or ceramic matrix composite (CMC) materials.

As shown in <FIG>, the inner liner <NUM> and the outer liner <NUM> may be encased within an outer casing <NUM>. An outer flow passage <NUM> of a diffuser cavity or pressure plenum <NUM> may be defined around the inner liner <NUM> and/or the outer liner <NUM>. The inner liner <NUM> and the outer liner <NUM> may extend from the bulkhead <NUM> towards a turbine nozzle or inlet <NUM> to the HP turbine <NUM> (<FIG>), thus at least partially defining a hot gas path between the combustor assembly <NUM> and the HP turbine <NUM>. A fuel nozzle assembly <NUM> may extend at least partially through the bulkhead <NUM> and dome assembly <NUM> to provide a fuel-air mixture <NUM> to the combustion chamber <NUM>. In various embodiments, the bulkhead <NUM> includes a fuel-air mixing structure attached thereto (e.g., a swirler or mixer assembly).

During operation of the engine <NUM>, as shown in <FIG> and <FIG> collectively, a volume of air as indicated schematically by arrows <NUM> enters the engine <NUM> through an associated inlet <NUM> of the nacelle <NUM> and/or fan assembly <NUM>. As the air <NUM> passes across the fan blades <NUM> a portion of the air as indicated schematically by arrows <NUM> is directed or routed into the bypass airflow passage <NUM> while another portion of the air as indicated schematically by arrow <NUM> is directed or routed into the LP compressor <NUM>. Air <NUM> is progressively compressed as it flows through the LP and HP compressors <NUM>, <NUM> towards the combustion system <NUM>. As shown in <FIG>, the now compressed air as indicated schematically by arrows <NUM> flows into a diffuser cavity or pressure plenum <NUM> of the combustion system <NUM>. The pressure plenum <NUM> generally surrounds the inner liner <NUM> and the outer liner <NUM>, and generally upstream of the combustion chamber <NUM>.

The compressed air <NUM> pressurizes the pressure plenum <NUM>. A first portion of the of the compressed air <NUM>, as indicated schematically by arrows <NUM>(a) flows from the pressure plenum <NUM> into the combustion chamber <NUM> through the fuel nozzle <NUM> (e.g., across a vane structure configured to promote fuel-air mixing) where it is mixed with the fuel <NUM> and burned, thus generating combustion gases, as indicated schematically by arrows <NUM>, within the combustor <NUM>. Typically, the LP and HP compressors <NUM>, <NUM> provide more compressed air to the pressure plenum <NUM> than is needed for combustion. Therefore, a second portion of the compressed air <NUM> as indicated schematically by arrows <NUM>(b) may be used for various purposes other than combustion. For example, as shown in <FIG>, compressed air <NUM>(b) may be routed into the outer flow passage <NUM> to provide cooling to the inner and outer liners <NUM>, <NUM>.

Referring back to <FIG> and <FIG> collectively, the combustion gases <NUM> generated in the combustion chamber <NUM> flow from the combustor assembly <NUM> into the HP turbine <NUM>, thus causing the HP rotor shaft <NUM> to rotate, thereby supporting operation of the HP compressor <NUM>. As shown in <FIG>, the combustion gases <NUM> are then routed through the LP turbine <NUM>, thus causing the LP rotor shaft <NUM> to rotate, thereby supporting operation of the LP compressor <NUM> and/or rotation of the fan shaft <NUM>. The combustion gases <NUM> are then exhausted through the jet exhaust nozzle section <NUM> of the core engine <NUM> to provide propulsive thrust.

Referring now to the circumferential flowpath view generally provided in <FIG>, the combustion system <NUM> further defines the fuel nozzle assembly <NUM> as including the at least two first fuel nozzles <NUM> and at least two second fuel nozzles <NUM> in alternating circumferential arrangement (i.e., alternating along circumferential direction C around the longitudinal centerline axis <NUM>). The first fuel nozzles <NUM> and the second fuel nozzles <NUM> are each coupled to a fuel supply system <NUM> providing a first supply of fuel <NUM> to the first fuel nozzles <NUM> and a second supply of fuel <NUM> to the second fuel nozzles <NUM>.

The overall supply of fuel <NUM>, and the first supply of fuel <NUM> and second supply of fuel <NUM> therefrom, may be split based on a volumetric flow rate or a mass flow rate.

In the embodiment generally provided in <FIG>, the fuel supply system <NUM> receives an overall supply of fuel <NUM> and then splits or divides the overall supply of fuel <NUM> into the first supply of fuel <NUM> and the second supply of fuel <NUM>. The first supply of fuel <NUM> and the second supply of fuel <NUM> together account for the overall supply of fuel <NUM> delivered to the combustion chamber <NUM> for combustion purposes. For example, a desired overall fuel-air ratio at the combustion chamber <NUM> is based on the overall supply of fuel <NUM>, of which the sum of the first supply of fuel <NUM> and the second supply of fuel <NUM> together at least approximately equal the overall supply of fuel <NUM>).

Referring still to <FIG>, the fuel supply system <NUM> may include a first fuel manifold <NUM> coupled to each of the first fuel nozzles <NUM>, and further providing the first supply of fuel <NUM> to each first fuel nozzle <NUM>. The fuel supply system <NUM> may further include a second fuel manifold <NUM> coupled to each of the second fuel nozzles <NUM>, and further providing the second supply of fuel <NUM> to each second fuel nozzle <NUM>.

In other embodiments, the overall supply of fuel <NUM> is provided from the fuel supply system <NUM> to the fuel nozzle assembly <NUM> including each of the first fuel nozzle <NUM> and the second fuel nozzle <NUM>. Each first fuel nozzle <NUM> and second fuel nozzle <NUM> includes a valve or metering orifice that then limits the portion of the overall supply of fuel <NUM> that egresses the plurality of fuel nozzles <NUM> and mixes with air <NUM>(a) and releases into the combustion chamber <NUM> as the fuel-air mixture <NUM>.

In various embodiments, the fuel supply system <NUM>, including valves, metering orifices, flow restrictors, or other flow or pressure alternating devices, provides at least <NUM>% of the overall supply of fuel <NUM> as the first supply of fuel <NUM> egressing the first fuel nozzle <NUM> and mixing with the air <NUM>(a) as a first fuel-air mixture at the combustion chamber <NUM>. The remainder (i.e., the difference between the overall supply of fuel <NUM> and the first supply of fuel <NUM>) flows through the each of the second fuel nozzles <NUM> as the second supply of fuel <NUM>, thereby producing a second fuel-air mixture at the combustion chamber <NUM> different from the first fuel-air mixture.

The first fuel nozzle <NUM> defines a richer burning fuel nozzle and the second fuel nozzle <NUM> defines a leaner burning fuel nozzle. For example, each first fuel nozzle <NUM> may define a local fuel-air equivalence ratio as providing more fuel in the first fuel-air mixture than is required for complete combustion. As another example, each first fuel nozzle <NUM> may define a local fuel-air equivalence ratio greater than that of the second fuel nozzles <NUM>, in which each of the first fuel nozzle <NUM> and the second fuel nozzle <NUM> define an equivalence ratio greater than <NUM>.

At part-power conditions, such as from ignition or light-off to below maximum power (e.g., low power, medium power, or sub-idle, idle, cruise, approach, climb conditions, etc.), the fuel supply system <NUM> provides at least <NUM>% of the overall supply of fuel <NUM> to the first fuel nozzles <NUM>. For example, at ignition or low power conditions, approximately <NUM>% to <NUM>% of the overall supply of fuel <NUM> may egress through the first fuel nozzles <NUM> as the first fuel-air mixture and the remaining <NUM>% to <NUM>% may egress through the second fuel nozzle <NUM> as the second fuel-air mixture. In various embodiments, the fuel split may define <NUM>/<NUM> to the first fuel nozzle <NUM> versus the second fuel nozzle <NUM>; or <NUM>/<NUM> to the first fuel nozzle <NUM> versus the second fuel nozzle <NUM>; or <NUM>/<NUM>, or <NUM>/<NUM>, or <NUM>/<NUM>, etc. The alternating circumferential arrangement of the first fuel nozzle <NUM> and the second fuel nozzle <NUM> defining such fuel splits provides circumferential temperature non-uniformity in the combustion chamber <NUM> while maintaining an overall desired fuel-air ratio. The alternating circumferential arrangement providing temperature non-uniformity alters the local stoichiometry and flame structure at the first fuel nozzle <NUM> versus the second fuel nozzle <NUM> which thereby suppresses combustion dynamics at part-power operating conditions of the gas turbine engine. Additionally, the alternating arrangement of first and second fuel nozzles <NUM>, <NUM> de-couples heat release from combustion pressure fluctuations, thereby mitigating formation or propagation of combustion dynamics within the combustion chamber <NUM>.

Referring to <FIG>, the combustion chamber <NUM> may define a primary combustion zone <NUM> adjacent to an exit of the plurality of fuel nozzles <NUM> at which the fuel-air mixture <NUM> is initially ignited and burned. The primary combustion zone <NUM> may generally define a region within the combustion chamber <NUM> at which maximum temperatures of the combustion gases <NUM> are produced and a dilution zone <NUM>. In various embodiments, the primary combustion zone <NUM> may further include a secondary combustion zone, in which an upstream end of the primary combustion zone <NUM> provides an initial temperature rise that facilitates further combustion at the downstream secondary combustion zone.

Referring to <FIG>, within the primary combustion zone <NUM>, the alternating circumferential arrangement of the first fuel nozzle <NUM> and the second fuel nozzle <NUM> may produce circumferential variations in temperature. However, the circumferential variations in temperature are at least partially mitigated by the combination of the plurality of fuel nozzles <NUM> defining a richer burning fuel nozzle and a leaner burning fuel nozzle. Still further, as the combustion gases <NUM> flow downstream from the primary combustion zone <NUM> to the dilution zone <NUM>, circumferential variations in temperature are further attenuated, such as to mitigate or eliminate adverse effects to durability of the turbine section <NUM>. As such, the combination of the decreased initial gradient of the combination of richer burning and leaner burning fuel nozzles (e.g., the first fuel nozzle <NUM> and the second fuel nozzle <NUM>) further facilitates reduction or elimination of adverse magnitudes of circumferential temperature variations at the turbine section <NUM>.

At maximum power or full load conditions, the fuel supply system <NUM> provides an approximately <NUM>/<NUM> or approximately equal quantity or portion of the overall supply of fuel <NUM> to each of the first fuel nozzles <NUM> and second fuel nozzles <NUM>. As such, at maximum power operating conditions, the fuel supply system <NUM> mitigates formation of circumferentially non-uniform temperature profiles (e.g., hot spots) along through the combustion chamber <NUM> that may adversely affect durability of the turbine section <NUM>.

Referring now to <FIG>, an exemplary flowchart outlining steps of a method of attenuating combustion dynamics is generally provided (hereinafter, "method <NUM>"). The method <NUM> may be implemented on a combustion system, such as the combustion system <NUM> generally provided and described in regard to <FIG>. It should be appreciated that steps of the method <NUM> may be re-arranged, omitted, or altered within the scope of the present disclosure.

The method <NUM> includes at <NUM> flowing, via a compressor section, an overall supply of air (e.g., <NUM>) to the combustion system; at <NUM> flowing, via a fuel supply system (e.g., <NUM>), an overall flow of fuel (e.g., <NUM>) to the combustion system (e.g., <NUM>); at <NUM> flowing, to a first fuel nozzle (e.g., <NUM>) of the combustion system, a first supply of fuel (e.g., <NUM>) defining a richer fuel-air mixture at the first fuel nozzle; at <NUM> flowing, to a second fuel nozzle (e.g., <NUM>) of the combustion system, a second supply of fuel (e.g., <NUM>) defining a leaner fuel-air mixture at the second fuel nozzle; and at <NUM> igniting the richer fuel-air mixture and the leaner fuel-air mixture to produce an overall fuel-air ratio at a combustion chamber of the combustion system.

The method <NUM> further includes at <NUM> determining a desired overall fuel-air ratio at the combustion chamber based at least on the overall supply of air and the overall supply of fuel. For example, as described in regard to <FIG>, a desired overall fuel-air ratio at the combustion chamber <NUM> is defined based on an operating condition of the engine <NUM>. The operating condition is a function of one or more of a low rotor speed (e.g., N<NUM> or NL, or the rotational speed of the LP rotor <NUM>), a high rotor speed (e.g., N<NUM> or NH, or the rotational speed of the HP rotor <NUM>), an overall supply of fuel (e.g., Wftotal), an overall supply of air (e.g., Wa, or Wa3 or air flow rate at the combustion system <NUM>), pressure at the combustion system <NUM> (e.g., P<NUM>), temperature at the combustion system <NUM> (e.g., T<NUM>), or engine pressure ratio (EPR), or combinations thereof.

The method <NUM> may further include at <NUM> determining a fuel split to the first fuel nozzle and to the second fuel nozzle based on the overall flow of fuel that is further based at least on an operating condition of the engine, the overall supply of air, and the desired overall fuel-air ratio. For example, as described in regard to <FIG>, the fuel split may be provided at the fuel supply system <NUM> to define the first supply of fuel <NUM> and the second supply of fuel <NUM> from the overall supply of fuel <NUM>. The fuel split defines the first supply of fuel <NUM> at the first fuel nozzles <NUM> as at least <NUM>% of the overall supply of fuel <NUM>. The remainder is provided to the second fuel nozzles <NUM> as the second supply of fuel <NUM>.

As a result of the fuel split, as well as the approximately equal flows of air (e.g., air <NUM>(a)) through and across the plurality of fuel nozzles <NUM>, including the first fuel nozzle <NUM> and the second fuel nozzle <NUM>, the fuel split defines a first equivalence ratio of a first fuel-air mixture from the first fuel nozzle <NUM> different from a second equivalence ratio of a second fuel-air mixture from the second fuel nozzle <NUM> when the operating condition of the engine is less than the maximum power operating condition. For example, the first fuel nozzle <NUM> may define the first equivalence ratio corresponding to a richer burning while the second fuel nozzle <NUM> may define the second equivalence ratio corresponding to a leaner burning at the combustion chamber <NUM>. As previously described in regard to <FIG>, the alternating circumferential arrangement of the first fuel nozzle <NUM> and the second fuel nozzle <NUM>, such as defining an alternating circumferential arrangement of a richer burning fuel nozzle and a leaner burning fuel nozzle, produces dissimilar flame structures and local stoichiometries at each fuel nozzle <NUM>, <NUM> that mitigate combustion dynamics while also mitigating production of emissions.

As described in regard to <FIG>, the fuel split defines a quantity of the first supply of fuel <NUM> to the first fuel nozzle <NUM> as between approximately <NUM>% and <NUM>% of the overall supply of fuel <NUM>. The fuel split further defines a quantity of the second supply of fuel <NUM> to the second fuel nozzle <NUM> as a difference of the overall supply of fuel <NUM> from the first supply of fuel <NUM> to the first fuel nozzle <NUM>. At maximum or high power operating conditions (e.g., full load, or takeoff condition, etc.), the fuel split defines an approximately <NUM>/<NUM> split of the first supply of fuel <NUM> to the first fuel nozzle <NUM> and the second supply of fuel <NUM> to the second fuel nozzle <NUM>. Furthermore, at maximum or high power, the fuel split defines an approximately equal equivalence ratio at the first fuel nozzle <NUM> and the second fuel nozzle <NUM> when the operating condition of the engine.

At ignition/re-light, low power, or medium power conditions, or more generally, conditions below maximum or high power, the fuel split may define ratios between the first supply of fuel <NUM> to the first fuel nozzle <NUM> and the second supply of fuel <NUM> to the second fuel nozzle <NUM> as previously mentioned (e.g., <NUM>/<NUM>, <NUM>/<NUM>, <NUM>, <NUM>, <NUM>/<NUM>, <NUM>/<NUM>, etc.). In various embodiments, the fuel split generally approaches approximately <NUM>/<NUM> as the operating condition of the engine <NUM> increases toward maximum or high power.

At <NUM>, determining the fuel split is based on one or more of a lookup table, a function, or a curve. For example, the function, such as a transfer function, or one or more tables, functions, curves, or references stored at a computer-device including memory and a processor (e.g., a full authority digital engine controller or FADEC), may utilize one or more of a low rotor speed (e.g., N<NUM> or NL, or the rotational speed of the LP rotor <NUM>), a high rotor speed (e.g., N<NUM> or NH, or the rotational speed of the HP rotor <NUM>), an overall supply of fuel (e.g., Wftotal), an overall supply of air (e.g., Wa, or Wa3 or air flow rate at the combustion system <NUM>), pressure at the combustion system <NUM> (e.g., P<NUM>), temperature at the combustion system <NUM> (e.g., T<NUM>), or engine pressure ratio (EPR), or combinations thereof to determine the fuel split.

The method <NUM> may further include at <NUM> measuring, via one or more sensors, a frequency, amplitude, or both, or magnitude of changes of a pressure at the combustion chamber, and a frequency, amplitude, or both of vibrations at the combustion chamber. For example, the engine <NUM> may further include one or more sensors <NUM> (shown in <FIG>) measuring, monitoring, or calculating a pressure at the combustion chamber <NUM>. For example, the sensor <NUM> may sense a dynamic pressure resulting from the heat release produced by ignition of the fuel-air mixture <NUM> resulting in the combustion gases <NUM> in the combustion chamber <NUM>. The dynamic pressure may result in frequencies, amplitudes, and changes in magnitudes thereof, that indicates combustion dynamics, or attenuations or excitations thereof resulting in one or more acoustic modes. The sensors <NUM> may further measure, monitor, or calculate a frequency, amplitude, or both of vibrations at the combustion chamber.

As such, the method <NUM> may further include at <NUM> determining one or more acoustic modes at the combustion chamber; at <NUM> determining a fuel split to the first fuel nozzle and to the second fuel nozzle based on the one or more acoustic modes to be attenuated at the combustion chamber; and at <NUM> adjusting the fuel split based on the desired overall fuel-air ratio and one or more of a frequency, amplitude, or both, or magnitude of changes thereof of a pressure at the combustion chamber, and a frequency, amplitude, or both of vibrations at the combustion chamber.

At <NUM>, determining the fuel split is further based at least on a desired overall fuel-air ratio at the combustion chamber and the one or more acoustic modes at the combustion chamber to be attenuated.

The first fuel nozzle <NUM> and the second fuel nozzle <NUM> may be configured to provide different flow rates of air <NUM>(a) therethrough for mixing with the first supply of fuel <NUM> and the second supply of fuel <NUM>, respectively. For example, the first fuel nozzle <NUM> may define volumes, cross sectional areas, metering orifices, etc. that may restrict or provide a flow of air <NUM>(a) through the first fuel nozzle <NUM> different from the second fuel nozzle <NUM>. In other embodiments, the engine <NUM> may be configured to provide variable flows to the first fuel nozzle <NUM> and the second fuel nozzle <NUM> such as to define the first equivalence ratio and the second equivalence ratio, respectively, from each fuel nozzle <NUM>, <NUM>.

As such, and referring to <FIG> in addition to <FIG>, the method <NUM> may further include at <NUM> flowing, through a first fuel-air mixing flowpath of the first fuel nozzle, a first supply of air from the overall supply of air from the compressor section; at <NUM> mixing the first supply of air with the first supply of fuel within the first fuel-air mixing flowpath of the first fuel nozzle to produce a richer fuel-air mixture; at <NUM> flowing, through a second fuel-air mixing flowpath of the second fuel nozzle, a second supply of air from the overall supply of air from the compressor section; and at <NUM> mixing the second supply of air with the second supply of fuel within the second fuel-air mixing flowpath of the second fuel nozzle to produce a leaner fuel-air mixture. Still further, the method may further include at <NUM> flowing an approximately equal first supply of fuel and second supply of fuel to produce an approximately equal fuel-air mixture at each of the first fuel nozzle and the second fuel nozzle at a maximum power operating condition.

The methods <NUM> and structures <NUM> for attenuating combustion dynamics generally provided herein vary the flow rate of the first supply of fuel <NUM> through the first fuel nozzle <NUM> relative to the second supply of fuel <NUM> through the second fuel nozzle <NUM> to change the local stoichiometry and flame structure of the alternating arrangement of the first and second fuel nozzles <NUM>, <NUM>. The alternating circumferential arrangement of the first fuel nozzle <NUM> and the second fuel nozzle <NUM>, such as to define an alternating arrangement of richer burning and leaner burning fuel nozzles, provides circumferential temperature non-uniformity while maintaining overall fuel-air ratio at the combustion chamber <NUM> exit. For example, the circumferential temperature non-uniformity is maintained within a primary combustion zone <NUM> of the combustion chamber <NUM> adjacent to the exit (e.g., downstream end) of the plurality of fuel nozzles <NUM> including the first and second fuel nozzles <NUM>, <NUM>. As such, the non-uniform flame structure from the alternating circumferential arrangement of the first and second fuel nozzles <NUM>, <NUM> suppresses combustion dynamics at part-power operating conditions of the gas turbine engine (e.g., from sub-idle to under maximum power). Furthermore, temperature non-uniformities are then suppressed as the combustion gases <NUM> flow downstream (i.e., toward the downstream end <NUM>), such as through a dilution zone <NUM> of the combustion chamber <NUM>, thereby mitigating temperature non-uniformity and adverse effects thereof (e.g., circumferential hot spots adversely affecting the turbine section).

The alternating circumferential arrangement of richer burning and leaner burning fuel nozzles moves the fuel-air ratio of each fuel nozzle <NUM> away from a critical stoichiometry with peak fuel nozzle swirler or mixer combustion dynamics. Furthermore, the alternating arrangement of richer burning and leaner burning fuel nozzles reduces emissions of oxides of nitrogen by moving each fuel nozzle <NUM> away from the stoichiometry producing maximum oxides of nitrogen. Still further, the aforementioned arrangement may further improve lean blow-out margin, thereby improving combustion stability and engine operability. Furthermore, the alternating arrangement of richer burning and leaner burning fuel nozzles maintains a desired overall fuel-air ratio of the combustion process while mitigating combustion dynamics.

All or part of the combustion system <NUM> may be part of a single, unitary component and may be manufactured from any number of processes commonly known by one skilled in the art. These manufacturing processes include, but are not limited to, those referred to as "additive manufacturing" or "3D printing". Additionally, any number of casting, machining, welding, brazing, or sintering processes, or any combination thereof may be utilized to construct the combustion system <NUM>, including. Furthermore, the combustor assembly may constitute one or more individual components that are mechanically joined (e.g. by use of bolts, nuts, rivets, or screws, or welding or brazing processes, or combinations thereof) or are positioned in space to achieve a substantially similar geometric, aerodynamic, or thermodynamic results as if manufactured or assembled as one or more components. Non-limiting examples of suitable materials include high-strength steels, nickel and cobalt-based alloys, and/or metal or ceramic matrix composites, or combinations thereof.

Claim 1:
A method (<NUM>) of operating a combustion system to attenuate combustion dynamics, wherein the combustion system comprises a plurality of fuel nozzles (<NUM>), the plurality of fuel nozzles (<NUM>) comprising at least two first fuel nozzles and at least two second fuel nozzles; the method (<NUM>) comprising:
(<NUM>) flowing, via a compressor section, an overall supply of air to the combustion system, wherein at least a portion of the overall supply of air is provided through the plurality of fuel nozzles (<NUM>);
(<NUM>) flowing, via a fuel supply system, an overall flow of fuel to the combustion system;
(<NUM>) flowing, to the at least two first fuel nozzles of the combustion system, a first supply of fuel defining a richer fuel-air mixture at the first fuel nozzles;
(<NUM>) flowing, to the at least two second fuel nozzles of the combustion system, a second supply of fuel defining a leaner fuel-air mixture at the second fuel nozzles, where the at least two first fuel nozzles and the at least two second fuel nozzles are in alternating circumferential arrangement along a circumferential direction in the combustion system; and
(<NUM>) igniting the richer fuel-air mixture and the leaner fuel-air mixture to produce an overall fuel-air ratio at a combustion chamber of the combustion system,
characterised in that the method further comprises :
(<NUM>) determining a desired overall fuel-air ratio at the combustion chamber based at least on the overall supply of air and the overall flow of fuel, and:
(<NUM>) determining a fuel split to the at least two first fuel nozzles and to the at least two second fuel nozzles based on the overall flow of fuel that is further based at least on an operating condition of the engine, the overall supply of air, and the desired overall fuel-air ratio.