US20190056109A1

US20190056109A1 - Main fuel nozzle for combustion dynamics attenuation

Info

Publication number: US20190056109A1
Application number: US15/681,918
Authority: US
Inventors: Hejie Li; Duane Douglas Thomsen
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 2017-08-21
Filing date: 2017-08-21
Publication date: 2019-02-21

Abstract

The present disclosure is directed to a combustor assembly for a gas turbine engine. The combustor assembly includes a fuel nozzle comprising a centerbody extended along a lengthwise direction, wherein the fuel nozzle defines a nozzle centerline extended through the centerbody along the lengthwise direction. The centerbody defines a plurality of exit openings in circumferential arrangement relative to the nozzle centerline. The plurality of exit openings defines two or more locations different from one another on the centerbody along the lengthwise direction.

Description

FIELD

The present subject matter relates generally to gas turbine engine combustion assemblies.

BACKGROUND

Pressure oscillations generally occur in combustion sections of gas turbine engines resulting from the ignition of a fuel and air mixture within a combustion chamber. While nominal pressure oscillations are a byproduct of combustion, increased magnitudes of pressure oscillations may result from generally operating a combustion section at lean conditions, such as to reduce combustion emissions. Increased pressure oscillations may damage combustion sections and/or accelerate structural degradation of the combustion section in gas turbine engines, thereby resulting in engine failure or increased engine maintenance costs. As gas turbine engines are increasingly challenged to reduce emissions, structures for attenuating combustion gas pressure oscillations are needed to enable reductions in gas turbine engine emissions while maintaining or improving the structural life of combustion sections.

BRIEF DESCRIPTION

Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention.
The present disclosure is directed to a combustor assembly for a gas turbine engine. The combustor assembly defines a combustor centerline and a radial direction extended from the combustor centerline. The radial direction and combustor centerline together defining a reference plane. The combustor assembly includes a fuel nozzle including a centerbody extended along a lengthwise direction. The fuel nozzle defines a nozzle centerline extended through the centerbody along the lengthwise direction. The centerbody defines a plurality of exit openings in circumferential arrangement relative to the nozzle centerline. The plurality of exit openings defines two or more locations different from one another on the centerbody along the lengthwise direction.
In various embodiments, the two or more locations along the lengthwise direction and the circumferential arrangement of the plurality of exit openings define a waveform. In various embodiments, the waveform is sinusoidal, sawtooth, or box.
In still various embodiments, the plurality of exit openings defines two or more cross sectional areas. Each cross sectional area is different from one another. In one embodiment, the plurality of exit openings provides a fuel to a combustion chamber at two or more flow rates each corresponding to the two or more cross sectional areas of the plurality of exit openings.
In another embodiment, the plurality of exit openings defines a first opening and a second opening, the first opening defining a first lengthwise location and the second opening defining a second lengthwise location different from the first lengthwise location.
In one embodiment, the plurality of exit openings defines a first opening and a second opening, the first opening defining a first cross sectional area and the second opening defining a second cross sectional area different from the first cross sectional area.
In still another embodiment, a single exit opening of the plurality of exit openings is defined at each circumferential location along the centerbody from the nozzle centerline.
In still yet another embodiment, the plurality of exit openings are defined in asymmetric circumferential arrangement through the centerbody.
Another aspect of the present disclosure is directed to a gas turbine engine including a combustor assembly comprising a plurality of fuel nozzles disposed in circumferential arrangement around the axial centerline of the gas turbine engine. Each fuel nozzle includes a centerbody extended along a lengthwise direction, and wherein the fuel nozzle defines a nozzle centerline extended through the centerbody along the lengthwise direction. The centerbody defines a plurality of exit openings in circumferential arrangement relative to the nozzle centerline. The plurality of exit openings defines two or more cross sectional areas different from one another on the centerbody.
In one embodiment, the plurality of exit openings are defined along the same reference plane.
In various embodiments, the plurality of exit openings is defined along two or more locations different from one another along the lengthwise direction. In one embodiment, the plurality of exit openings provides a fuel to a combustion chamber at two or more flow rates each corresponding to the two or more cross sectional areas of the plurality of exit openings. In still various embodiments, the two or more locations along the lengthwise direction and the circumferential arrangement of the plurality of exit openings defines a waveform along the centerbody. In one embodiment, the waveform is sinusoidal, sawtooth, or box.
In another embodiment, the plurality of exit openings defines a first opening and a second opening, the first opening defining a first lengthwise location and the second opening defining a second lengthwise location different from the first lengthwise location.
In still another embodiment, a single exit opening of the plurality of exit openings is defined at each circumferential location along the centerbody from the nozzle centerline.
In still various embodiments, the plurality of fuel nozzles defines a first nozzle and a second nozzle. The first nozzle defines a plurality of exit openings, the plurality of exit openings defining a first opening defining a first cross sectional area and a second opening defining a second cross sectional area different from the first cross sectional area. The second nozzle defines the plurality of exit openings at a plurality of locations along the lengthwise direction. In one embodiment, the plurality of fuel nozzles is defined in alternating circumferential arrangement. In another embodiment, at least half of the plurality of fuel nozzles defines the first nozzle and the remainder of the plurality of fuel nozzles defines the second nozzle. At least two or more of the first fuel nozzle or the second fuel nozzle are in sequential circumferential arrangement.
These and other features, aspects and advantages of the present invention will become better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and, together with the description, serve to explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended drawings, in which:

FIG. 1 is a schematic cross-sectional view of an exemplary embodiment of a gas turbine engine;

FIG. 2 is a cross sectional side view of an exemplary embodiment of a combustor assembly of the gas turbine engine generally provided in FIG. 1;

FIG. 3 is a perspective view of an exemplary embodiment of a fuel nozzle of the combustor assembly generally provided in FIG. 2;

FIG. 4 is an axial view of an embodiment of the fuel nozzle generally provided in FIG. 3; and

FIG. 5 is an axial view of an embodiment of the fuel nozzle generally provided in FIG. 3.

Repeat use of reference characters in the present specification and drawings is intended to represent the same or analogous features or elements of the present invention.

DETAILED DESCRIPTION

Reference now will be made in detail to embodiments of the invention, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation of the invention, not limitation of the invention. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made in the present invention without departing from the scope or spirit of the invention. For instance, features illustrated or described as part of one embodiment can be used with another embodiment to yield a still further embodiment. Thus, it is intended that the present invention covers such modifications and variations as come within the scope of the appended claims and their equivalents.
As used herein, the terms “first”, “second”, and “third” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components.
The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows. The terms “upstream of” or “downstream of” generally refer to directions toward “upstream 99” or toward “downstream 98”, respectively, as provided in the figures.
Embodiments of a combustor assembly for a gas turbine engine including a fuel nozzle generally provided that may desirably alter the heat release characteristics of each fuel nozzle to mitigate undesired combustion dynamics. The fuel nozzle defines a plurality of exit openings egressing a main fuel flow to the combustion chamber. The plurality of exit openings are defined through a centerbody of the fuel nozzle at two or more locations or reference planes along the lengthwise direction of the fuel nozzle. The plurality of exit openings may further define two or more cross sectional areas. The fuel nozzle providing a main fuel flow is configured to provide a fuel-air mixture capable of engine operation up to a maximum or high power condition.
The combustor assembly including the embodiments of the fuel nozzle shown and described herein may attenuate pressure oscillations characterized by high pressure fluctuations that are sustained in a combustion chamber of a combustion section. Embodiments of the fuel nozzle may mitigate such pressure oscillations by altering the heat release characteristics of each flame from each fuel nozzle. Altering the heat release characteristics, such as flame structure, characteristic time, or both, for each fuel nozzle may then decouple heat release from pressure fluctuations, thereby mitigating undesired combustion dynamics.
Referring now to the drawings, FIG. 1 is a schematic partially cross-sectioned side view of an exemplary high by-pass turbofan engine 10 herein referred to as “engine 10” 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 propulsion systems and turbomachinery in general, including turbojet, turboprop, and turboshaft gas turbine engines and marine and industrial turbine engines and auxiliary power units. As shown in FIG. 1, the engine 10 has a longitudinal or axial centerline axis 12 that extends there through for reference purposes and generally along an axial direction A. The engine 10 further defines a radial direction R extended from the axial centerline 12, and a circumferential direction C (shown in FIGS. 2 and 6) around the axial centerline 12. The engine 10 further defines an upstream end 99 and a downstream 98 generally opposite of the upstream end 99 along the axial direction A. In general, the engine 10 may include a fan assembly 14 and a core engine 16 disposed downstream from the fan assembly 14.
The core engine 16 may generally include a substantially tubular outer casing 18 that defines an annular inlet 20. The outer casing 18 encases or at least partially forms, in serial flow relationship, a compressor section having a booster or low pressure (LP) compressor 22, a high pressure (HP) compressor 24, a combustion section 26, a turbine section including a high pressure (HP) turbine 28, a low pressure (LP) turbine 30 and a jet exhaust nozzle section 32. A high pressure (HP) rotor shaft 34 drivingly connects the HP turbine 28 to the HP compressor 24. A low pressure (LP) rotor shaft 36 drivingly connects the LP turbine 30 to the LP compressor 22. The LP rotor shaft 36 may also be connected to a fan shaft 38 of the fan assembly 14. In particular embodiments, as shown in FIG. 1, the LP rotor shaft 36 may be connected to the fan shaft 38 by way of a reduction gear 40 such as in an indirect-drive or geared-drive configuration. In other embodiments, the engine 10 may further include an intermediate pressure (IP) compressor and turbine rotatable with an intermediate pressure shaft.
As shown in FIG. 1, the fan assembly 14 includes a plurality of fan blades 42 that are coupled to and that extend radially outwardly from the fan shaft 38. An annular fan casing or nacelle 44 circumferentially surrounds the fan assembly 14 and/or at least a portion of the core engine 16. In one embodiment, the nacelle 44 may be supported relative to the core engine 16 by a plurality of circumferentially-spaced outlet guide vanes or struts 46. Moreover, at least a portion of the nacelle 44 may extend over an outer portion of the core engine 16 so as to define a bypass airflow passage 48 therebetween.
FIG. 2 is a cross sectional side view of an exemplary combustion section 26 of the core engine 16 as shown in FIG. 1. As shown in FIG. 2, the combustion section 26 may generally include an annular type combustor 50 having an annular inner liner 52, an annular outer liner 54 and a dome wall 56 that extends radially between upstream ends 58, 60 of the inner liner 52 and the outer liner 54 respectfully. In other embodiments of the combustion section 26, the combustion assembly 50 may be a can or can-annular type. As shown in FIG. 2, the inner liner 52 is radially spaced from the outer liner 54 with respect to axial centerline 12 (FIG. 1) and defines a generally annular combustion chamber 62 therebetween.
As shown in FIG. 2, the inner liner 52 and the outer liner 54 may be encased within an outer casing 64. An outer flow passage 66 may be defined around the inner liner 52, the outer liner 54, or both. The inner liner 52 and the outer liner 54 may extend from the dome wall 56 towards a turbine nozzle or inlet 68 to the HP turbine 28 (FIG. 1), thus at least partially defining a hot gas path between the combustor assembly 50 and the HP turbine 28. A fuel nozzle 70 may extend at least partially through the dome wall 56 and provide a fuel-air mixture 72 to the combustion chamber 62.
During operation of the engine 10, as shown in FIGS. 1 and 2 collectively, a volume of air as indicated schematically by arrows 74 enters the engine 10 through an associated inlet 76 of the nacelle 44 and/or fan assembly 14. As the air 74 passes across the fan blades 42 a portion of the air as indicated schematically by arrows 78 is directed or routed into the bypass airflow passage 48 while another portion of the air as indicated schematically by arrow 80 is directed or routed into the LP compressor 22. Air 80 is progressively compressed as it flows through the LP and HP compressors 22, 24 towards the combustion section 26. As shown in FIG. 2, the now compressed air as indicated schematically by arrows 82 flows across a compressor exit guide vane (CEGV) 67 and through a prediffuser 65 into a diffuser cavity or head end portion 84 of the combustion section 26.
The prediffuser 65 and CEGV 67 condition the flow of compressed air 82 to the fuel nozzle 70. The compressed air 82 pressurizes the diffuser cavity 84. The compressed air 82 enters the fuel nozzle 70 to mix with a fuel. The fuel nozzles 70 premix fuel and air 82 within the array of fuel injectors with little or no swirl to the resulting fuel-air mixture 72 exiting the fuel nozzle 70. After premixing the fuel and air 82 within the fuel nozzles 70, the fuel-air mixture 72 burns from each of the plurality of fuel nozzles 70 as an array of flames.
Referring still to FIGS. 1 and 2 collectively, the combustion gases 86 generated in the combustion chamber 62 flow from the combustor assembly 50 into the HP turbine 28, thus causing the HP rotor shaft 34 to rotate, thereby supporting operation of the HP compressor 24. As shown in FIG. 1, the combustion gases 86 are then routed through the LP turbine 30, thus causing the LP rotor shaft 36 to rotate, thereby supporting operation of the LP compressor 22 and/or rotation of the fan shaft 38. The combustion gases 86 are then exhausted through the jet exhaust nozzle section 32 of the core engine 16 to provide propulsive thrust.
As the fuel-air mixture burns, pressure oscillations occur within the combustion chamber 62. These pressure oscillations may be driven, at least in part, by a coupling between the flame's unsteady heat release dynamics, the overall acoustics of the combustor 50 and transient fluid dynamics within the combustor 50. The pressure oscillations generally result in undesirable high-amplitude, self-sustaining pressure oscillations within the combustor 50. These pressure oscillations may result in intense, frequently single-frequency or multiple-frequency dominated acoustic waves that may propagate within the generally closed combustion section 26.
Depending, at least in part, on the operating mode of the combustor 50, these pressure oscillations may generate acoustic waves at a multitude of low or high frequencies. These acoustic waves may propagate downstream from the combustion chamber 62 towards the high pressure turbine 28 and/or upstream from the combustion chamber 62 back towards the diffuser cavity 84 and/or the outlet of the HP compressor 24. In particular, as previously provided, low frequency acoustic waves, such as those that occur during engine startup and/or during a low power to idle operating condition, and/or higher frequency waves, which may occur at other operating conditions, may reduce operability margin of the turbofan engine and/or may increase external combustion noise, vibration, or harmonics.
Referring now to the exemplary embodiment of the combustor assembly 50 including the fuel nozzle 70 generally provided in FIG. 3, the fuel nozzle 70 includes a centerbody 105 extended along the lengthwise direction L. The fuel nozzle 70 defines a nozzle centerline 11 extended through the centerbody 105 of the fuel nozzle 70 along the lengthwise direction L. The fuel nozzle 70 defines one or more exit openings 107 in circumferential arrangement on the centerbody 105 relative to the nozzle centerline 11. In various embodiments, the exit openings 107 define a main fuel flow outlet from the fuel nozzle 70 to the combustion chamber 62. For example, the exit openings 107 may be configured to provide a flow of fuel to operate the combustor assembly 50 and the engine 10 at a maximum or high power condition or less.
Referring to the axial views of the exemplary embodiments of the fuel nozzle 70 generally provided in FIGS. 4-5, in addition to the perspective view generally provided in FIG. 3, the plurality of exit openings 107 are defined through the centerbody 105 at two or more locations different from one another along the lengthwise direction L, such as shown as a first exit opening 108 and a second exit opening 109. For example, a first reference plane 114 is defined along a radial direction RR extended from the nozzle centerline 11 at a first position along the lengthwise direction L at the first exit opening 108 defined through the centerbody 105. A second reference plane 116 is defined along the radial direction RR at a second position along the lengthwise direction at the second exit opening 109 defined through the centerbody 105. More particularly, the location along the lengthwise direction L of the plurality of exit openings 107 is defined in relationship to the downstream-most end of the centerbody 105, shown schematically by arrow 106.
The two or more locations at which the plurality of exit openings 107 is defined may further define a waveform along the centerbody 105. For example, the waveform may be sinusoidal, sawtooth, box, triangle, or another suitable waveform. In various embodiments, the waveform may be irregular along the centerbody 105, such as defining a variable frequency or amplitude. The waveform, as either a constant frequency or amplitude or as a variable frequency or amplitude, may define a third reference plane, a fourth reference plane, an Nth reference plane, etc. of the plurality of exit openings 107 such as to define a third exit opening, a fourth exit opening, an Nth reference opening, etc.
It should be appreciated that the first reference plane 114, the second reference plane 116, etc. may be defined through a center point of the plurality of exit openings 107. However, in other embodiments, each reference plane may be defined relative to a perimeter or another geometric feature of the exit openings 107.
The two or more lengthwise locations of the plurality of exit openings 107 may mitigate combustion dynamics by altering the heat release characteristics of each flame from each fuel nozzle 70. More specifically, the two or more lengthwise locations of the exit openings 107 of each fuel nozzle 70 may alter the flame structure, characteristic time, or both, for each fuel nozzle 70, thereby decoupling heat release from pressure fluctuations and mitigating undesired combustion dynamics.
Referring now to FIGS. 4-5, in various embodiments, the fuel nozzle 70 may more particularly define the first reference plane 114, at which the first exit opening 108 is disposed along the lengthwise direction L, at a first distance 115 relative to the downstream-most end 106 of the centerbody 105. The second reference plane 116, at which the second exit opening 109 is disposed along the lengthwise direction L, is defined at a second distance 117 relative to the downstream-most end 106 of the centerbody 105. Still further, as shown in FIGS. 4-5, the first exit opening 108 and the second exit opening 109 may occupy unique locations along the circumferential direction relative to the nozzle centerline 11 such that neither the first exit opening 108 or the second exit opening 109 is upstream or downstream of one another along the same radial location along the centerbody 105 (i.e., the plurality of exit openings 107 are staggered). It should be appreciated that a third exit opening, a fourth exit opening, an Nth exit opening may be disposed similarly.
Referring now to FIG. 5, in one embodiment, the plurality of exit openings 107 defines two or more cross sectional areas through the centerbody 105 different from one another. For example, the fuel nozzle 70 defines a first exit opening 108 defining a first cross sectional area and a second exit opening 109 defining a second cross sectional area greater or lesser than the first cross sectional area. The plurality of exit openings 107 provide a fuel to the combustion chamber 62 at two or more pressures or flow rates corresponding to the two or more cross sectional areas through the centerbody 105. The two or more cross sectional areas of the exit openings 107 providing two or more pressures or flow rates of fuel to the combustion chamber 62 may mitigate such pressure oscillations by altering the heat release characteristics of each flame from each fuel nozzle 70. More specifically, the two or more exit openings 107 of each fuel nozzle 70 may alter the flame structure, characteristic time, or both, for each fuel nozzle 70, thereby decoupling heat release from pressure fluctuations and mitigating undesired combustion dynamics.
In one embodiment, the plurality of exit openings 107 of each fuel nozzle 70 may define a nominal first exit opening 108 of the first cross sectional area and the second exit opening 109 of the second cross sectional area up to approximately 50% greater or lesser than the first cross sectional area. It should be appreciated that a volume of a fuel passage within the fuel nozzle 70 extending in fluid communication with each exit opening 107 may generally correspond to the cross sectional area defined by each exit opening 107 (e.g., first cross sectional area corresponding to the first exit opening 108, the second cross sectional area corresponding to the second exit opening 109, etc.). Still further, it should be appreciated that the fuel nozzle 70 may define a third exit opening corresponding to a third cross sectional area, a fourth exit opening corresponding to a fourth cross sectional area, etc., in which each exit opening and cross sectional area defines a different pressure, flow rate, or both of the fuel egressing therefrom into the combustion chamber 62.
In still various embodiments, the plurality of exit openings 107 each define a unique cross sectional area at each circumferential location through the centerbody 105. For example, for N quantity of exit openings 107, the centerbody 105 defines N cross sectional areas. In another embodiment, the plurality of exit openings 107 is defined in asymmetric circumferential arrangement through the centerbody 105. For example, two or more exit openings 107 are defined unequally along the circumference of the centerbody 105 relative to two or more other exit openings 107.
Referring to FIGS. 9-10, the engine 10 may define a plurality of the fuel nozzles 70 defining the embodiments generally provided in and described in regard to FIGS. 3-5 disposed in circumferential arrangement around the axial centerline A. In various embodiments, the engine 10 defines a first fuel nozzle and a second fuel nozzle. The first fuel nozzle defines a first waveform of the plurality of exit openings 107 and the second fuel nozzle defines a second waveform of the plurality of exit openings 107. In another embodiment, the first fuel nozzle defines a first cross sectional area of the plurality of exit openings 107 and the second fuel nozzle defines a second cross sectional area of the plurality of exit openings 107. In still another embodiment, the first fuel nozzle defines a first exit opening 108 and a second exit opening 109 of two different cross sectional areas at two different locations along the lengthwise direction L. The second fuel nozzle defines a first exit opening 108 at two different locations along the lengthwise direction L.
In one embodiment, the plurality of fuel nozzles 70 are defined in alternating circumferential arrangement around the axial centerline 12 of the engine 10. For example, the plurality of fuel nozzles 70 may define the first fuel nozzle, the second fuel nozzle, the first fuel nozzle, etc. along the circumferential direction relative to the axial centerline 12.
In various embodiments, the engine 10 defines at least half of the plurality of fuel nozzles 70 as either the first fuel nozzle or the second fuel nozzle, in which the remainder are defined as either the second fuel nozzle or the first fuel nozzle. In one embodiment, at least two or more of the first fuel nozzle or the second fuel nozzle are in sequential circumferential arrangement around the axial centerline 12 of the engine 10. For example, the engine 10 may define, in sequential arrangement around the axial centerline 12, the first fuel nozzle, another first fuel nozzle, and one or more additional first fuel nozzles, then one or more of a second fuel nozzle, then one or more of a first fuel nozzle.
The various embodiments of the engine 10 may provide a flow of fuel through the centerbody 105 and egressing from the plurality of exit openings 107 into the combustion chamber 62. The waveform definition of the plurality of exit openings 107 along the centerbody 105, the first cross sectional area of the first exit opening 108 and the second cross sectional area of the second exit opening 109, or a combination thereof provides a circumferentially asymmetric flame within the combustion chamber 62 relative to the axial centerline 12.
All or part of the combustor assembly 50 and the fuel nozzle 70 may each 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 fuel nozzle 70. Furthermore, the combustor assembly 50 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.
This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.

Claims

What is claimed is:

1. A combustor assembly for a gas turbine engine, the combustor assembly defining a combustor centerline and a radial direction extended from the combustor centerline, the radial direction and combustor centerline together defining a reference plane, the combustor assembly comprising:

a fuel nozzle comprising a centerbody extended along a lengthwise direction, wherein the fuel nozzle defines a nozzle centerline extended through the centerbody along the lengthwise direction, the centerbody defining a plurality of exit openings in circumferential arrangement relative to the nozzle centerline, the plurality of exit openings defining two or more locations different from one another on the centerbody along the lengthwise direction.

2. The combustor assembly of claim 1, wherein the two or more locations along the lengthwise direction and the circumferential arrangement of the plurality of exit openings defines a waveform.

3. The combustor assembly of claim 2, wherein the waveform is sinusoidal, sawtooth, or box.

4. The combustor assembly of claim 1, wherein the plurality of exit openings defines two or more cross sectional areas, wherein each cross sectional area is different from one another.

5. The combustor assembly of claim 4, wherein the plurality of exit openings provides a fuel to a combustion chamber at two or more flow rates each corresponding to the two or more cross sectional areas of the plurality of exit openings.

6. The combustor assembly of claim 1, wherein the plurality of exit openings defines a first opening and a second opening, the first opening defining a first lengthwise location and the second opening defining a second lengthwise location different from the first lengthwise location.

7. The combustor assembly of claim 1, wherein the plurality of exit openings defines a first opening and a second opening, the first opening defining a first cross sectional area and the second opening defining a second cross sectional area different from the first cross sectional area.

8. The combustor assembly of claim 1, wherein a single exit opening of the plurality of exit openings is defined at each circumferential location along the centerbody from the nozzle centerline.

9. The combustor assembly of claim 1, wherein the plurality of exit openings are defined in asymmetric circumferential arrangement through the centerbody.

10. A gas turbine engine defining an axial centerline and a radial direction extended from the axial centerline, the radial direction and axial centerline together defining a reference plane, the gas turbine engine comprising:

a combustor assembly comprising a plurality of fuel nozzles disposed in circumferential arrangement around the axial centerline of the gas turbine engine, wherein each fuel nozzle comprises a centerbody extended along a lengthwise direction, and wherein the fuel nozzle defines a nozzle centerline extended through the centerbody along the lengthwise direction, the centerbody defining a plurality of exit openings in circumferential arrangement relative to the nozzle centerline, the plurality of exit openings defining two or more cross sectional areas different from one another on the centerbody.

11. The gas turbine engine of claim 10, wherein the plurality of exit openings are defined along the same reference plane.

12. The gas turbine engine of claim 10, wherein the plurality of exit openings are defined along two or more locations different from one another along the lengthwise direction.

13. The gas turbine engine of claim 12, wherein the plurality of exit openings provides a fuel to a combustion chamber at two or more flow rates each corresponding to the two or more cross sectional areas of the plurality of exit openings.

14. The gas turbine engine of claim 12, wherein the two or more locations along the lengthwise direction and the circumferential arrangement of the plurality of exit openings defines a waveform along the centerbody.

15. The gas turbine engine of claim 14, wherein the waveform is sinusoidal, sawtooth, or box.

16. The gas turbine engine of claim 10, wherein the plurality of exit openings defines a first opening and a second opening, the first opening defining a first lengthwise location and the second opening defining a second lengthwise location different from the first lengthwise location.

17. The gas turbine engine of claim 10, wherein a single exit opening of the plurality of exit openings is defined at each circumferential location along the centerbody from the nozzle centerline.

18. The gas turbine engine of claim 10, wherein the plurality of fuel nozzles defines a first nozzle and a second nozzle, and wherein the first nozzle defines a plurality of exit openings, the plurality of exit openings comprising a first opening defining a first cross sectional area and a second opening defining a second cross sectional area different from the first cross sectional area, and wherein the second nozzle defines the plurality of exit openings at a plurality of locations along the lengthwise direction.

19. The gas turbine engine of claim 18, wherein the plurality of fuel nozzles are defined in alternating circumferential arrangement.

20. The gas turbine engine of claim 18, wherein at least half of the plurality of fuel nozzles defines the first nozzle and the remainder of the plurality of fuel nozzles defines the second nozzle, and wherein at least two or more of the first fuel nozzle or the second fuel nozzle are in sequential circumferential arrangement.