Dilution Structure for Gas Turbine Engine Combustor

Abstract

The present disclosure is directed to a combustor assembly for a gas turbine engine. The combustor assembly includes a liner defining a combustion chamber therewithin and a pressure plenum surrounding the liner. The liner defines an opening. The liner includes a walled chute disposed at least partially within the opening. The walled chute defines an inner surface at least partially defining a jet destabilizer.

Description

FIELD

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

BACKGROUND

Combustion assemblies for gas turbine engines generally include orifices in the combustion liners to dilute the combustion gases within the combustion chamber with air from the diffuser cavity. The air may be employed to mix with an over rich combustion gas mixture to complete the combustion process; to stabilize combustion flames within the recirculation zone of the combustion chamber; to minimize oxides of nitrogen emissions; or to decrease combustion gas temperature before egressing to the turbine section.

Although dilution orifices provide known benefits, there is a need for structures that may provide and improve upon these benefits via egressing the air into the combustion chamber in increasingly detailed or specific modes.

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 includes a liner defining a combustion chamber therewithin and a pressure plenum surrounding the liner. The liner defines an opening. The liner includes a walled chute disposed at least partially within the opening. The walled chute defines an inner surface at least partially defining a jet destabilizer.

In various embodiments, the walled chute includes a span of the inner surface from a cold end proximate to the pressure plenum to a hot end proximate to the combustion chamber. The jet destabilizer includes a contoured surface is defined within approximately 33% of the span from the hot end. In one embodiment, the contoured surface defines a chevron, a waveform, a rib structure, or a vane structure.

In still various embodiments, the inner surface at least partially defines a groove. In one embodiment, the groove is defined substantially helical. In another embodiment, the groove is defined substantially perpendicular to the inner surface.

In still yet various embodiments, the jet destabilizer defines a member extended from the inner surface at least partially across the opening. In one embodiment, the member is extended from the inner surface of the walled chute substantially perpendicular to a direction of flow of combustion gases formed within the combustion chamber. In another embodiment, the member is extended from the inner surface of the walled chute substantially co-directional to a direction of flow of combustion gases formed within the combustion chamber.

In one embodiment, the walled chute is extended at least partially into the combustion chamber.

Another aspect of the present disclosure is directed to a gas turbine engine including a combustor assembly. The combustor assembly includes a liner defining a combustion chamber therebetween and a pressure plenum surrounding the liner. The liner defines an opening. The liner includes a walled chute disposed at least partially within the opening. The walled chute includes an inner surface. The inner surface includes a jet destabilizer in a first flow passage defined within the walled chute.

In various embodiments, the jet destabilizer includes a member extended across a diameter of the opening of the liner. In one embodiment, the member is extended from the inner surface substantially perpendicular to a direction of flow of combustion gases formed within the combustion chamber. In another embodiment, the member is extended from the inner surface substantially co-directional to a direction of flow of combustion gases formed within the combustion chamber.

In still various embodiments, the jet destabilizer is defined at the inner surface between a cold end and a hot end of the first flow passage. In one embodiment, the jet destabilizer is defined within 33% of the walled chute from the cold end or the hot end.

In one embodiment, the walled chute comprises equal to or less than six times a diameter of the opening of the liner.

In another embodiment, the jet destabilizer comprises a groove defined at the inner surface of the walled chute.

In still another embodiment, the jet destabilizer comprises a chevron, a waveform, a rib structure, or a vane structure at the walled chute.

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 figures, in which:

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

FIG. 2 is a perspective cross sectional view of an exemplary embodiment of a combustor assembly of the exemplary engine shown in FIG. 1;

FIG. 3-4 are cross sectional side views of a portion of an exemplary embodiment of a combustor assembly;

FIGS. 5-7 are cross sectional side views of exemplary embodiments of walled chutes of the combustor assembly of FIGS. 2-4;

FIGS. 8-11 are perspective views of exemplary embodiments of walled chutes of the combustor assembly of FIGS. 2-4;

FIGS. 12-14 are cross sectional side views of exemplary embodiments of walled chutes of the combustor assembly of FIGS. 2-4;

FIGS. 15-16 are perspective views of exemplary embodiments of walled chutes of the combustor assembly of FIGS. 2-4; and

FIGS. 17-18 are top-down views of exemplary embodiments of portions of the combustor assembly including embodiments of the walled chute of FIGS. 12-15.

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.

Embodiments of combustor assembly dilution structures are generally provided that may improve emissions and combustion gas quenching via egressing the air into the combustion chamber in increasingly detailed or specific modes. The various embodiments of combustor assemblies generally define a walled chute configured to egress air from the diffuser cavity to the combustion chamber in multiple or tailored modes.

Referring now to the drawings, FIG. 1 is a schematic partially cross-sectioned side view of an exemplary high bypass 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 turbomachinery in general, including turbojet, turboprop, and turboshaft gas turbine engines, including marine and industrial turbine engines and auxiliary power units. As shown in FIG. 1, the engine 10 has a longitudinal or axial engine centerline axis 12 that extends there through for reference purposes. The engine 10 defines a longitudinal direction L and an upstream end 99 and a downstream end 98 along the longitudinal direction L. The upstream end 99 generally corresponds to an end of the engine 10 along the longitudinal direction L from which air enters the engine 10 and the downstream end 98 generally corresponds to an end at which air exits the engine 10, generally opposite of the upstream end 99 along the longitudinal direction L. 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 compressor and turbine rotatable with an intermediate pressure shaft altogether defining a three-spool gas turbine engine.

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 bulkhead 56 that extends radially between upstream ends of the inner liner 52 and the outer liner 54 respectively. In other embodiments of the combustion section 26, the combustion assembly 50 may be a can-annular type. The combustor 50 further includes a dome assembly 57 extended radially between the inner liner 52 and the outer liner 54 downstream of the bulkhead 56. As shown in FIG. 2, the inner liner 52 is radially spaced from the outer liner 54 with respect to engine centerline 12 (FIG. 1) and defines a generally annular combustion chamber 62 therebetween. In particular embodiments, the inner liner 52, the outer liner 54, and/or the dome assembly 57 may be at least partially or entirely formed from metal alloys or ceramic matrix composite (CMC) materials.

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 of a diffuser cavity or pressure plenum 84 may be defined around the inner liner 52 and/or the outer liner 54. The inner liner 52 and the outer liner 54 may extend from the bulkhead 56 towards a turbine nozzle or inlet 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 bulkhead 56 to provide a fuel 72 to mix with the air 82(a) and burn at the combustion chamber 62. In various embodiments, the bulkhead 56 includes a fuel-air mixing structure attached thereto (e.g., a swirler assembly).

Referring still to FIG. 2, the inner liner 52 and the outer liner 54 each define one or more openings 105 through the liners 52, 54. A walled chute 100 is disposed at least partially within the opening 105. In various embodiments, the walled chute 100 is extended at least partially into the combustion chamber 62. In other embodiments, the walled chute 100 is extended at least partially into the pressure plenum 84. In still other embodiments, the walled chute 100 is approximately flush or even to the liner 52, 54 to which the walled chute 100 is attached and disposed in the opening 105. The walled chute 100 generally defines a walled enclosure defining a first flow passage 111 therethrough from the pressure plenum 84 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 into a diffuser cavity or pressure plenum 84 of the combustion section 26. The pressure plenum 84 generally surrounds the inner liner 52 and the outer liner 54, and generally upstream of the combustion chamber 62.

The compressed air 82 pressurizes the pressure plenum 84. A first portion of the of the compressed air 82, as indicated schematically by arrows 82(a) flows from the pressure plenum 84 into the combustion chamber 62 where it is mixed with the fuel 72 and burned, thus generating combustion gases, as indicated schematically by arrows 86, within the combustor 50. Typically, the LP and HP compressors 22, 24 provide more compressed air to the pressure plenum 84 than is needed for combustion. Therefore, a second portion of the compressed air 82 as indicated schematically by arrows 82(b) may be used for various purposes other than combustion. For example, as shown in FIG. 2, compressed air 82(b) may be routed into the outer flow passage 66 to provide cooling to the inner and outer liners 52, 54.

Additionally, at least a portion of compressed air 82(b) flows out of the pressure plenum 84 into the combustion chamber 62 via the first flow passage 111 defined by the walled chute 100, such as depicted via arrows 83. A portion of the compressed air 82(b), shown as air 83, egresses from the pressure plenum 84 through the first flow passage 111 into the combustion chamber 62.

Referring now to FIGS. 3-11, the walled chute 100 defines an inner surface 101 at the first flow passage 111. The inner surface 101 at least partially defines a jet destabilizer. In one embodiment, the jet destabilizer defines a contoured surface 103. In various embodiments, the walled chute 100 defines a span 107 of the inner surface 101 from a cold end 106 proximate to the pressure plenum 84 to a hot end 108 proximate to the combustion chamber 62.

Referring to FIGS. 5-11, in various embodiments, the contoured surface 103 may define a chevron, a waveform, a rib structure, or a vane structure. For example, referring to FIGS. 5-7, the contoured surface 103 may be defined over 50% or more of the inner surface 101 of the walled chute 100. The contoured surface 103 may define a plurality of grooves 104. In one embodiment, such as generally provided in regard to FIG. 5, the contoured surface 103 defines the plurality of grooves 104 extended substantially perpendicular relative to the flow of air 83 through the first flow passage 111 defined within the walled chute 100.

In another embodiment, such as generally provided in regard to FIG. 6, the contoured surface 103 defines the plurality of grooves 104 extended at an acute angle relative to the flow of air 83 through the first flow passage 111. Alternatively, the contoured surface 103 defines the plurality of grooves 104 extended at an acute angle relative to the walled chute 100 extended through the liner 52, 54. For example, in various embodiments, the grooves 104 may be extended at an acute angle greater than zero degrees and less than 90 degrees relative to the walled chute 100 and/or the inner surface 101 of the walled chute 100.

In yet another embodiment, such as generally provided in regard to FIG. 7, the contoured surface 103 defines the plurality of grooves 104 extended substantially helical along the inner surface 101. For example, the grooves 104 may be extended substantially along the span 107 of the inner surface 101.

Referring to FIGS. 5-11, the contoured surface 103 may be defined protruding into the inner surface 101 of the walled chute 100. In other embodiments, the contoured surface 103 may be defined extended from the inner surface 101 of the walled chute 100, such as into the first flow passage 111.

In one embodiment, such as generally provided in regard to FIGS. 8-11, the contoured surface 103 is defined within approximately 33% of the span 107 from the hot end 108. In other embodiments, the contoured surface 103 is defined within approximately 10% of the span 107 from the hot end 108. In other embodiments, the contoured surface 103 may be defined within approximately 33% of the span 107 from the cold end 106. Referring to FIGS. 8-11, the contoured surface 103 may generally define a chevron or a waveform at least partially along the walled chute 100. In one embodiment, such as generally shown in regard to FIGS. 8-9, the contoured surface 103 defining a chevron or waveform 113 may be defined substantially inward toward the first flow passage 111. In another embodiment, such as generally shown in regard to FIG. 10, the contoured surface 103 defining the chevron or waveform 113 may be defined substantially along direction of the flow of air 83 through the walled chute 100.

It should be appreciated that various embodiments of the contoured surface 103 may defining the chevrons or waveforms 113 may define a triangle waveform, square or rectangular or step waveform, a sinusoidal waveform, a sawtooth waveform, or other waveform. Additionally, or alternatively, the chevron or waveforms 113 may be defined regular (e.g., constant frequency) along the walled chute 100. In other embodiments, the chevron or waveforms 113 may be defined irregularly (e.g., variable, asymmetric, or irregular frequency) along the walled chute 100.

Embodiments of the contoured surface 103 may generally enable, promote, or increase turbulence in the flow of air 83 from the pressure plenum 84 to the combustion chamber 62. The increased turbulence of the flow of air 83 may improve mixing of the flow of air 83 with the combustion gases 86 such as to decrease production of nitrogen oxides (e.g., NOx), improve durability of the combustor assembly 50 (e.g., improve durability at the liners 52, 54), or both. As another example, the walled chute 100 including the contoured surface 103 may further improve mixing of the flow of air 83 with the combustion gases 86 while mitigating losses in penetration of the flow of air 83 with the combustion gases 86 into the combustion chamber 62.

Referring to FIGS. 2-18, in various embodiments, the walled chute 100 may define the span 107 as equal to or greater than a diameter 109 (FIG. 2) of the opening 105 in the liner 52, 54. In one embodiment, the span 107 is less than or equal to six times (6×) the diameter 109 of the opening 105. In another embodiment, the span 107 is less than or equal to four times (4×) the diameter 109 of the opening 105. In still various embodiments, the span 107 of the walled chute 100 may extend into the combustion chamber 62 from the liner 52, 54 by six times (6×) the diameter 109 or less. In still yet various embodiments, the span 107 of the walled chute 100 may extend into the combustion chamber 62 from the liner 52, 54 by four times (4×) the diameter 109 or less. For example, the walled chute 100 may extend approximately flush or even into the liner 52, 54, or into the pressure plenum 84, and additionally extend into the combustion chamber 62 (e.g., FIG. 4), and a portion of the walled chute 100 extended into the combustion chamber 62 may be extended into the combustion chamber 62 by six times or less of the diameter 109 of the opening 105.

Referring now to FIGS. 12-18, in various embodiments, the walled chute 100 may further define the jet destabilizer as a member 110 extended from the inner surface 101 at least partially across the first flow passage 111. For example, the member 110 may be extended across the diameter 109 of the opening 105.

In one embodiment, such as generally provided in regard to FIG. 12, the member 110 may be extended across the first flow passage 111 approximately equidistant between the cold end 106 and the hot end 108 of the walled chute 100. In other embodiments, such as generally depicted in FIGS. 12-16, the member 110 may be extended through the walled chute 100 within the liner 52, 54.

In other embodiments, such as generally provided in regard to FIGS. 12-16, the member 110 may be extended across the first flow passage 111 within approximately 33% of the span 107 from the cold end 106. In still other embodiments, the member 110 may be extended across the first flow passage 111 within approximately 33% of the span 107 from the hot end 108.

In another embodiment of the walled chute 100, such as generally provided in regard to FIGS. 14-17, the member 110 is extended from the inner surface 101 (FIGS. 14-15) of the walled chute 100 substantially perpendicular (FIG. 17) to a direction of flow of combustion gases 86 formed within the combustion chamber 62 (FIG. 2). The member 110 may define a jet destabilizer disposed substantially perpendicular to the cross flow direction of the combustion gases 86 (i.e., perpendicular to the flow of combustion gases 86 toward the turbines 28, 30). The walled chute 100 defining the member 110 as a jet destabilizer member may split a counter-rotating vortex pair (CVP) into two or more pairs (e.g., air 83(a) and air 83(b) in FIG. 14), thereby adding additional vorticity or wake from the flow of air 83 to the jet flow of combustion gases 86. The additional vorticity may induce cross-wise perturbations that may further be amplified or destabilized to enable oscillation to the flow of air 83 defining a dilution jet to the combustion gases 86. The oscillation of the flow of air 83 may improve penetration and mixing of the flow of air 83 with the combustion gases 86 to reduce production of nitrogen oxides (i.e., NOx).

Referring now to FIGS. 14-15 and FIG. 18, in yet another embodiment, the member 110 is extended from the inner surface 101 of the walled chute 100 substantially co-directional (FIG. 18) or parallel to a direction of flow of combustion gases 86 formed within the combustion chamber 62. The member 110 defining the jet destabilizer may weaken the CVP, thereby improving or promoting spread of the air 83 over the liners 52, 54 such as to improve heat transfer (e.g., cooling) at the liner 52, 54. As such, the member 110 disposed substantially co-directional to the flow of combustion gases 86 may further improve durability of the combustor assembly 50, such as at the liner 52, 54.

Various embodiments of the engine 10 and combustor assembly 50 may define a rich burn combustor in which the walled chute 100 may define dilution jets providing additional mixing air (e.g., air 83) with a mixture of combustion gases (e.g., combustion gases 86) to improve or complete the combustion process. The walled chute 100 may further define dilution jets that further enable or augment a combustion recirculation zone within the combustion chamber 62 to stabilize a flame therein. Still further, the walled chute 100 may define dilution jets that may relatively rapidly quench the combustion gases 86 to minimize production of nitrogen oxides. Furthermore, various embodiments of the combustor assembly 50 and walled chute 100 shown and described herein may enable customization of a distribution of combustion gas temperature to improve durability of components at or downstream of the combustor assembly 50 (e.g., the liners 52, 54, the HP turbine 28).

Still further, the walled chute 100 may generally define the member 110 as a bluff-body device such as to provide a jet destabilizer to modify counter rotating vortex pairs (CVP) formed in jets in cross flow (JIC). For example, the portion of air 83 provided through the walled chute 100 may define a CVP formed relative to the flow of combustion gases 86 defining a JIC.

All or part of the combustor assembly 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 combustor 50, including, but not limited to, the liners 52, 54, the walled chute 100, the member 110, or combinations thereof. 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.

Various embodiments of the walled chute 100 including the member 110 may define the member 110 of one or more cross sectional areas, such as, but not limited to, a circular cross section (e.g., shown in FIGS. 16-18), a rectangular cross section, a ovular or racetrack cross section, an airfoil or teardrop cross section, a polygonal cross section, or an oblong cross section, or another suitable cross section, or combinations thereof.

Additionally, or alternatively, various embodiments of the walled chute 100 and/or the opening 105 through which the walled chute 100 is disposed may define one or more cross sectional areas, such as, but not limited to, a circular cross section (e.g., shown in FIGS. 2-18), a rectangular cross section, a ovular or racetrack cross section, an airfoil or teardrop cross section, a polygonal cross section, or an oblong cross section, or another suitable cross section, 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

1. A combustor assembly for a gas turbine engine, the combustor assembly comprising:

a liner defining a combustion chamber therewithin and a pressure plenum surrounding the liner, wherein the liner comprises an opening, and wherein the liner comprises a walled chute disposed at least partially within the opening, and further wherein the walled chute comprises an inner surface, and wherein the inner surface comprises a jet destabilizer.

2. The combustor assembly of claim 1, wherein the walled chute defines a span of the inner surface from a cold end proximate to the pressure plenum to a hot end proximate to the combustion chamber, and wherein the jet destabilizer comprises a contoured surface defined within approximately 33% of the span from the hot end.

3. The combustor assembly of claim 2, wherein the contoured surface comprises a chevron, a waveform, a rib structure, or a vane structure.

4. The combustor assembly of claim 1, wherein the inner surface comprises a groove.

5. The combustor assembly of claim 4, wherein the groove is defined substantially helical.

6. The combustor assembly of claim 4, wherein the groove is defined substantially perpendicular to the inner surface.

7. The combustor assembly of claim 1, wherein the jet destabilizer comprises a member extended from the inner surface at least partially across the opening.

8. The combustor assembly of claim 7, wherein the member is extended from the inner surface of the walled chute substantially perpendicular to a direction of flow of combustion gases formed within the combustion chamber.

9. The combustor assembly of claim 7, wherein the member is extended from the inner surface of the walled chute substantially co-directional to a direction of flow of combustion gases formed within the combustion chamber.

10. The combustor assembly of claim 1, wherein the walled chute is extended at least partially into the combustion chamber.

11. The combustor assembly of claim 1, wherein the walled chute is extended at least partially into the pressure plenum.

12. A gas turbine engine, the gas turbine engine comprising:

a combustor assembly comprising a liner defining a combustion chamber therebetween and a pressure plenum surrounding the liner, wherein the liner comprises an opening, and wherein the liner comprises a walled chute disposed at least partially within the opening, and further wherein the walled chute comprises an inner surface, and wherein the inner surface comprises a jet destabilizer in a first flow passage defined within the walled chute.

13. The gas turbine engine of claim 12, wherein the jet destabilizer comprises a member extended across a diameter of the opening of the liner.

14. The gas turbine engine of claim 13, wherein the member is extended from the inner surface substantially perpendicular to a direction of flow of combustion gases formed within the combustion chamber.

15. The gas turbine engine of claim 13, wherein the member is extended from the inner surface substantially co-directional to a direction of flow of combustion gases formed within the combustion chamber.

16. The gas turbine engine of claim 12, wherein the jet destabilizer is defined at the inner surface between a cold end and a hot end of the first flow passage.

17. The gas turbine engine of claim 16, wherein the jet destabilizer is defined within 33% of the walled chute from the cold end or the hot end.

18. The gas turbine engine of claim 12, wherein the walled chute is equal to or less than six times a diameter of the opening of the liner.

19. The gas turbine engine of claim 12, wherein the jet destabilizer comprises a groove defined at the inner surface of the walled chute.

20. The gas turbine engine of claim 12, wherein the jet destabilizer comprises a chevron, a waveform, a rib structure, or a vane structure at the walled chute.