阅读说明：本技术 用于燃气涡轮发动机燃烧器的稀释结构 (Dilution structure for gas turbine engine combustor ) 是由 帕鲁马鲁·乌坎蒂 赫兰雅·库马尔·纳斯 贾扬斯·塞卡尔 亚瑟·韦斯利·约翰森 古鲁纳斯·甘迪 于 2019-08-01 设计创作，主要内容包括：本公开针对一种用于燃气涡轮发动机的燃烧器组件。燃烧器组件包括衬里和压力气室,衬里限定在其内的燃烧室,压力气室围绕衬里。衬里限定开口。衬里包括至少部分地安置在开口内的有壁槽。有壁槽限定内表面,内表面至少部分地限定射流扰乱器。(The present disclosure is directed to a combustor assembly for a gas turbine engine. The combustor assembly includes a liner defining a combustion chamber therein and a pressure plenum surrounding the liner. The liner defines an opening. The liner includes a walled channel at least partially disposed within the opening. The walled channel defines an inner surface that at least partially defines a jet disruptor.)

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

a liner defining a combustion chamber therein and a plenum surrounding the liner, wherein the liner comprises an opening, and wherein the liner comprises a walled groove disposed at least partially within the opening, and further wherein the walled groove comprises an inner surface, and wherein the inner surface comprises a jet disruptor.

2. The burner assembly of claim 1 wherein said walled groove defines a span from said inner surface proximate a cold end of said plenum to a hot end of said combustion chamber, and wherein said jet disrupter comprises a contoured surface defined within approximately 33% of said span from said hot end.

3. The combustor assembly of claim 2, wherein the contoured surface comprises a V-shape, a wave shape, a rib structure, or a vane structure.

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

5. The burner assembly of claim 4 wherein the groove defines a substantially helical shape.

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

7. The burner assembly of claim 1 wherein the jet disrupter comprises a member extending from the inner surface at least partially across the opening.

8. The burner assembly of claim 7 wherein said member extends from said inner surface of said walled channel substantially perpendicular to a direction of flow of combustion gases formed within said combustion chamber.

9. The burner assembly of claim 7 wherein said member extends from said inner surface of said walled channel substantially co-directionally with the direction of flow of combustion gases formed within said combustion chamber.

10. The burner assembly of claim 1 wherein said walled channel extends at least partially into said combustion chamber.

Technical Field

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

Background

Combustion assemblies for gas turbine engines generally include apertures in the combustion liner to dilute the combustion gases within the combustion chamber with air from the diffuser cavity. Air may be employed to mix with the overly rich combustion gas mixture to complete the combustion process; to stabilize the combustion flame within the recirculation zone of the combustion chamber; to minimize nitrogen oxide emissions; or to reduce the combustion gas temperature prior to flowing out to the turbine section.

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

Disclosure of Invention

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 therein and a pressure plenum surrounding the liner. The liner defines an opening. The liner includes a walled channel at least partially disposed within the opening. The walled channel defines an inner surface that at least partially defines a jet disrupter (jetdestabilizer).

In various embodiments, the walled groove comprises a span from the inner surface proximate the cold end of the plenum to the hot end of the combustion chamber. The jet disrupter includes a contoured surface defined within approximately 33% of the span from the hot end. In one embodiment, the contoured surface defines a V-shape, a wave shape, a rib structure, or a vane structure.

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

In various embodiments, the jet disrupter defines a member extending from the inner surface at least partially across the opening. In one embodiment, the member extends from the inner surface of the walled channel substantially perpendicular to a flow direction of combustion gases formed within the combustion chamber. In another embodiment, the member extends from the inner surface of the walled channel substantially co-directionally with the direction of flow of combustion gases formed within the combustion chamber.

In one embodiment, the walled groove extends 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 channel at least partially disposed within the opening. The walled channel includes an inner surface. The inner surface includes a jet disruptor in a first flow channel defined within the walled slot.

In various embodiments, the jet disrupter comprises a member that extends across a diameter of the opening of the liner. In one embodiment, the member extends from the inner surface substantially perpendicular to a flow direction of combustion gases formed within the combustion chamber. In another embodiment, the member extends from the inner surface substantially co-directionally with a flow direction of combustion gases formed within the combustion chamber.

In still other embodiments, the jet disrupter is defined at the inner surface between the cold end and the hot end of the first flow channel. In one embodiment, the jet disrupters are defined within 33% of the walled slots from the cold or hot ends.

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

In another embodiment, the jet disrupter comprises a groove defined at an inner surface of the walled trough.

In yet another embodiment, the jet disrupter comprises a V-shaped, wave-shaped, rib-shaped structure or vane structure at the wall slot.

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.

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 including 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;

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 slots of the burner assembly of FIGS. 2-4;

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

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

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

fig. 17-18 are top and bottom views of exemplary embodiments of portions of a combustor assembly including the walled trough embodiment of fig. 12-15.

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

Detailed Description

Reference will now be made in detail to embodiments of the invention, one or more examples of which are illustrated in the accompanying 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 cover the modifications and variations of this invention provided they come within the scope of the appended claims and their equivalents.

The terms "first," "second," and "third" as used herein may be used interchangeably to distinguish one element from another without intending to indicate the position or importance of the various elements.

The terms "upstream" and "downstream" refer to relative directions with respect to fluid flow in a fluid path. For example, "upstream" refers to the direction of fluid flow therefrom, and "downstream" refers to the direction of fluid flow thereto.

Embodiments of combustor assembly dilution structures are generally provided that, in increasingly detailed or specific modes, may improve emissions and combustion gas quenching via outflow of air into the combustion chamber. Various embodiments of the combustor assembly generally define wall slots configured to flow air from the diffuser cavity into the combustion chamber in a variety or customized modes.

Referring now to the drawings, FIG. 1 is a schematic partial cross-sectional side view of an exemplary high bypass turbofan engine 10, referred to herein as "engine 10", that may incorporate various embodiments of the present disclosure. Although further described below with reference to turbofan engines, the present disclosure is generally applicable to turbomachines as well, including turbojets, turboprops, and turboshaft gas turbine engines, including marine and industrial turbine engines and auxiliary power units. As shown in FIG. 1, for reference purposes, the engine 10 has a longitudinal or axial engine centerline axis 12 extending therethrough. The engine 10 defines a longitudinal direction L and upstream and downstream ends 99, 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 of the engine 10 from which air exits, generally opposite 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 of the fan assembly 14.

The core engine 16 may generally include a substantially cylindrical outer casing 18, the outer casing 18 defining an annular inlet 20. The outer housing 18 encloses or at least partially forms in continuous 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 an injection exhaust nozzle section 32. A High Pressure (HP) spool 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. LP rotor shaft 36 may also be connected to a fan shaft 38 of fan assembly 14. In certain embodiments, as shown in FIG. 1, the LP rotor shaft 36 may be connected to the fan shaft 38 via a reduction gear 40, for example, in an indirect drive or gear drive configuration. In other embodiments, engine 10 may further include an Intermediate Pressure (IP) compressor and a turbine rotatable with the intermediate pressure shaft, collectively defining a three-shaft gas turbine engine.

As shown in FIG. 1, the fan assembly 14 includes a plurality of fan blades 42, the plurality of fan blades 42 coupled to the fan shaft 38 and extending radially outward from the fan shaft 38. An annular fan casing or nacelle 44 circumferentially surrounds at least a portion of fan assembly 14 and/or 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 posts 46. Additionally, at least a portion of nacelle 44 may extend over an exterior portion of core engine 16 to define a bypass airflow passage 48 therebetween.

FIG. 2 is a cross-sectional side view of an exemplary combustion section 26 of core engine 16 shown in FIG. 1. As shown in FIG. 2, combustion section 26 may generally include an annular combustor 50 having an annular inner liner 52, an annular outer liner 54, and a diaphragm 56, with diaphragm 56 extending radially between upstream ends of inner liner 52 and outer liner 54, respectively. In other embodiments of combustion section 26, combustion assembly 50 may be an annular can. Combustor 50 further includes a dome assembly 57 extending radially between inner liner 52 and outer liner 54 downstream of diaphragm 56. As shown in FIG. 2, the inner liner 52 is radially spaced from the outer liner 54 about the engine centerline 12 (FIG. 1) and defines a generally annular combustion chamber 62 therebetween. In particular embodiments, inner liner 52, outer liner 54, and/or dome assembly 54 may be at least partially or entirely formed of a metal alloy or a Ceramic Matrix Composite (CMC) material.

As shown in FIG. 2, the inner liner 52 and the outer liner 54 may be enclosed within an outer shell 64. The outer flow channels 66 of the diffuser cavity or plenum 84 may be defined around the inner liner 52 and/or the outer liner 54. Inner and outer liners 52, 54 may extend from a diaphragm 56 toward a turbine nozzle or inlet to HP turbine 28 (FIG. 1), thereby at least partially defining a hot gas path between combustor assembly 50 and HP turbine 28. The fuel nozzles 70 may extend at least partially through the membrane 56 to provide fuel 72, mix with air 82(a), and combust at the combustion chamber 62. In various embodiments, the baffle 56 includes a fuel-air mixing structure (e.g., a swirler assembly) attached thereto.

Still referring to FIG. 2, the inner liner 52 and the outer liner 54 each define one or more openings 105 through the liners 52, 54. The walled channel 100 is at least partially disposed within the opening 105. In various embodiments, the walled slot 100 extends at least partially into the combustion chamber 62. In other embodiments, walled slots 100 extend at least partially into the plenum 84. In still other embodiments, the walled channel 100 is approximately flush or smoothly transitioning with the liners 52, 54 to which the walled channel 100 is attached, and is disposed in the opening 105. The walled slot 100 generally defines a walled enclosure defining a first flow passage 111 therethrough from the pressure chamber 84 to the combustion chamber 62.

During operation of engine 10, as shown collectively in fig. 1 and 2, a volume of air, indicated schematically by arrow 74, enters engine 10 through nacelle 44 and/or an associated inlet 76 of fan assembly 14. As air 74 passes through fan blades 42, a portion of the air, schematically indicated by arrow 78, is channeled or conveyed into bypass airflow passage 48, and another portion of air 58, schematically indicated by arrow 80, is channeled or conveyed into LP compressor 22. As air 80 flows through LP compressor 22 and HP compressor 24 toward combustion section 26, air 80 is gradually compressed. As shown in FIG. 2, the compressed air that is now indicated schematically by arrow 82 flows into a diffuser cavity or plenum 84 of the combustion section 26. A plenum 84 generally surrounds inner liner 52 and outer liner 54, and is generally upstream of combustion chamber 62.

The compressed air 82 pressurizes a diffuser chamber 84. A first portion of the compressed air 82, indicated schematically by arrow 82(a), flows from the plenum 84 into the combustion chamber 62, where the compressed air 82 mixes with the fuel 72 and combusts, thus producing combustion gases, indicated schematically by arrow 86, within the combustor 50. In general, LP compressor 22 and HP compressor 24 provide more compressed air to plenum 84 than is required for combustion. Thus, the second portion of the compressed air 82, schematically indicated by arrow 82(b), may be used for various purposes other than combustion. For example, as shown in FIG. 2, compressed air 82(b) may be delivered into outer flow channel 66 to provide cooling to inner liner 52 and outer liner 54.

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

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

Referring to fig. 5-11, in various embodiments, contoured surface 103 may define a V-shape, a wave shape, a rib structure, or a vane structure. For example, referring to fig. 5-7, contoured surface 103 may be defined over 50% or more of inner surface 101 of walled channel 100. Contoured surface 103 may define a plurality of grooves 104. In one embodiment, such as generally provided with reference to fig. 5, the contoured surface 103 defines a plurality of grooves 104, the plurality of grooves 104 extending substantially perpendicular with respect to the flow of air 83 through the first flow channel 111 defined within the walled channel 100.

In another embodiment, such as generally provided with reference to fig. 6, the contoured surface 103 defines a plurality of grooves 104, the plurality of grooves 104 extending at an acute angle relative to the air 83 passing through the first flow channel 111. Alternatively, the contoured surface 103 defines a plurality of grooves 104, the plurality of grooves 104 extending at an acute angle relative to the walled channel 100 extending through the liners 52, 54. For example, in various embodiments, the groove 104 may extend at an acute angle greater than zero degrees and less than 90 degrees relative to the walled groove 100 and/or the inner surface 101 of the walled groove 100.

In yet another embodiment, such as generally provided with reference to fig. 7, contoured surface 103 defines a plurality of grooves 104, the plurality of grooves 104 extending substantially helically along inner surface 101. For example, the groove 104 may extend substantially along the span 107 of the inner surface 101.

Referring to fig. 5-11, a contoured surface 103 may be defined to protrude into the inner surface 101 of the walled channel 100. In other embodiments, the contoured surface 103 may be defined to extend from the inner surface 101 of the walled channel 100, such as into the first flow channel 111.

In one embodiment, such as generally provided with reference to fig. 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 10% of the span 107 from the cold end 16. Referring to fig. 8-11, contoured surface 103 may define a V-shape or wave shape at least partially along walled channel 100. In one embodiment, such as generally illustrated with reference to fig. 8-9, the contoured surface 103 defining the V-shape or wave 113 may be defined substantially inwardly toward the first flow channel 111. In another embodiment, such as generally provided with reference to fig. 10, the contoured surface 103 defining the V-shape or wave 113 may be defined substantially along the flow direction of the air 83 through the walled channel 100.

It should be understood that various embodiments of the contoured surface 103 that may define a V-shape or waveform 113 may define a triangular waveform, a square or rectangular or stepped waveform, a sinusoidal waveform, a sawtooth waveform, or other waveform. Additionally or alternatively, the V-shape or waveform 113 may be defined regularly (e.g., constant frequency) along the walled slot 100. In other embodiments, the V-shape or waveform 113 may be defined irregularly (e.g., variable, asymmetric, or irregular frequency) along the walled slide 100.

Embodiments with contoured surfaces 103 may generally achieve, promote, or increase turbulence in the flow of air 83 from the 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 combustion gases 86 to reduce production of nitrogen oxides (e.g., NOx), improve durability of combustor assembly 50 (e.g., improve durability of liners 52, 54), or both. As another example, the walled groove 100 including the contoured surface 103 may further improve mixing of the air 83 flow with the combustion gases 86 while mitigating penetration losses of the air 83 flow and the combustion gases 86 into the combustion chamber 62.

Referring to fig. 2-18, in various embodiments, the walled channel 100 may define the span 107 to be equal to or greater than the 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 (6x) the diameter 109 of the opening 105. In another embodiment, span 107 is less than or equal to four times (4x) diameter 109 of opening 105. In still other embodiments, the span 107 of the walled slot 100 may extend from the liners 52, 54 into the combustion chamber 62 at less than six times (6x) the diameter 109. Also in various embodiments, the span 107 of the walled slot 100 may extend from the liners 52, 54 into the combustion chamber 62 less than four times (4x) the diameter 109. For example, the walled channel 100 may extend substantially flush or with a smooth transition into the liner 52, 54 or into the plenum 84, additionally into the combustion chamber 62 (as in FIG. 4), and a portion of the walled channel 100 extending into the combustion chamber 62 may extend into the combustion chamber 62 less than six times the diameter 109 of the opening 105.

Referring now to fig. 12-18, in various embodiments, the walled groove 100 may further define the jet disruptor as a member 110 extending from the inner surface 101 at least partially across the first flow channel 111. For example, the member 110 may extend across the diameter 109 of the opening 105.

In one embodiment, such as generally provided with reference to fig. 12, the member 110 may extend across the first flow channel 111 approximately equidistant between the cold end 106 and the hot end 108 of the walled channel 100. In other embodiments, such as generally depicted in fig. 12-16, the member 110 may extend across the walled channel 100 within the liners 52, 54.

In other embodiments, such as generally provided with reference to fig. 12-16, the member 110 may extend across the first flow channel 111 within approximately 33% of the span 107 from the cold end 106. In still other embodiments, the member 110 may extend across the first flow channel 111 within approximately 33% of the span 107 from the hot end 108.

In another embodiment of the walled channel 100, such as generally provided with reference to fig. 14-17, the member 110 extends from the inner surface 101 (fig. 14-15) of the walled channel 100 substantially perpendicular (fig. 17) to a flow direction of the combustion gases 86 formed within the combustion chamber 62 (fig. 2). The components 110 may define jet disruptors that are disposed substantially perpendicular to a direction of cross-flow of the combustion gases 86 (i.e., perpendicular to the flow of the combustion gases 86 toward the turbines 28, 30). The walled channel 100 defining the member 110 as a jet disruptor member may divide the counter-rotating vortex pair (CVP) into more than two pairs (e.g., air 83(a) and air 83(b) in fig. 14), thereby adding additional swirl or wake from the air 83 flow to the jet of combustion gases 86. The additional swirl may cause lateral disturbances that may be further amplified or destabilized to achieve oscillations that define the flow of air 83 that dilutes the injection of 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 fig. 14-15 and 18, in yet another embodiment, the member 110 extends from the inner surface 101 of the walled channel 100 substantially co-directional (fig. 18) or parallel to the flow direction of the combustion gases 86 formed within the combustion chamber 62. Member 110 defining the jet disruptor may weaken the CVP, thereby improving or promoting diffusion of air 83 over liners 52, 54 to improve heat transfer (e.g., cooling) at liners 52, 54. As such, the components 110 disposed substantially co-directional with the flow of the combustion gases 86 may further improve the durability of the combustor assembly 50, for example, at the liners 52, 54.

Various embodiments of engine 10 and combustor assembly 50 may define a rich-burn combustor, wherein walled slots 100 may define dilution jets providing additional mixing of a mixture of air (e.g., air 83) and combustion gases (e.g., combustion gases 86) to improve or complete the combustion process. The walled slots 100 may further define dilution jets that further enable or enhance a combustion recirculation zone within the combustion chamber 62 to stabilize the flame therein. Still further, the walled slots 100 may define dilution injections that may quench the combustion gases 86 relatively quickly to minimize production of nitrogen oxides. Moreover, the various embodiments of combustor assembly 50 and walled slot 100 shown and described herein may enable tailoring of the distribution of combustion gas temperatures to improve the durability of components at or downstream of combustor assembly 50 (e.g., liners 52, 54, HP turbine 28).

Still further, the walled slot 100 may generally define the component 110 as a bluff body device to provide a jet disruptor to modify the counter-rotating vortex pair (CVP) formed in a cross-flow injection (JIC). For example, a portion of the air 83 provided by the walled groove 100 may define a CVP that is established relative to the JIC-defined flow of the combustion gases 86.

All or part of the burner assembly may be part of a single, unitary assembly and may be manufactured by any number of processes known to those skilled in the art. These manufacturing processes include, but are not limited to, those manufacturing processes referred to as "additive manufacturing" or "3D printing. Additionally, the combustor 50 may be constructed using any number or combination of casting, machining, welding, brazing, or sintering processes, including but not limited to liners 52, 54, walled channels 100, components 110, or combinations thereof. Moreover, the burner assembly may constitute one or more separate components mechanically joined (e.g., by using bolts, nuts, rivets or screws, or a welding or brazing process, or a combination thereof) or positioned in space to achieve substantially similar geometric, aerodynamic, or thermodynamic results as if fabricated or assembled into one or more components. Non-limiting examples of suitable materials include high strength steels, nickel and cobalt based alloys, and/or metallic or ceramic matrix composites, or combinations thereof.

Various embodiments of the walled channel 100 including the member 110 may define more than one cross-sectional area of the member 110, such as, but not limited to, a circular cross-section (as shown in fig. 16-18), a rectangular cross-section, an oval or racetrack shaped cross-section, an airfoil or tear-drop shaped cross-section, a polygonal cross-section, or an elliptical cross-section, or other suitable cross-section, or combinations thereof.

Additionally or alternatively, various embodiments of the walled channel 100 and/or the opening 105 through which the walled channel 100 is disposed may define more than one cross-sectional area, such as, but not limited to, a circular cross-section (as shown in fig. 2-18), a rectangular cross-section, an oval or racetrack shaped cross-section, an airfoil or tear drop shaped cross-section, a polygonal cross-section, or an elliptical cross-section, or other 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 scope of the patent rights to the invention is defined by the claims and may include other examples that occur readily to those skilled in the art. Such other examples are intended to be within the scope of the claims if the examples include structural elements that do not differ from the literal language of the claims, or if the examples include equivalent structural elements with insubstantial differences from the literal languages of the claims.

Further aspects of the invention are provided by the subject matter of the following clauses:

1. a combustor assembly for a gas turbine engine, the combustor assembly comprising: a liner defining a combustion chamber therein and a plenum surrounding the liner, wherein the liner comprises an opening, and wherein the liner comprises a walled groove disposed at least partially within the opening, and further wherein the walled groove comprises an inner surface, and wherein the inner surface comprises a jet disruptor.

2. The burner assembly according to any preceding claim, wherein the walled slot defines a span from the cold end proximate the plenum to the inner surface proximate a hot end of the combustion chamber, and wherein the jet disrupter comprises a contoured surface defined within approximately 33% of the span from the hot end.

3. A burner assembly according to any of the preceding items, wherein the contoured surface comprises a V-shape, a wave shape, a rib structure, or a vane structure.

4. The burner assembly of any preceding claim wherein the inner surface comprises a groove.

5. The burner assembly of any preceding claim wherein the groove defines a substantially helical shape.

6. The burner assembly of any preceding claim wherein the groove is defined substantially perpendicular to the inner surface.

7. The burner assembly of any preceding claim, wherein the jet disrupter comprises a member extending from the inner surface at least partially across the opening.

8. The burner assembly of any preceding claim wherein the member extends from the inner surface of the walled trough substantially perpendicular to a flow direction of combustion gases formed within the combustion chamber.

9. A burner assembly according to any preceding claim wherein the member extends from the inner surface of the walled trough substantially co-directionally with the direction of flow of combustion gases formed within the combustion chamber.

10. The burner assembly of any preceding claim wherein the walled slot extends at least partially into the combustion chamber.

11. The burner assembly of any preceding claim wherein the walled slot extends at least partially into the 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 groove disposed at least partially within the opening, and wherein the walled groove comprises an inner surface, and wherein the inner surface comprises a jet disruptor in a first flow channel defined within the walled groove.

13. The gas turbine engine of any preceding item, wherein the jet disrupter comprises a member extending across a diameter of the opening of the liner.

14. The gas turbine engine of any preceding item, wherein the component extends from the inner surface substantially perpendicular to a flow direction of combustion gases formed within the combustion chamber.

15. The gas turbine engine of any preceding claim, wherein the component extends from the inner surface substantially co-directionally with a flow direction of combustion gases formed within the combustion chamber.

16. The gas turbine engine of any preceding item, wherein the jet disrupter 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 any preceding item, wherein the jet disruptor is defined within 33% of the walled slot from the cold end or the hot end.

18. The gas turbine engine of any preceding item, wherein the walled groove is equal to or less than six times a diameter of the opening of the liner.

19. The gas turbine engine of any preceding item, wherein the jet disruptor comprises a groove defined at the inner surface of the walled groove.

20. The gas turbine engine of any preceding item, wherein the jet disrupter comprises a V-shape, a wave shape, a rib structure, or a vane structure at the walled trough.