Patent Application
20190056108
The present subject matter relates generally to turbine engine combustion assemblies.
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.
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 comprising a fuel nozzle and an annular shroud. The fuel nozzle comprises a centerbody extended along a lengthwise direction. The fuel nozzle defines a nozzle centerline extended through the centerbody of the fuel nozzle along the lengthwise direction. The fuel nozzle defines a plurality of exit openings in circumferential arrangement on the centerbody relative to the nozzle centerline. The annular shroud surrounds the centerbody of the fuel nozzle. At least a portion of the shroud defines a contoured structure defining a waveform.
In one embodiment, the waveform is triangle, sinusoidal, or box.
In another embodiment, the contoured structure of the shroud extends along the lengthwise direction.
In various embodiments, the combustor assembly defines a second reference plane along the radial direction from the nozzle centerline at a position along the lengthwise direction. The plurality of exit openings on the centerbody is defined at least approximately along the second reference plane. In one embodiment, the combustor assembly defines a first reference plane along the radial direction from the nozzle centerline at a position along the lengthwise direction. The shroud and the centerbody each define a downstream-most end approximately co-planar at the first reference plane. In another embodiment, the combustor assembly defines a third reference plane along the radial direction from the nozzle centerline at a position along the lengthwise direction. The third reference plane is defined downstream of the second reference plane along the lengthwise direction. A downstream-most end of the shroud is defined at least approximately at the third reference plane.
In still various embodiments, the contoured structure of the shroud extends at least partially along a radial direction relative to the nozzle centerline. In one embodiment, the contoured structure of the shroud further extends at least partially along a circumferential direction relative to the nozzle centerline.
In still yet various embodiments, the exit openings define two or more cross sectional areas through the centerbody different from one another. In one embodiment, the plurality of exit openings defines a first exit opening of a first cross sectional area and a second exit opening of a second cross sectional area different from the first cross sectional area.
Another aspect of the present disclosure is directed to a gas turbine engine defining an axial centerline, a radial direction extended therefrom, and a circumferential direction around the axial centerline. The gas turbine engine includes a combustor assembly disposed generally concentric to the axial centerline of the gas turbine engine. The combustor assembly includes a plurality of fuel nozzles disposed in circumferential arrangement around the axial centerline. Each fuel nozzle comprises a centerbody extended along a lengthwise direction and defining a nozzle centerline therethrough, and wherein an annular shroud is defined around the centerbody, and wherein at least a portion of the shroud defines a contoured structure defining a waveform, and wherein each fuel nozzle defines a plurality of exit openings in circumferential arrangement on the centerbody relative to the nozzle centerline.
In various embodiments of the gas turbine engine, the combustor assembly defines a second reference plane along the radial direction from the nozzle centerline at a position along the lengthwise direction. The plurality of exit openings on the centerbody is defined at least approximately along the second reference plane. In one embodiment, the combustor assembly defines a first reference plane along the radial direction from the nozzle centerline at a position along the lengthwise direction, and the shroud and the centerbody each define a downstream-most end approximately co-planar at the first reference plane. The first reference plane relative to the second reference plane defines a first immersion depth of the fuel nozzle. In another embodiment, the combustor assembly defines a third reference plane along the radial direction from the nozzle centerline at a position along the lengthwise direction, and wherein the third reference plane is defined downstream of the second reference plane along the lengthwise direction. A downstream-most end of the shroud is defined at least approximately at the third reference plane. The third reference plane relative to the second reference plane defines a second immersion depth of the fuel nozzle.
In one embodiment of the gas turbine engine, the waveform is triangle, sinusoidal, or box.
In another embodiment, the contoured structure of the shroud extends along the lengthwise direction.
In various embodiments, the contoured structure of the shroud extends at least partially along a radial direction relative to the nozzle centerline. In one embodiment, the contoured structure of the shroud further extends at least partially along a circumferential direction relative to the nozzle centerline.
In another embodiment of the gas turbine engine, the combustor assembly defines a first annular shroud and a second annular shroud, in which the first annular shroud defines a first waveform different from a second waveform of the second annular shroud.
In still another embodiment of the gas turbine engine, the fuel nozzle is configured to provide a flow of fuel through the centerbody and egressing from the exit openings into a combustion chamber of the combustor assembly, and wherein the contoured structure of the annular shroud provides a circumferentially asymmetric flame relative to the axial centerline within the combustion chamber.
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.
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:
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.
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 from a given location or feature toward “upstream end 99” or toward “downstream end 98”, respectively, as provided in the figures.
Embodiments of a combustor assembly for a gas turbine engine including a fuel nozzle and annular shroud are generally provided that may desirably alter the heat release characteristics of each fuel nozzle and annular shroud combination to mitigate undesired combustion dynamics. The annular shroud generally defines a mixer surrounding each fuel nozzle, such as defining a flow passage between one or more main fuel injection openings in the fuel nozzle and a flow of air from a diffuser cavity to a combustion chamber.
The combustor assembly including the embodiments of the fuel nozzle and annular shroud 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 and annular shroud 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,
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
As shown in
As shown in
During operation of the engine 10, as shown in
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
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
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 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 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.
Referring back to
In another embodiment, such as generally provided in
Regarding
Referring now to
Referring now to
The fuel nozzle 70 defines a reference plane from the nozzle centerline 11 and the radial direction RR along the nozzle centerline 11. The shroud 110 defines a downstream-most end 111 and the centerbody 105 of the fuel nozzle 70 defines a downstream-most end 106. Referring to
Referring still to
Referring now to
In other embodiments, the third reference plane 118 is defined downstream along the lengthwise direction L of the first reference plane 114. The downstream-most end 111 of the shroud 110 defines a distance 117 along the lengthwise direction L from the second reference plane 116 defining the planar location of the plurality of exit openings 107 greater than the distance 115 of the first reference plane 114 to the second reference plane 116. For example, the distance 117 along the lengthwise direction L from the downstream-most end 111 of the shroud 110 is greater than the distance 115 along the lengthwise direction L from the downstream-most end 106 of the centerbody 105.
It should be appreciated that the second reference plane 116 may be defined through a center point of the plurality of exit openings 107. However, in other embodiments, the second reference plane 116 may be defined relative to a perimeter or another geometric feature of the exit openings 107. In still various embodiments, the distance 117 of the downstream-most end 111 of the shroud 110 may be greater than the distance 115 of the downstream-most end 106 of the centerbody 105.
Referring to
In various embodiments, the engine 10 defines a first fuel nozzle and a second fuel nozzle. The first fuel nozzle defines the first immersion depth (i.e., the distance 115, such as generally provided in
In still various embodiments, the first fuel nozzle and the second fuel nozzle may each define one or more of the contoured structure 113 generally described and shown in regard to
For example, in various embodiments, the combustor assembly 50 may define a plurality of the fuel nozzles 70 in which up to half of the total plurality of fuel nozzles 70 defines the shroud 110 relative to the exit openings 107 of the first immersion depth (e.g., the distance 115, such as generally provided in
In still various embodiments, the plurality of fuel nozzles 70 may dispose the shroud 110 embodiments as generally provided in regard to
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 contoured structure 113 of the annular shroud 110 provides a circumferentially asymmetric flame within the combustion chamber 62 relative to the axial centerline 12.
All or part of the combustor assembly 50, fuel nozzle 70, and annular shroud 110 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 and the shroud 110. 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.
