Patent Description:
A gas turbine engine includes a turbine that is driven by combustion of a combustible fuel within a combustor of the engine. A turbine engine utilizes a fuel injector assembly to inject the combustible fuel into the combustor. The fuel injector assembly can mix the fuel with air prior to injection in order to achieve efficient combustion.

A prior art gas turbine engine is disclosed in <CIT>, wherein a combustion chamber for a gas turbine system is described with at least one premix burner which has a burner outlet for supplying a fuel-oxidant mixture into the combustion chamber and at least one gas-permeable perforated section in a wall of the combustion chamber and / or a front plate of the premix burner. The premix burner is in fluid communication with a means for supplying a combustible gas or gas mixture to the combustion chamber.

A prior art fuel injector for a combustion engine is disclosed in <CIT>, wherein the fuel injector for a combustion engine includes an injector head including a nozzle, a premixer, and a distributor structured to distribute a plurality of different fuels to different sets of fueling orifices in the premixer. A pilot assembly of the fuel injector is coupled to the premixer and includes a first fueling passage for a first fuel and a second fueling passage for a second fuel. Multiple sets of fueling orifices are positioned within the fuel injector, the fueling orifice sets being selectively connectable to a plurality of different fuel supplies, and both located and sized so as to accommodate a wide range of flow rates to enable a combustion engine coupled with the fuel injector to operate on fuels having a range of Wobbe indices and compositions.

A prior art fuel delivery system for a gas turbine engine is disclosed in <CIT>, wherein the fuel delivery system for a turbine engine has at least one fuel injector having an upstream orifice arrangement that produces a first pressure drop of flowing fuel and a downstream orifice arrangement that produces a second pressure drop of the flowing fuel. The upstream orifice arrangement or the downstream orifice arrangement includes a noncylindrical orifice. A prior art turbine engine is disclosed in <CIT>, wherein a combustor comprises a fuel injector provided with a gas flow restrictor.

According to the present invention there is provided a turbine engine as defined in the appended independent apparatus claim. Further preferable features of the turbine engine of the present invention are defined in the appended dependent apparatus claims.

A full and enabling disclosure, 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:.

Aspects of the disclosure described herein are generally directed to a combustion section for a turbine engine. The combustion section including a fuel injector defining a fuel inlet for the combustion section. The fuel injector having at least one fuel channel fluidly coupled to a fuel circuit. A flow restrictor can be provided within the fuel channel and includes a set of fuel orifices. The fuel within the fuel injector can be any suitable fuel. As a non-limiting example, the fuel can contain hydrogen (hereinafter, hydrogen-containing fuel) that is mixed with at least one airflow within a fuel-air mixing assembly downstream of the fuel injector. Hydrogen-containing fuel typically has a wider flammable range and a faster burning velocity than traditional fuels, such as petroleum-based fuels or petroleum and synthetic fuel blends. The burn temperatures for hydrogen-containing fuel can be higher than the burn temperatures of traditional fuel, such that existing engine designs for traditional fuels would not be capable of operating under the heightened temperatures. The fuel injector, as described herein, provides for a fuel injector with stages that restrict the flow of fuel through the fuel channel (e.g., the flow restrictor). This, in turn, can affect the volume and coverage of the hydrogen-containing fuel as the fuel enters the fuel-air mixing assembly.

For purposes of illustration, the present disclosure will be described with respect to the turbine for an aircraft turbine engine. It will be understood, however, that aspects of the disclosure described herein are not so limited and may have general applicability within an engine, including compressors, power generation turbines, as well as in non-aircraft applications, such as other mobile applications and non-mobile industrial, commercial, and residential applications.

Reference will now be made in detail to the combustor architecture, and in particular the fuel injector and swirler for providing fuel to the combustor located within a turbine engine, one or more examples of which are illustrated in the accompanying drawings. Like or similar designations in the drawings and description have been used to refer to like or similar parts of the disclosure.

The terms "forward" and "aft" refer to relative positions within a turbine engine or vehicle, and refer to the normal operational attitude of the turbine engine or vehicle. For example, with regard to a turbine engine, forward refers to a position closer to an engine and aft refers to a position closer to an engine nozzle or exhaust.

As used herein, the term "upstream" refers to a direction that is opposite the fluid flow direction, and the term "downstream" refers to a direction that is in the same direction as the fluid flow. The term "fore" or "forward" means in front of something and "aft" or "rearward" means behind something. For example, when used in terms of fluid flow, fore/forward can mean upstream and aft/rearward can mean downstream.

The term "fluid" may be a gas or a liquid. The term "fluid communication" means that a fluid is capable of making the connection between the areas specified.

Additionally, as used herein, the terms "radial" or "radially" refer to a direction away from a common center. For example, in the overall context of a turbine engine, radial refers to a direction along a ray extending between a center longitudinal axis of the engine and an outer engine circumference.

Furthermore, as used herein, the term "set" or a "set" of elements can be any number of elements, including only one.

All directional references (e.g., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, front, back, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, forward, aft, etc.) are used only for identification purposes to aid the reader's understanding of the present disclosure, and should not be construed as limiting, particularly as to the position, orientation, or use of aspects of the disclosure described herein. Connection references (e.g., attached, coupled, connected, and joined) are to be construed broadly and can include intermediate members between a collection of elements and relative movement between elements unless otherwise indicated. As such, connection references do not necessarily infer that two elements are directly connected and in fixed relation to one another. The exemplary drawings are for purposes of illustration only and the dimensions, positions, order and relative sizes reflected in the drawings attached hereto can vary. Furthermore, as used herein, the term "set" or a "set" of elements can be any number of elements, including only one.

Accordingly, a value modified by a term or terms, such as "about", "approximately", "generally", and "substantially", are not to be limited to the precise value specified. For example, the approximating language may refer to being within a <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM> percent margin in either individual values, range(s) of values and/or endpoints defining range(s) of values.

<FIG> is a schematic view of a turbine engine <NUM>. As a non-limiting example, the turbine engine <NUM> can be used within an aircraft. The turbine engine <NUM> can include, at least, a compressor section <NUM>, a combustion section <NUM>, and a turbine section <NUM>. A drive shaft <NUM> rotationally couples the compressor and turbine sections <NUM>, <NUM>, such that rotation of one affects the rotation of the other, and defines a rotational axis <NUM> for the turbine engine <NUM>.

The compressor section <NUM> can include a low-pressure (LP) compressor <NUM>, and a high-pressure (HP) compressor <NUM> serially fluidly coupled to one another. The turbine section <NUM> can include an HP turbine <NUM>, and an LP turbine <NUM> serially fluidly coupled to one another. The drive shaft <NUM> can operatively couple the LP compressor <NUM>, the HP compressor <NUM>, the HP turbine <NUM> and the LP turbine <NUM> together. Alternatively, the drive shaft <NUM> can include an LP drive shaft (not illustrated) and an HP drive shaft (not illustrated). The LP drive shaft can couple the LP compressor <NUM> to the LP turbine <NUM>, and the HP drive shaft can couple the HP compressor <NUM> to the HP turbine <NUM>. An LP spool can be defined as the combination of the LP compressor <NUM>, the LP turbine <NUM>, and the LP drive shaft such that the rotation of the LP turbine <NUM> can apply a driving force to the LP drive shaft, which in turn can rotate the LP compressor <NUM>. An HP spool can be defined as the combination of the HP compressor <NUM>, the HP turbine <NUM>, and the HP drive shaft such that the rotation of the HP turbine <NUM> can apply a driving force to the HP drive shaft which in turn can rotate the HP compressor <NUM>.

The compressor section <NUM> can include a plurality of axially spaced stages. Each stage includes a set of circumferentially-spaced rotating blades and a set of circumferentially-spaced stationary vanes. The compressor blades for a stage of the compressor section <NUM> can be mounted to a disk, which is mounted to the drive shaft <NUM>. Each set of blades for a given stage can have its own disk. The vanes of the compressor section <NUM> can be mounted to a casing which can extend circumferentially about the turbine engine <NUM>. It will be appreciated that the representation of the compressor section <NUM> is merely schematic and that there can be any number of stages. Further, it is contemplated, that there can be any other number of components within the compressor section <NUM>.

Similar to the compressor section <NUM>, the turbine section <NUM> can include a plurality of axially spaced stages, with each stage having a set of circumferentially-spaced, rotating blades and a set of circumferentially-spaced, stationary vanes. The turbine blades for a stage of the turbine section <NUM> can be mounted to a disk which is mounted to the drive shaft <NUM>. Each set of blades for a given stage can have its own disk. The vanes of the turbine section can be mounted to the casing in a circumferential manner. It is noted that there can be any number of blades, vanes and turbine stages as the illustrated turbine section is merely a schematic representation. Further, it is contemplated, that there can be any other number of components within the turbine section <NUM>.

The combustion section <NUM> can be provided serially between the compressor section <NUM> and the turbine section <NUM>. The combustion section <NUM> can be fluidly coupled to at least a portion of the compressor section <NUM> and the turbine section <NUM> such that the combustion section <NUM> at least partially fluidly couples the compressor section <NUM> to the turbine section <NUM>. As a non-limiting example, the combustion section <NUM> can be fluidly coupled to the HP compressor <NUM> at an upstream end of the combustion section <NUM> and to the HP turbine <NUM> at a downstream end of the combustion section <NUM>.

During operation of the turbine engine <NUM>, ambient or atmospheric air is drawn into the compressor section <NUM> via a fan (not illustrated) upstream of the compressor section <NUM>, where the air is compressed defining a pressurized air. The pressurized air can then flow into the combustion section <NUM> where the pressurized air is mixed with fuel and ignited, thereby generating combustion gases. Some work is extracted from these combustion gases by the HP turbine <NUM>, which drives the HP compressor <NUM>. The combustion gases are discharged into the LP turbine <NUM>, which extracts additional work to drive the LP compressor <NUM>, and the exhaust gas is ultimately discharged from the turbine engine <NUM> via an exhaust section (not illustrated) downstream of the turbine section <NUM>. The driving of the LP turbine <NUM> drives the LP spool to rotate the fan (not illustrated) and the LP compressor <NUM>. The pressurized airflow and the combustion gases can together define a working airflow that flows through the fan, compressor section <NUM>, combustion section <NUM>, and turbine section <NUM> of the turbine engine <NUM>.

<FIG> depicts a schematic cross-sectional view of a generic combustion section suitable for use as combustion section <NUM> located between a compressor section <NUM> and a turbine section <NUM> of a turbine engine. The combustion section <NUM> can include an annular arrangement of fuel injectors <NUM> each connected to a combustor <NUM>. It should be appreciated that the annular arrangement of fuel injectors <NUM> can be one or multiple fuel injectors, which can each have different characteristics. Only a single fuel injector of the annular arrangement of fuel injectors <NUM> is illustrated is for illustrative purposes only and is not intended to be limiting. The combustor <NUM> can have a can, can-annular, or annular arrangement depending on the type of turbine engine in which the combustor <NUM> is located. In a non-limiting example, an annular arrangement is illustrated and disposed within a casing <NUM>. The combustor <NUM> can include an annular combustor liner <NUM>, a dome assembly <NUM> including a dome wall <NUM> which together define a combustion chamber <NUM> about a longitudinal axis (LA). A compressed air passageway <NUM> can be defined at least in part by both annular combustor liner <NUM> and the casing <NUM>. The fuel injector of the annular arrangement of fuel injectors <NUM> is fluidly coupled to the combustion chamber <NUM>. A passage can fluidly connect the compressed air passageway <NUM> and the combustor <NUM>. The passage can be defined by at least one set of dilution openings <NUM> located in the annular combustor liner <NUM>.

The fuel injector of the annular arrangement of fuel injectors <NUM> can be coupled to and disposed within the dome assembly <NUM> upstream of a flare cone <NUM> to define a fuel outlet <NUM>. The fuel injector of the annular arrangement of fuel injectors <NUM> can include a fuel inlet <NUM> that can be adapted to receive a flow of a fuel (F) (e.g., a hydrogen-containing fuel) and a linear fuel passageway <NUM> extending between the fuel inlet <NUM> and the fuel outlet <NUM>. A swirler <NUM> can be provided at a dome inlet <NUM> to swirl incoming air in proximity to fuel (F) exiting the fuel injector of the annular arrangement of fuel injectors <NUM> and provide a homogeneous mixture of air and fuel entering the combustor <NUM>.

The annular combustor liner <NUM> can be defined by a wall <NUM> having an outer surface <NUM> and an inner surface <NUM> at least partially defining the combustion chamber <NUM>. The wall <NUM> can be made of one continuous monolithic portion or be multiple monolithic portions assembled together to define the annular combustor liner <NUM>. By way of non-limiting example, the outer surface <NUM> can define a first piece of the wall <NUM> while the inner surface <NUM> can define a second piece of the wall <NUM> that when assembled together form the annular combustor liner <NUM>. As described herein, the wall <NUM> includes the at least one set of dilution openings <NUM>. It is further contemplated that the annular combustor liner <NUM> can be any type of annular combustor liner <NUM>, including but not limited to a double walled liner or a tile liner. An igniter <NUM> can be provided at the wall <NUM> and fluidly coupled to the combustion chamber <NUM>.

During operation, compressed air (C) can flow from the compressor section <NUM> to the combustor <NUM> through the compressed air passageway <NUM>. The at least one set of dilution openings <NUM> in the annular combustor liner <NUM> allow passage of at least a portion of the compressed air (C), the portion defining a dilution airflow (D), from the compressed air passageway <NUM> to the combustion chamber <NUM>.

Some compressed air (C) can be mixed with the fuel (F) from the fuel injector of the annular arrangement of fuel injectors <NUM> and upon entering the combustor <NUM> are ignited within the combustion chamber <NUM> by one or more igniters <NUM> to generate combustion gas (G). The combustion gas (G) is mixed using the dilution airflow (D) supplied through the at least one set of dilution openings <NUM>, and mixes within the combustion chamber <NUM>, after which the combustion gas (G) flows through a combustor outlet <NUM> and exits into the turbine section <NUM>.

<FIG> is a cross-sectional side view of a fuel injector <NUM> suitable for use as at least one of the fuel injectors of the annular array of fuel injectors <NUM> of <FIG>. The fuel injector <NUM> can include a fuel channel <NUM> configured to receive the flow of the fuel (F) from the fuel inlet <NUM> (<FIG>). The fuel injector <NUM> can include a wall <NUM> provided along a distal end of the fuel injector <NUM>. A first set of fuel orifices <NUM> can be provided along and extend through at least a portion of the wall <NUM> and define an outlet of the fuel injector <NUM> for the fuel (F) from the fuel injector <NUM>.

A flow restrictor <NUM> can be provided within a portion of the fuel channel <NUM>. As a non-limiting example, the flow restrictor <NUM> can be provided upstream of the first set of fuel orifices <NUM>. The flow restrictor <NUM> can span the entire fuel channel <NUM> and oppose the flow of the fuel (F). In other words, the flow restrictor <NUM> can act as an impedance or restriction point of the flow of fuel (F) within the flow restrictor <NUM>. In order for the fuel (F) to flow out of the first set of fuel orifices <NUM>, the fuel (F) must first flow through the flow restrictor <NUM>. The flow restrictor <NUM> can include a second set of fuel orifices <NUM> within the flow restrictor <NUM>. As a non-limiting example, both the flow restrictor <NUM> and the portion of the fuel channel <NUM> that the flow restrictor <NUM> is provided within can be circular. As such, the flow restrictor <NUM> can be a circumferential flow restrictor. The second set of fuel orifices <NUM> can be circumferentially and radially spaced throughout the flow restrictor <NUM> and extend axially through the entire flow restrictor <NUM> with respect to a centerline axis <NUM> of the flow restrictor <NUM>. In other words, the second set of fuel orifices <NUM> can form multiple radially-spaced rows of circumferentially-spaced orifices. Alternatively, the flow restrictor <NUM> can be any suitable shape that corresponds to the portion of the fuel injector <NUM> that the flow restrictor <NUM> is provided within. As such, the flow restrictor <NUM> and the fuel injector <NUM> can take any suitable shape. As a non-limiting example, the fuel injector <NUM> can include a venturi with the flow restrictor <NUM> provided within the venturi. The flow restrictor <NUM> can increase or decrease in cross-sectional area to conform to the shape of the venturi. Similarly, the first set of fuel orifices <NUM> can be circumferentially or radially spaced with respect to one another. It will be appreciated that each orifice of the first set of fuel orifices <NUM> and the second set of fuel orifices <NUM> can be sized or shaped any suitable size or shape. Further, there can be any number of one or more orifices of the first set of fuel orifices <NUM> and one or more orifices of the second set of fuel orifices <NUM>.

The flow restrictor <NUM> can span across the fuel channel <NUM> in any suitable direction. As a non-limiting example, the flow restrictor <NUM> can extend perpendicularly, as illustrated, or non-perpendicularly across the fuel channel <NUM> with respect to the centerline axis <NUM>. Further, the flow restrictor <NUM> can have any suitable length between radially opposing walls of the fuel channel <NUM>. As a non-limiting example, the length of the flow restrictor <NUM> can be equal to a diameter of the fuel channel <NUM>. Alternatively, the length of the flow restrictor <NUM> can be non-equal to the diameter of the fuel channel <NUM>. The flow restrictor <NUM> can have any suitable shape, size or form. As a non-limiting example, the flow restrictor <NUM> can be a planar flow restrictor that extends linearly across the fuel channel <NUM> when viewed from a plane parallel to the centerline axis <NUM> and intersecting the flow restrictor <NUM>. Alternatively, the flow restrictor <NUM> can be a non-planar flow restrictor that extends non-linearly across the fuel channel <NUM> when viewed from a plane parallel to the centerline axis <NUM> and intersecting the flow restrictor <NUM> (e.g., the flow restrictor <NUM> can be formed as an undulating wave). It will be further appreciated that the flow restrictor <NUM> can be symmetric or non-symmetric about a plane parallel to the centerline axis <NUM> and intersecting the flow restrictor <NUM>.

During operation, the flow of the fuel (F) can flow through the second set of fuel orifices <NUM>, into a cavity <NUM>, and through the first set of fuel orifices <NUM>. The fuel (F) can then flow into the combustor <NUM> (<FIG>). When entering the combustor <NUM>, the fuel (F) can mix with a flow of air (e.g., the compressed air (C)). The fuel (F) and the flow of air can each include acoustic oscillations (e.g., the acoustic oscillations of the hydrogen-containing fuel), which can interact with one another to generate combustion dynamics. As used herein, the term "combustion dynamics" or iterations thereof, can refer to the generation of acoustic pressure oscillations that occur within the combustor. As a non-limiting example, acoustic pressure oscillations can occur within the fuel injector <NUM> or the fuel inlet <NUM> (<FIG>), which can increase or compound with combustion dynamics within the combustor <NUM>. It is contemplated that the mitigation, elimination, or control of the combustion dynamics can result in a greater efficiency or longer life cycle of the turbine engine <NUM>.

As the fuel (F) flows through the fuel injector <NUM>, two stages of pressure drops of the fuel (F) can be experienced. The first pressure drop being across the second set of fuel orifices <NUM>, the second pressure drop being across the first set of fuel orifices <NUM>. In other words, the fuel (F) can have a first pressure upstream of the second set of fuel orifices <NUM>, a second pressure downstream of the second set of fuel orifices <NUM> and within the cavity <NUM>, and a third pressure downstream of the first set of fuel orifices <NUM>. The pressure change from the first pressure to the second pressure can define the first pressure drop, while the pressure change from the second pressure to the third pressure can define the second pressure drop. As a non-limiting example, the second pressure drop across the second set of fuel orifices <NUM> can be larger than the first pressure drop across the first set of fuel orifices <NUM> such that combustion dynamics occurring in the combustion chamber <NUM> can propagate upstream of the first set of fuel orifices <NUM> and into the cavity <NUM>, but not through the flow restrictor <NUM>. In other words, the two pressure drops can be used to locate or isolate combustion dynamics in the fuel injector <NUM> or the fuel inlet that the fuel injector is a coupled to (e.g., the fuel inlet <NUM> of <FIG>) from the combustion dynamics within the combustion chamber downstream of the fuel injector <NUM>. This ultimately reduces or otherwise controls the overall combustion dynamics of the combustion section.

The acoustic oscillations of the fuel (F) can be at least partially dependent on acoustic impedance of the fuel injector <NUM>, which is a function of the two pressure drops and the volume of the cavity <NUM>. As a non-limiting example, the acoustic impedance can be controlled by changing the location, number, sizing, or formation of the second set of fuel orifices <NUM> or the first set of fuel orifices <NUM>, or by changing a volume of the cavity <NUM>. As a non-limiting example, the axial length, with respect to the centerline axis <NUM>, or volume of the cavity <NUM> can be sized with respect to the acoustic oscillation of the fuel (F), such that the acoustic oscillation of the fuel (F) within the cavity <NUM> and downstream of the first set of fuel orifices <NUM> can offset or counteract the acoustic pressure oscillations of the compressed air (C), which can ultimately reduce or otherwise control the overall combustion dynamics in the combustion chamber <NUM> (<FIG>). As a non-limiting example, the axial length of the cavity <NUM> can be a quarter of the wavelength of the acoustic oscillation of the fuel (F). As such, the fuel injector <NUM> can be further defined as an acoustic resonator that can be used to mitigate combustion dynamics within the combustion chamber and the fuel injector <NUM>.

Further, the fuel injector <NUM> can be used to distribute the fuel (F) in a desired fashion before the fuel (F) enters the combustion chamber <NUM>. As a non-limiting example, the geometric configuration of the flow restrictor <NUM> (e.g., the number, position, and sizing of the second set of fuel orifices <NUM>) and the positioning of the first set of fuel orifices <NUM> can be used to generate a profile of the fuel (F) as the fuel (F) enters the combustion chamber <NUM>. The profile of the fuel (F) can determine a shape of the flame or flame shape after the fuel (F) is mixed with the compressed air (C) and ignited within the combustor <NUM>. The flame shape can directly affect the combustor exit velocity and temperature profile of the combustion gases exiting the combustion chamber <NUM> (e.g., after combustion has occurred) , the pollutant emissions, and the combustion dynamics within the combustor <NUM>.

<FIG> is a cross-sectional side view of a fuel injector <NUM> for use in a turbine engine in accordance with the claims and which is suitable for use as at least one of the fuel injectors of the annular array of fuel injectors <NUM> of <FIG>. The fuel injector <NUM> is similar to the fuel injector <NUM>, therefore, like parts will be identified with like numerals increased to the <NUM> series, with it being understood that the description of the like parts of the fuel injector <NUM> applies to the fuel injector <NUM> unless otherwise noted.

The fuel injector <NUM> includes a first fuel channel <NUM> that terminates at a distal end of the fuel injector <NUM> defined by a wall <NUM>. The first fuel channel <NUM> is configured to receive the fuel (F). A first set of fuel orifices <NUM> can be provided within and extend through the wall <NUM> to define an outlet of the fuel injector <NUM>. A flow restrictor <NUM> can be provided within the first fuel channel <NUM> upstream of the first set of fuel orifices <NUM> and be defined by a centerline axis <NUM>. The flow restrictor <NUM> can include a second set of fuel orifices <NUM> fluidly coupled to the fuel (F), with at least a portion of the second set of fuel orifices <NUM> being directly fluidly coupled to a cavity <NUM>.

The flow restrictor <NUM> is similar to the flow restrictor <NUM>, except, the flow restrictor <NUM> is not formed as a planar flow restrictor with respect to a plane normal to the centerline axis <NUM> and intersecting the flow restrictor <NUM>. The flow restrictor <NUM> instead includes a first portion including a first subset <NUM> of the second set of fuel orifices <NUM>, and a second portion including a second subset <NUM> of the second set of fuel orifices <NUM>. The second portion can extend axially outwardly, in a first direction, from a plane normal to the centerline axis <NUM> and intersecting the first portion. The second portion can define a first protrusion <NUM>. The flow restrictor <NUM> can further include a third portion extending axially outwardly from the plane in a second direction, opposite the first direction, to define a second protrusion <NUM>. The second protrusion <NUM> can confront, contact, be coupled to, or integrally formed with a portion of the fuel injector <NUM> including the first set of fuel orifices <NUM>. A second fuel channel <NUM> can be formed by the first protrusion <NUM> and the second protrusion <NUM> and extend axially through the flow restrictor <NUM>. As such, a first subset <NUM> of the first set of fuel orifices <NUM> can be fluidly coupled to the first fuel channel <NUM>, while a second subset <NUM>, different from the first subset <NUM>, of the first set of fuel orifices <NUM> can be fluidly coupled to the second fuel channel <NUM>. The cavity <NUM> and the second fuel channel <NUM> can form concentric circles or any other suitable shape where the cavity <NUM> envelopes the second fuel channel <NUM> when viewed in a plane normal to centerline axis <NUM> and intersecting the cavity <NUM> and the second fuel channel <NUM>. As a non-limiting example, an axial length of the second fuel channel <NUM> can be larger than the axial length of the cavity <NUM>. It will be appreciated, however, that the axial length of the second fuel channel <NUM> can be smaller than the axial length of the cavity <NUM>. As such, the first protrusion <NUM> can be formed as a recess within the flow restrictor <NUM> such that the fuel orifices of the second set of fuel orifices <NUM> provided on the first protrusion <NUM> are downstream of the remaining fuel orifices of the second set of fuel orifices <NUM>.

The fuel injector <NUM> can be defined by a first segment <NUM> and a second segment <NUM> that together define the fuel channel of the fuel injector <NUM> (e.g., the combination of the first fuel channel <NUM> and the second fuel channel <NUM>). The first segment <NUM> can be defined by the space between the first set of fuel orifices <NUM> and the second set of fuel orifices <NUM> fluidly coupled to the cavity <NUM>. The second segment <NUM> can be defined by the space between the first set of fuel orifices <NUM> and the fuel orifice of the second set of fuel orifices <NUM> fluidly coupled to the second fuel channel <NUM>. The first segment <NUM> can envelope or at least partially surround the second segment <NUM>. As a non-limiting example, the fuel injector <NUM> can be a tubular fuel injector such that the first segment <NUM> circumscribes the second segment <NUM> when viewed along a plane normal to the centerline axis <NUM> and intersecting the first segment <NUM> and the second segment <NUM>. An outlet of the first segment <NUM> can be defined by the first subset <NUM> of the first set of fuel orifices <NUM>. An inlet of the first segment <NUM> can be defined by the first subset <NUM> of the second set of fuel orifices <NUM>. An outlet of the second segment <NUM> can be defined by the second subset <NUM> of the first set of fuel orifices <NUM>. An inlet of the second segment <NUM> can be defined by the second subset <NUM>, different from the first subset <NUM>, of the second set of fuel orifices <NUM>. The inlet of the first segment <NUM> can be spaced from the inlet of the second segment <NUM>. The fuel orifices that define the outlets and inlets of the first segment <NUM> and the second segment <NUM> can be varying sizes or shapes, or the same size or shape with respect to one another. The volume or axial length with respect to the centerline axis <NUM> of the first segment <NUM> and the second segment <NUM> can be sized to mitigate the combustion dynamics within the fuel injector <NUM>. Although two segments are illustrated, it will be appreciated that the flow restrictor <NUM> can partition the fuel injector <NUM> into any number of concentric or non-concentric segments.

The second set of fuel orifices <NUM>, similar to the second set of fuel orifices <NUM> of <FIG>, can be circumferentially and radially spaced about the flow restrictor <NUM>. The second set of fuel orifices <NUM>, however, can further be axially spaced about the flow restrictor <NUM>. As illustrated, at least a portion of the second set of fuel orifices <NUM> can be provided on a forward or upstream portion of the first protrusion <NUM>, while the remaining orifices of the second set of fuel orifices <NUM> can be provided on a portion of the flow restrictor <NUM> downstream of the first protrusion <NUM>. The fuel orifices of the second set of fuel orifices <NUM> that are provided on the first protrusion <NUM> can be directly fluidly coupled to the second fuel channel <NUM>, while the remaining orifices of the second set of fuel orifices <NUM> can be directly fluidly coupled to the cavity <NUM>.

Each orifice of the second set of fuel orifices <NUM> can be defined by a cross-sectional area with respect to a plane normal to the centerline axis <NUM> and intersecting the respective fuel orifice of the second set of fuel orifices <NUM>. As illustrated, the fuel orifices of the second set of fuel orifices <NUM> provided on the first protrusion <NUM> can be defined by a first cross-sectional area, while the cross-sectional area of the fuel orifices of the second set of fuel orifices <NUM> downstream of the first protrusion <NUM> can be defined by a second cross-sectional area, which can be different from the first cross-sectional area. In other words, the fuel orifices of the second set of fuel orifices <NUM> provided on the first protrusion <NUM> can be different from the remaining fuel orifices of the second set of fuel orifices <NUM>. As a non-limiting example, each cross-sectional area can be equal. The sizing of each of the second set of fuel orifices <NUM> can be used to control the mass flow rates of the fuel (F) within the cavity <NUM> and the second fuel channel <NUM>. As a non-limiting example, a smaller cross-sectional area will result in a smaller fuel flow rate. As illustrated, the fuel (F) within the cavity <NUM> would have a larger flow rate than the fuel (F) within the second fuel channel <NUM>.

The first set of fuel orifices <NUM> can be positioned to correspond to one of either the cavity <NUM> or the second fuel channel <NUM>. As a non-limiting example, any number of one more fuel orifices of the first set of fuel orifices <NUM> can be directly fluidly coupled to the cavity <NUM> while the remaining fuel orifices of the first set of fuel orifices <NUM> can be directly fluidly coupled to the second fuel channel <NUM>. The sizing or placement of the second set of fuel orifices <NUM> and the positioning and sizing of the first set of fuel orifices <NUM> can be used to change the fuel flow rates in the second fuel channel <NUM> and the cavity <NUM>, and therefore change the profile of the fuel (F) as the fuel (F) exits the first set of fuel orifices <NUM> and flows into the combustor <NUM>.

<FIG> is a cross-sectional side view of a fuel injector <NUM> for use in a turbine engine in accordance with the claims and which is suitable for use as at least one of the annular array of fuel injectors <NUM> of <FIG>. The fuel injector <NUM> is similar to the fuel injectors <NUM>, <NUM>, therefore, like parts will be identified with like numerals increased to the <NUM> series, with it being understood that the description of the like parts of the fuel injector <NUM>, <NUM> applies to the fuel injector <NUM> unless otherwise noted.

The fuel injector <NUM> can included a wall <NUM> provided along a distal end of the fuel injector <NUM>. The fuel injector <NUM> includes a first set of fuel orifices <NUM> that can be provided along and extend through at least a portion of the wall <NUM> to define an outlet of the fuel injector <NUM>. A flow restrictor <NUM> can be provided within the fuel injector <NUM> upstream of the first set of fuel orifices <NUM> and be defined by a centerline axis <NUM>. The flow restrictor <NUM> can include a second set of fuel orifices <NUM>, with at least a portion of the second set of fuel orifices <NUM> being directly fluidly coupled to a cavity <NUM>.

The fuel injector <NUM> is similar to the fuel injector <NUM>, <NUM>, except that the fuel injector <NUM> is fluidly coupled to a flow of a first fuel (F1) from a first fluid circuit, and a flow of a second fuel (F2) from a second fluid circuit. The first fuel (F1) and the second fuel (F2) can each contain a hydrogen-containing fuel having a respective percentage of hydrogen. The percent hydrogen of the first fuel (F1) can be equal to or non-equal to the percent hydrogen of the second fuel (F2). The difference between the first fuel (F1) and the second fuel (F2) can be the chemical makeup of the fluid flow or a pressure, volume, or velocity of the fluid flow. As a non-limiting example, the first fuel (F1) can contain between <NUM>% and <NUM>% of hydrogen. As a non-limiting example, the second fuel (F2) can contain between <NUM>% and <NUM>% hydrogen. As a non-limiting example, the first fuel (F1) can contain <NUM>% hydrogen and <NUM>% other fuel, while the second fuel (F2) can contain <NUM>% hydrogen. As a non-limiting example, both the first fuel (F1) and the second fuel (F2) can contain <NUM>% hydrogen. Alternatively, the first fuel (F1) and the second fuel (F2) can be identical fuels, with differing or equal mass flow rate, pressure, volume, or velocity. It is contemplated that one of the first fuel (F1) or the second fuel (F2) can contain a non-hydrogen-containing fuel. It will be appreciated that the fuel injector <NUM> can be fluidly coupled to any number of one or more fluid circuits containing any sustainable fuel.

The fuel injector <NUM> can be split into a first fuel channel <NUM> and a second fuel channel <NUM>. The first fuel channel <NUM> can be fluidly coupled to the first fuel (F1), while the second fuel channel <NUM> can be fluidly coupled to the second fuel (F2). It will be appreciated that the fuel injector <NUM> can be split into any number of one or more fuel channels. As a non-limiting example, the number of fuel channels can correspond to the number of fuel circuits that the fuel injector <NUM> is fluidly coupled to.

The flow restrictor <NUM> can be similar to the flow restrictor <NUM> of <FIG> in that the flow restrictor <NUM> includes a first portion, a second portion defining a first protrusion <NUM> and a third portion defining a second protrusion <NUM>, which together define at least a portion of the second fuel channel <NUM>. As illustrated, the first protrusion <NUM> can extend axially through the fuel injector <NUM>. It is contemplated that the first protrusion <NUM> can further define a hose or conduit that is directly fluidly coupled to a portion of the second fuel circuit containing the second fuel (F2). The second set of fuel orifices <NUM>, like the second set of fuel orifices <NUM> (<FIG>), can be axially, radially, and circumferentially spaced about the flow restrictor <NUM> with respect to the centerline axis <NUM>. An axially forward or upstream fuel orifice(s) of the second set of fuel orifices <NUM> can be provided within a portion of the second fuel channel <NUM>, while the remaining fuel orifice(s) can be provided within the first fuel channel <NUM>. The portion of the first fuel channel <NUM> downstream of the second set of fuel orifices <NUM> within the first fuel channel <NUM> can define a cavity <NUM>. Similar to the second set of fuel orifices <NUM> (<FIG>), the second set of fuel orifices <NUM> can be varied in shape, size, or cross-sectional area to affect the profile of the fluid flow that flows through the respective fuel orifices <NUM>. Similar to the second set of fuel orifices <NUM> (<FIG>), the second set of fuel orifices <NUM> can be varied in placement with respect to the cavity <NUM> and the second fuel channel <NUM>. Similar to the flow restrictor <NUM> (<FIG>), the flow restrictor <NUM> include a first segment <NUM> and a second segment <NUM>, with the first segment <NUM> circumscribing the second segment <NUM>. The first segment <NUM> can extend between a first subset <NUM> of the first set of fuel orifices <NUM> and a first subset <NUM> of the second set of fuel orifices <NUM>. The second segment <NUM> can extend between a second subset <NUM>, different from the first subset <NUM>, of the first set of fuel orifices <NUM> and a second subset <NUM>, different from the first subset <NUM>, of the second set of fuel orifices <NUM>.

Further, the variation of the pressure, volume, or velocity of the first fuel (F1) with respect to the second fuel (F2), the placement of the first set of fuel orifices <NUM>, or the second set of fuel orifices <NUM> can be used to further control the profile of the fluid flow (e.g., combined fluid flow of the first fuel (F1) and the second fuel (F2)) that flows out of the fuel injector <NUM> and into the combustor <NUM> (<FIG>).

Benefits associated with the disclosure as described herein include an improvement to the fuel profile, flame shape, combustor exit temperature profile, pollutant emissions, and combustion dynamics within the combustor when compared to a conventional combustor. For example, conventional combustors can include a fuel injector with a fuel channel that feeds to a set of orifices that define an outlet of the fuel injector. In the conventional fuel injector, there is nothing between the inlet of the fuel and the set of orifices. This, in turn, creates a large pressure differential between a portion of the fuel injector upstream of the set of orifices and downstream the set of orifices. This, in turn, results in a non-controlled profile of the fuel exiting the fuel injector and ultimately a non-controlled flame shape. This results in uncontrolled and undesirable combustion dynamics, which can ultimately reduce the overall efficiency or lifespan of the turbine engine including a conventional combustor. The fuel injector as described herein, however, includes the flow restrictor with the set of fuel orifices, and the creation of the cavity or the concentric segments, which can all be used to adjust or otherwise control the profile of the fuel exiting the first set of fuel orifices, and control the combustion dynamics within the combustion chamber. The variation of the profile, as discussed herein, can ultimately control the flame shape, which can ultimately control the combustion dynamics within the combustor. Further, the axial length or volume of the cavity or the segments can be sized with respect to the acoustic oscillation of the flow of the fuel within the fuel injector, such that the acoustic oscillation of the fuel within the cavity or the segments and downstream of the first set of fuel orifices can offset or counteract the acoustic pressure oscillation of the air flow oscillations in the combustor, and ultimately reduce the combustion dynamics in the combustion chamber. The control, mitigation, or counteraction of the combustion dynamics can ultimately result in a turbine engine with a higher efficiency when compared to a conventional turbine engine including a conventional combustor. Further, the controlled flame shape can also impact combustion pollutant emissions and the combustor exit temperature profile, which can result in an eco-friendlier and more efficient turbine engine, respectively, when compared to a conventional turbine engine.

Further benefits of the present disclosure include a combustor with a fuel injector containing a flow of fuel with a hydrogen-containing fuel. Hydrogen-containing fuels have a higher flame temperature than traditional fuels (e.g., fuels not containing hydrogen). That is, hydrogen or a hydrogen mixed fuel typically has a wider flammable range and a faster burning velocity than traditional fuels such petroleum-based fuels, or petroleum and synthetic fuel blends. Further, the hydrogen within the hydrogen-containing fuel is a compressible gas. As such, the fuel can oscillate and interact with the combustion dynamics of the combustor. This, in turn, can increase the overall combustion dynamics of the combustor. Therefore, the many of the combustion components designed for traditional fuels would not be suitable for hydrogen or hydrogen mixed fuels. The fuel injector, as described herein, however, can be used in instances where hydrogen-containing fuels are used. The fuel injector includes the flow restrictor, the cavity, and the first set of fuel orifices. When the fuel flows against the flow restrictor and through the second set of fuel orifices, the fuel has a first pressure upstream of the flow restrictor and a second pressure downstream of the flow restrictor (e.g., within the cavity), thus defining a first pressure drop. The fuel within the cavity and at the second pressure can then flow through the first set of fuel orifices, where the fuel is then at a third pressure downstream of the first set of fuel orifices, thus defining a second pressure drop. The first pressure drop is larger than the second pressure drop. The two stages or pressure drops, in turn, results in a relatively small pressure drop is experienced across the second set of fuel orifices when compared to a conventional fuel injector that does not include the flow restrictor (e.g., only a single set of orifices defining the outlet of the fuel injector). The pressure drop across the flow restrictor upstream of the first set of fuel orifices, the pressure drop across the first set of fuel orifices, and the inclusion of the segments or cavity can be used to create fuel flow oscillations that will cancel or counteract the air flow oscillations in the combustor, resulting in reduced overall combustion dynamics of the combustor, which in turn can increase the lifespan and efficiency of the turbine engine when compared to a conventional turbine engine using traditional fuels.

To the extent not already described, the different features and structures of the various aspects can be used in combination, or in substitution with each other as desired. That one feature is not illustrated in all of the examples is not meant to be construed that it cannot be so illustrated, but is done for brevity of description. Thus, the various features of the different aspects can be mixed and matched as desired to form new aspects, whether or not the new aspects are expressly described. All combinations or permutations of features described herein are covered by this disclosure.

Claim 1:
A turbine engine (<NUM>) comprising:
a compressor section (<NUM>), a combustion section (<NUM>), and a turbine section (<NUM>) in serial flow arrangement; and
a combustor (<NUM>), provided within the combustion section (<NUM>), the combustor comprising:
a dome wall and a combustor liner collectively forming a combustion chamber (<NUM>) and a plurality of fuel injectors (<NUM>, <NUM>) extending through a respective portion of the dome wall, the fuel injectors (<NUM>, <NUM>) provided in an annular arrangement, the fuel injectors comprising:
a fuel channel (<NUM>, <NUM>, <NUM>, <NUM>), wherein the fuel channel defines a centerline axis, terminates at a distal end, and is fluidly coupled to a fuel (F, F1, F2);
a wall (<NUM>, <NUM>) provided within the fuel channel, the wall including a first set of fuel orifices (<NUM>, <NUM>) extending therethrough and fluidly coupling the fuel channel (<NUM>, <NUM>, <NUM>, <NUM>) to the combustion chamber (<NUM>); and
a flow restrictor (<NUM>, <NUM>) located within the fuel channel (<NUM>, <NUM>, <NUM>, <NUM>), upstream of and spaced from the first set of fuel orifices (<NUM>, <NUM>), and having a second set of fuel orifices (<NUM>, <NUM>) fluidly coupled to the first set of fuel orifices (<NUM>, <NUM>);
wherein a diameter of the fuel channel (<NUM>, <NUM>, <NUM>, <NUM>) at the wall (<NUM>, <NUM>) is the same as the diameter of the fuel channel at the flow restrictor (<NUM>, <NUM>)wherein the flow restrictor (<NUM>, <NUM>) includes a first portion and a second portion, and wherein the second portion extends axially outwardly from a plane normal to the centerline axis (<NUM>, <NUM>) and intersecting the first portion to define a first protrusion (<NUM>, <NUM>).