COMBUSTOR WTH PRE-MIXING FUEL NOZZLE ASSEMBLY

Abstract

A premixing fuel nozzle assembly including a plurality of premixers for use in a combustor section of a turbine engine is provided. Each of the premixers of the pre-mixing fuel nozzle assembly including an annular outer ring, an annular inner hub, configured co-annular with the outer ring, a plurality of swirler vanes, a plurality of fuel injection orifices formed in each vane and an air-assist orifice formed about at least one of the plurality of fuel injection orifices. The plurality of swirler vanes extend radially outward from the annular inner hub toward the annular outer ring. The air-assist orifice generating a co-annular air-assist jacket about a fuel jet injected via the fuel injection orifice to provide additional momentum and thus mixing of the fuel jet with a downstream crossflow of fluid. A combustor section including the pre-mixing fuel nozzle assembly and method of assembling the pre-mixing fuel nozzle are additionally disclosed.

Description

BACKGROUND

The present disclosure relates in general to turbine systems, and more particularly relates to a low NO_x combustor with a pre-mixing fuel nozzle assembly.

Heavy duty gas turbine systems are widely utilized in fields such as power generation. For example, a conventional heavy duty gas turbine system includes a compressor section, a combustor section, and at least one turbine section. The compressor section is configured to compress air as the air flows through the compressor section. The air is then flowed from the compressor section to the combustor section, where it is mixed with fuel and combusted, generating a hot gas flow. The hot gas flow is provided to the turbine section, which utilizes the hot gas flow by extracting energy from it to power the compressor, an electrical generator, and/or other various loads.

The fuel supplied to the combustor may be a liquid fuel, a gaseous fuel, or a combination of liquid and gaseous fuels, depending on various factors such as the operating mode, operating level, and availability of various fuels. If the liquid fuel, gaseous fuel, and/or other fluids are not evenly mixed with the compressed working fluid prior to combustion, localized hot spots may form in the combustor, particularly near the nozzle exits. The localized hot spots may increase the production of pollutant emissions. These pollutant emissions generally include carbon oxides (CO_x), nitrogen oxides (NO_x), sulfur oxides (SO_x), and particulate matter (PM). In particular, the hotspots may increase production of nitrous oxides (NO_x) in the fuel rich regions, while the fuel lean regions may increase the production of carbon monoxide and unburned hydrocarbons, all of which are undesirable exhaust emissions

The primary challenge in developing heavy duty gas turbine systems is managing these pollutant emissions, and in particular, NO_x emissions. These pollutant emissions are highly regulated in the United States and elsewhere. Accordingly, low NO_x combustors for gas turbines are known in the industry. The lowest NO_x emissions that can be achieved with a given combustor is directly related to its ability to increase fuel and air mixing. At least some known turbine assemblies facilitate reducing NO_x emissions by using pre-mixing technology. Pre-mixing fuel and air facilitates inhibiting the temperature of combustion gases such that the combustion temperature does not rise above the threshold where NO_x emissions are formed.

One type of known pre-mixing nozzle is a swirling annular fuel nozzle or “swozzle” which typically includes a number of vanes extending between the inner hub and an outer shroud. The vanes are circumferentially spaced apart and include fuel injection openings. The fuel injection openings are supplied fuel by an internal circuit and receive fuel through fuel passages that extend radially outward from the fuel entry openings in the inner hub. In operation, air traveling axially through the swozzle is swirled by the vanes and fuel traveling radially through the swozzle is injected into the swirling air flow. The fuel enters the combustor in a jet, in crossflow arrangement, and mixes with the turbulent crossflow to achieve a uniform fuel/air mixture before entering the combustor. The ability to achieve this uniform mixture is dependent on the fuel jet penetration. The jet fuel penetration for a given orifice diameter is directly related to the fuel density and velocity. The fuel density can vary continuously throughout gas turbine operation with change in fuel composition. The fuel orifice exit velocity can vary depending on the operating condition. An ability to achieve good jet penetration characteristics is necessary throughout the various operation scenarios.

Accordingly, an improved technique is needed to reduce pollutant emissions, such as NO_x emissions, from a gas turbine combustor. More particularly, an improved system for fuel injection is needed to provide increased jet penetration and thereby mixing improvement and dynamics mitigation in the premixer of a gas turbine combustor.

BRIEF DESCRIPTION

In accordance with one exemplary embodiment a premixer of a fuel nozzle assembly is disclosed. The premixer of the fuel nozzle assembly including an annular outer ring, an annular inner hub, a plurality of swirler vanes, a plurality of fuel injection orifices and an air-assist orifice. The annular inner hub is configured co-annular with the outer ring. The plurality of swirler vanes extend radially outward from the annular inner hub toward the annular outer ring. The plurality of fuel injection orifices are formed in each van, with the air-assist orifice is formed about at least one of the plurality of fuel injection orifices.

In accordance with another embodiment, a premixer of a fuel nozzle assembly is disclosed. The premixer including an annular outer ring, an annular inner hub, configured co-annular with the outer ring, a plurality of swirler vanes extending radially outward from the annular inner hub toward the annular outer ring. Each of the plurality of swirler vanes includes a fuel passage portion and an airfoil portion. The premixer further including a plurality of fuel injection orifices formed in each vane and through the airfoil portion, an air-assist orifice formed co-annular with and about at least one of the plurality of fuel injection orifices and through the airfoil portion, one or more fuel passages defined within each swirler vane and extending through the fuel passage portion and the airfoil portion and one or more air passages extending through the fuel passage portion and the airfoil portion. The one or more fuel passages terminating at one or more of the plurality of fuel injection orifices and the one or more air passages terminating at one or more of the air-assist orifices.

In accordance with another embodiment, a combustor section of a gas turbine engine is disclosed. The combustor section including a combustion chamber and a fuel nozzle assembly associated with the combustion chamber. The fuel nozzle assembly including a plurality of swirler vanes configured to swirl airflow. Each of the plurality of swirler vanes including a plurality of fuel injection orifices configured to inject a plurality of fuel jets into the airflow and an air-assist orifice configured about at least one of the plurality of fuel injection orifices and form an air-assist jacket about one or more of the fuel jets.

In accordance with yet another embodiment a method of assembly a fuel nozzle assembly is disclosed. The method including providing an outer ring, positioning an inner hub coaxially within the outer ring such that a plenum is formed therebetween, coupling a swirler between the outer ring and the inner hub, defining a plurality of fuel injection orifices in the airfoil portion of one or more of the plurality of swirler vanes and defining an air-assist orifice about at least one of the plurality of fuel injection orifices to form an air-assist jacket about one or more of the fuel jets. The swirler including a plurality of swirler vanes is configured to rotate fluid flowing therethrough in a downstream direction, each of the plurality of swirler vanes comprising a fuel passage portion and an airfoil portion. The plurality of fuel injection orifices are configured to inject a plurality of fuel jets into the plenum and the air-assist orifice is configured to provide increased momentum to the fuel jet during injection into the plenum.

DRAWINGS

These and other features and aspects of embodiments of the present invention will become better understood when the following detailed description is read with reference to the accompanying drawings in which like characters represent like parts throughout the drawings, wherein:

FIG. 1 is schematic view of an exemplary turbine engine, according to one or more embodiments shown or described herein;

FIG. 2 is a perspective cross-sectional schematic view of an exemplary pre-mixing fuel nozzle assembly that may be used with the turbine engine shown in FIG. 1, according to one or more embodiments shown or described herein;

FIG. 3 is a schematic cross-sectional view of a portion of the exemplary pre-mixing fuel nozzle assembly that may be used with the turbine engine shown in FIG. 1, according to one or more embodiments shown or described herein;

FIG. 4 is a perspective view illustrating an embodiment of a swozzle that may be used within the pre-mixing fuel nozzle assembly shown in FIGS. 2 and 3, according to one or more embodiments shown or described herein;

FIG. 5 is a cross-sectional view of an embodiment of a vane of the swozzle shown in FIG. 4, according to one or more embodiments shown or described herein;

FIG. 6 is a schematic illustration of a portion of the swozzle of FIG. 3, according to one or more embodiments shown or described herein;

FIG. 7 is a cross-section view of modeling performed on a three-dimensional model illustrating fuel jet penetration, according to one or more embodiments shown or described herein;

FIG. 8 is a graphical plot illustrating fuel jet penetration, according to one or more embodiments shown or described herein; and

FIG. 9 is a graphical plot illustrating fuel/air mixing effectiveness, according to one or more embodiments shown or described herein.

DETAILED DESCRIPTION

The present disclosure is directed to systems for improving the injection of fuel (e.g., liquid and/or gas) into a fuel nozzle, thereby enhancing premixing of the fuel (e.g., premixing fuel and air), fuel Wobbe capability (i.e., interchangeability of fuels used) and control over the fuel-air profile. In particular, embodiments of the present disclosure include a pre-mixing fuel nozzle assembly for a combustor of a turbo engine that utilizes an air-assist orifice configured co-annular with a fuel orifice to form an air jacket about the fuel orifice. By choosing an optimal flow and dimension of the air-assist orifice and thus the resultant co-annular air jacket, fuel jet penetration is improved, resulting in enhanced mixing of the inlet fuel and air while providing control over the fuel-air profile, and thus reducing emissions. By varying the flow rate through the air-assist orifice the fuel impedance characteristics can be tuned to reduce the combustion instability within the combustor.

One or more specific embodiments of the present disclosure will be described below. In an effort to provide a concise description of these embodiments, all features of an actual implementation may not be described in the specification. It should be appreciated that in the development of any such actual implementation, as in any engineering or design project, numerous implementation-specific decisions must be made to achieve the developers' specific goals, such as compliance with system-related and business-related constraints, which may vary from one implementation to another. Moreover, it should be appreciated that such a development effort might be complex and time consuming, but would nevertheless be a routine undertaking of design, fabrication, and manufacture for those of ordinary skill having the benefit of this disclosure.

When introducing elements of various embodiments of the present invention, the articles “a,” “an,” “the,” and “said” are intended to mean that there are one or more of the elements. The terms “comprising,” “including,” and “having” are intended to be inclusive and mean that there may be additional elements other than the listed elements.

FIG. 1 is a schematic view of an exemplary turbine engine 100. More specifically, in the exemplary embodiment turbine engine 100 is a gas turbine engine. While the exemplary embodiment illustrates a gas turbine engine, the present invention is not limited to any one particular engine, and one of ordinary skill in the art will appreciate that the pre-mixing fuel nozzle assembly, as described herein may be used in connection with other turbine engines.

In the exemplary embodiment, turbine engine 100 includes an intake section 102, a compressor section 104 downstream from the intake section 102, a combustor section 106 downstream from the compressor section 104, a turbine section 108 downstream from the combustor section 106, and an exhaust section 110. The turbine section 108 is coupled to the compressor section 104 via a rotor shaft 112. In the exemplary embodiment, the combustor section 106 includes a plurality of combustors 114. The combustor section 106 is coupled to the compressor section 104 such that each of the plurality of combustor 114 is in flow communication with the compressor section 104. A combustor can 116, including a pre-mixing fuel nozzle assembly, disclosed herein, is coupled within each of the plurality of combustor 114. The turbine section 108 is coupled to the compressor section 104 and to a load 118 such as, but not limited to, an electrical generator and/or a mechanical drive application through the rotor shaft 112. In the exemplary embodiment, each of the compressor section 104 and the turbine section 108 includes at least one rotor disk assembly 120 coupled to the rotor shaft 112 to form a rotor assembly 122.

During operation, the intake section 102 channels air towards the compressor section 104 wherein the air is compressed to a higher pressure and temperature prior to being discharged towards the combustor section 106. The compressed air is mixed with fuel and other fluids and then ignited to generate combustion gases that are channeled towards the turbine section 108. More specifically, the fuel mixture is ignited to generate high temperature combustion gases that are channeled towards the turbine section 108. The turbine section 108 converts the energy from the gas stream to mechanical rotational energy, as the combustion gases impart rotational energy to the turbine section 108 and to the rotor assembly 122.

FIGS. 2 and 3 are a schematic perspective cross-sectional schematic view and a schematic cross-sectional schematic view of a portion of one of the combustor cans 116 of FIG. 1. In the exemplary embodiment, the combustor can 116 includes a substantially cylindrical shroud 130 having a first end portion 132, a second end portion 134, and a mixing zone 136 defined there between. On an upstream end, the combustor can 116 includes an end cover 138. The end cover 138 may have included therein a fuel manifold 140. In addition, the end cover 138 supports a number of fuel nozzle assemblies 142 (of which only one is shown in FIG. 3) in fluid communication with the fuel manifold 140, that can communicate fuel into the combustor section 106 (FIG. 1). In an embodiment, the combustor can 116 includes a pre-mixing fuel nozzle assembly 142 for the fuel system (i.e., the engine 100 of FIGS. 1 and 2), configured to include a downstream swirling annular fuel nozzle assembly, referred to herein as a premixer or “swozzle” assembly 144. Each of the premixer assemblies 144 includes an internal fuel circuit (described presently). It is anticipated that in an embodiment, each of the premixer assemblies 144, and more particularly, the internal fuel circuits, are not identical, thereby allowing a flow of fuel to each of the premixer assemblies 144 to be different. The swozzle or premixer assembly 144 includes a fuel-air mixer, or swirler, 146, one or more downstream orifices 147, one or more upstream orifices 150 and connecting passages 152 defined within an extended centerbody 167. The one or more downstream orifices 147 are configured according to this disclosure. A system and method including one or more downstream orifices formed to define a first fluid path and an upstream orifice located upstream and formed to define a second fluid path in a turbomachine is described in detail in U.S. Pat. No. 8,322,140, entitled, “Fuel System Acoustic Feature to Mitigate Combustion Dynamics for Multi-Nozzle Dry Low NOx Combustion System and Method”, issued on Dec. 4, 2012 by Kim et al., assigned to the same assignee and incorporated herein by reference.

Each of the downstream orifices 147 is formed to define a first fluid path along which fluid, such as fuel flow, is directed to flow toward the mixing zone 136 and then the combustor section 106 (FIG. 1) to which the pre-mixing fuel nozzle assembly 142 is fluidly coupled. In an embodiment, the upstream orifice 150 is located upstream from the downstream orifices 147 relative to a flow direction of the connecting passages 152 and is formed to define a second fluid path along which the fluid is directed to flow toward the downstream orifices 147. The connecting passages 152 are disposed to fluidly couple the first and second fluid paths of the downstream and upstream orifices 147, 150, respectively, to one another. With this configuration, at least one of a radial size and an axial distance (D_A) of the upstream orifice 150, relative to the downstream orifices 147, may be set at variable values to cooperatively detune an acoustic impedance of the fuel system to thereby prevent or substantially mitigate fuel/air ratio fluctuation-driven combustion dynamics.

Each of the downstream orifices 147 is located at an axial location and has a radial size. The upstream orifice 150 is located at an axial location and has a radial size, which may be set at variable values and opens up to the connecting passages 152.

In accordance with embodiments, the pre-mixing fuel nozzle assembly 142 may be plural in number. That is, as shown in FIG. 1, the pre-mixing fuel nozzle assembly 142 may be provided as any number of separate fuel nozzle assemblies 142. In any case, each of the fuel nozzle assemblies 142 has a substantially similar axial length as measured from the common manifold 140, and is fluidly coupled to the common manifold 140 to thereby receive a common fuel flow.

In the exemplary embodiment, fuel 154 is channeled into the fuel manifold 140 via a fuel inlet passage 156. The fuel 154 is channeled into the swozzle or premixer assembly 144 to combine with air assist 158 that is also channeled through the swozzle or premixer assembly 144 (described presently). In an embodiment, air may be obtained from compressor air through the head end. A fuel-air mixture 160 then exits the combustor can 116 for use in the combustor section 106 (shown in FIG. 1).

Referring now to FIG. 4, illustrated is a portion of a combustor can 116, and more particularly the swozzle or premixer assembly 144. In the exemplary embodiment, the swozzle or premixer assembly 144 includes an outer ring, or shroud, 164 and an inner hub 166 disposed about the extended centerbody 167 (FIG. 3). The inner hub 166 is coaxially disposed within the outer ring 164 about a central axis 168. As such, in the exemplary embodiment, a plenum 170 is defined between the outer ring 164 and the inner hub 166. As used herein, the term “axial”, “axially”, or “coaxially” refers to a direction along or substantially parallel to the central axis 168. Furthermore, as used herein, the term “radial” or “radially” refers to a direction substantially perpendicular to the central axis 168.

Furthermore, in the exemplary embodiment, the plenum 170 includes a swirler 172 defined therein. The swirler 172 includes a plurality of swirler vanes 174 such as a first swirler vane 176, a second swirler vane 178, a third swirler vane 180, a fourth swirler vane 182, a fifth swirler vane 184, a sixth swirler vane 186, a seventh swirler vane 188, an eighth swirler vane 190, a ninth swirler vane 192, and a tenth swirler vane 194. Although the swirler 172 is shown as including ten swirler vanes, it should be understood that swirler 172 may include any suitable number of swirler vanes such that the swozzle or premixer assembly 144 functions as described herein. Furthermore, in the exemplary embodiment, the swirler vanes 176, 178, 180, 182, 184, 186, 188, 190, 192 and 194 are spaced circumferentially about the inner hub 166 such that a plurality of inner passages 195 are defined therebetween. As an example, in the exemplary embodiment, a first inner passage 196 is defined between the swirler vanes 176 and 178, with additional inner passages 195 defined between each adjacent set of swirler vanes 174. The swirler vanes 174 are coupled to the inner hub 166 and the outer ring 164 such that the swirler vanes 174 extend from the inner hub 166 to the outer ring 164. In the exemplary embodiment, the swozzle or premixer assembly 144 includes a leading edge 200 and a trailing edge 202. The swirler 172 is configured to rotate fluid flowing therethrough.

In the exemplary embodiment, the swirler 172 further includes a plurality of fuel injection conduits defined therein for delivering fuel to the plenum 170 and a plurality of air injection conduits defined therein for delivering air assist to the fuel entering the plenum 170. More specifically, each of the plurality of swirler vanes 174 includes a plurality of the fuel injection orifices 148 and one or more air-assist orifices 206, each formed about one or more of the fuel injection orifices 148. Each vane 174 includes a fuel passage portion and an airfoil portion (described presently). The fuel passage portion is positioned adjacent to inner hub 166, while the airfoil portion extends away from the fuel passage portion in an axial direction. The fuel injection orifices 148 and air-assist orifices 206 are formed through the airfoil portion, such as through one or both of a pressure side and a suction sides of the airfoil portion.

The fuel injection orifices 148 and air-assist orifices 206 are supplied fuel 154 and air 158, respectively, by internal circuits and receive the fuel 154 and the air 158 through fuel and air passages that extend radially outward toward the fuel injection orifices 148 and air-assist orifices 206 in the airfoil portion or each of the plurality of vanes 174. FIG. 5 is cross-sectional view of an embodiment of a portion of one such vane 174. Each fuel injection orifice 148 (of which only one is illustrated) is in fluid communication with a fuel passage 212 in the vane 174, which in turn is in fluid communication with a fuel manifold 140 (FIG. 3) and a source of fuel. Each air-assist orifice 206 (of which only one is illustrated surrounding the one fuel injection orifice 148) is in fluid communication with an air passage 214 in the vane 174, which in turn is in fluid communication with a source of air, such as compressor air or any other source of high pressure air. In an embodiment, the fuel passage 212 extends from a fuel entry opening 216 through a fluid passage portion 218 of the vane 174 to an airfoil portion 220 of the vane 174, terminating at the fuel injection orifices 148. In an embodiment the fuel passage 212 is in fluid communication with a fuel plenum 221. Furthermore, the air passage 214 extends from an air entry opening 222 through the fluid passage portion 218 of the vane 174 to the airfoil portion 220 of the vane 174, terminating at the air-assist orifices 206. In an embodiment, the air passage 214 is in fluid communication with an air plenum 223. In some embodiments, the swirler vanes 174 may additionally be configured to be cooled by the flow of fuel 154 or air 158 through the fuel passage 212 and/or air passage 214. More particularly, fuel passage 212 and/or air passage 214 may be positioned to direct fuel 154 and/or air 158 over a portion of an inner surface of the airfoil portion 220 for cooling purposes before the fuel 154 and/or air 158 exit the airfoil portion 220 through the fuel injection orifices 148 and/or the air-assist orifices 206. For example, the air passage 214 may extend axially through the vane 174 to a surface of the airfoil portion 220, where the air passage 214 may branch and travel along an inner surface of the airfoil portion 220 toward the air-assist orifices 206, which are positioned on opposite sides of the airfoil portion 220. However, other configurations are possible.

Known swozzle assemblies, or air-fuel mixers, are typically unable to deliver a uniform mixture of fuel 154 and crossflow of fluid 228 into the combustor section 106 (FIG. 1) due to inadequate fuel penetration. It has been found that the jet fuel penetration for a given fuel injection orifice diameter is directly related to the fuel velocity and density, which can vary continuously throughout the gas turbine operation depending on the operating condition. As best illustrated in FIG. 6, in an embodiment disclosed herein, a diameter “D” of each of the fuel injection orifices 148, referenced D₁, D₂, D₃, increases radially, where D₁<D₂<D₃, to accommodate an increasing cross-sectional area, with a resultant pressure drop, radially from D₁ to D₃ across the fuel injection orifices 148. As best illustrated in FIG. 7, to achieve good fuel jet penetration characteristics throughout various operation scenarios, a co-annular air jacket 224, formed via the one or more air-assist orifices 206, about a fuel jet 226, formed via the one or more fuel injection orifices 148, can supply the additional momentum to the fuel jet 226 thus providing and maintaining good penetration throughout various operating modes. As indicated in FIG. 7, an annular area of the air-assist orifice 206 is optimized to provide a minimum jet momentum ratio “X”, which will add to the momentum ratio of the fuel jet “Y”. The combined jet momentum ratio of the co-annular air jacket 224 along with the fuel jet 226 is thus X+Y. This combined momentum provides additional penetration of the fuel jet 226 with a crossflow of fluid 228, thus allowing enhanced mixing downstream. This is particularly advantageous when the fuel jet momentum ratio varies significantly. A minimum fuel jet ratio ensures that even at low fuel flow rates or low fuel jet momentum ratio, the fuel jet 226 penetration is not degraded.

In addition it is a well-known fact that fuel air oscillations generated at an exit of the fuel injection orifice 148 propagates downstream into a flame in the combustor section 106 (FIG. 1) and cause unwanted flame oscillations. These unwanted flame oscillations can drive pressure variations in the combustor, thus reducing the durability of the combustor structural components. The inclusion of the co-annular air-assist jacket 224 about one or more of the fuel jets 226 provides a control means for these fuel air oscillations.

In operation, the crossflow of fluid 228, typically air, traveling axially through the swozzle or premixer assembly 144, and more particularly the swirler 172, is swirled by the swirler vanes 174 and the fuel jet 226 and co-annular air-assist jacket 224 formed thereabout, traveling radially through the swirler 172 are injected into the swirling crossflow of fluid 228.

Referring more specifically to FIGS. 8 and 9, illustrated in graphical plots 230 and 250, respectively, are jet trajectory and air-fuel mixing both plotted with and without the air-assist air jacket, as disclosed herein. More specifically, illustrated in FIG. 8 in graphical plot 230, are axial locations within a mixing zone duct, of a fuel jet penetration, plotted on an axis 232 (X-axis) as a function of the radial locations within the mixing zone, of a jet penetration, plotted on an axis 234 (Y-axis) with and without the use of an air-assist jacket, as disclosed herein. More specifically, illustrated is the calculated penetration distance, both axially (in an x-direction) and radially (in a z-direction), of a fuel jet from an injection point 236 toward an exit of the mixing zone 238. Line 240 shows the fuel jet penetration values and fuel jet trajectory in the presence of an air-assist jacket, such as air-assist jacket 224 (FIG. 4) disclosed herein. Line 242 shows the fuel jet penetration values fuel jet trajectory without the presence of an air-assist jacket, such as air-assist jacket 224 (FIG. 4) disclosed herein. As seen clearly from plot 230, improved fuel jet penetration is achieved in the presence of the air-assist jacket 224 (FIG. 4). In addition, the use of the air-assist jacket results in improved fuel spread across the duct and thereby better mixing of the fuel and air.

Illustrated in FIG. 9 in graphical plot 250, are axial locations within a mixing zone duct, of a fuel jet penetration, plotted on an axis 252 (X-axis) as a function of the mixing effectiveness of the fuel and air, plotted on an axis 254 (Y-axis) with and without the use of an air-assist jacket, as disclosed herein. More specifically, illustrated is the calculated mixing effectiveness of a fuel jet axially, from an injection point 256 toward a duct exit 258. Line 260 shows the mixing effectiveness values in the presence of an air-assist jacket, such as air-assist jacket 224 (FIG. 4) disclosed herein. Line 262 shows the mixing effectiveness values without the presence of an air-assist jacket, such as air-assist jacket 224 (FIG. 4) disclosed herein. As seen clearly from plot 250, an improved degree of mixing of the fuel and air, noted at 264, is achieved in the presence of the air-jacket assist 224 (FIG. 4), as disclosed herein.

Referring again to FIGS. 2-6, a range of other configurations are also envisioned within the scope of the present disclosure. For example, any number of vanes 174 may be positioned within the swozzle assembly 116, the end cover 138 may have any number of fuel manifolds 140, and the fuel manifolds 140 may communicate fuel into any number or combination of the swirler vanes 174.

In operation, air 158 from the compressor section 104 is driven by a pressure differential along the connecting passages 152 toward the swozzle or premixer assembly 144. At least a portion of the air 158 is directed into the swirler 172, and more particularly through the swirler vanes 174 and air-assist orifices 206 in a generally radially direction, to form the co-annular air jackets 224. In addition, the fuel 154 is directed into the swirler 172, and more particularly through the swirler vanes 174 and the fuel injection orifices 148, having the co-annular air jackets 224 formed thereabout, in a generally radially direction. In addition, a portion of the crossflow fluid 228 is directed between the swirler vanes 174, which swirls the crossflow fluid 228. The fuel 154, air 158 and crossflow fluid 228 create the air-fuel mixture 160. The air-fuel mixture 160 travels axially through the assembly 142 into the combustors 114. Such a configuration differs from known fuel-nozzle assemblies by the inclusion of the air-assist orifices 206 to form the co-annular air-assist jackets 224 about the fuel jets 226 and provide enhanced mixing and penetration of the fuel jets 226 and the crossflow fluid 228.

It is anticipated that some or all of the swirler vanes 174 may be fueled via one or more fuel jet orifices 148 and/or include air-assist orifices 206 depending on the operating mode. For example, fuel jets 226 and/or air-assist fuel jackets 224 may be provided to a distinct sub-set of the swirler vanes 174, or fuel jets 226 and/or air-assist jackets 224 may be provided to all of the swirler vanes 174. In the illustrated embodiment, all ten vanes 174 are fueled, with at least a portion of the fuel jet orifices 148 including air-assist orifices 206 and thus air-assist jackets 224.

The co-annular air-assist jackets 224 can supply additional momentum to the fuel jets 226 thus maintaining good penetration throughout the operating modes. In addition, it is a well-known fact that fuel air oscillations generated at the exits of the fuel injection orifices 148 propagate downstream into a combustor flame and cause unwanted flame oscillations which can drive pressure variations in the combustor, thus reducing the durability of the combustor structural components. The inclusion of the air-assist jackets 224 can provide a means by which this fuel air oscillations originating at the exits of the fuel injection orifices 148 can be controlled. It is a known fact that current combustors aim to achieve optimal fuel splits across multiple premixers to maintain dynamics as well as reduce emissions. The air-assist jackets 224 may therefore additionally play a role of modulating the fuel air oscillations by employing them as an additional means for control.

Accordingly, disclosed herein is a pre-mixing fuel nozzle assembly and method of assembly the fuel nozzle, for use in a combustor section of a turbine engine including an annular outer ring, an annular inner hub, configured co-annular with the outer ring, a plurality of swirler vanes extending radially outward from the annular inner hub toward the annular outer ring, a plurality of fuel injection orifices formed in each vane and an air-assist orifice formed about at least one of the plurality of fuel injection orifices. The air-assist orifice generating a co-annular air-assist jacket about a fuel jet injected via the fuel injection orifice to provide additional momentum and thus mixing of the fuel jet with a downstream crossflow of fluid.

It is to be understood that not necessarily all such objects or advantages described above may be achieved in accordance with any particular embodiment. Thus, for example, those skilled in the art will recognize that the systems and techniques described herein may be embodied or carried out in a manner that achieves or improves one advantage or group of advantages as taught herein without necessarily achieving other objects or advantages as may be taught or suggested herein.

While the technology has been described in detail in connection with only a limited number of embodiments, it should be readily understood that the specification is not limited to such disclosed embodiments. Rather, the technology can be modified to incorporate any number of variations, alterations, substitutions or equivalent arrangements not heretofore described, but which are commensurate with the spirit and scope of the claims. Additionally, while various embodiments of the technology have been described, it is to be understood that aspects of the specification may include only some of the described embodiments. Accordingly, the specification is not to be seen as limited by the foregoing description, but is only limited by the scope of the appended claims.

Claims

1. A premixer of a fuel nozzle assembly, comprising: an annular outer ring;an annular inner hub, configured co-annular with the outer ring;a plurality of swirler vanes extending radially outward from the annular inner hub toward the annular outer ring;a plurality of fuel injection orifices formed in each vane; andan air-assist orifice formed about at least one of the plurality of fuel injection orifices.

2. The premixer as claimed in claim 1, wherein each of the plurality of swirler vanes includes: a fuel passage portion;an airfoil portion;one or more fuel passages extending through the fuel passage portion and the airfoil portion, the one or more fuel passages terminating at one or more of the plurality of fuel injection orifices formed through the airfoil portion; andone or more air passage extending through the fuel passage portion and the airfoil portion, the one or more air passages terminating at one or more of the air-assist orifices formed through the airfoil portion and about the fuel injection orifice.

3. The premixer as claimed in claim 2, wherein each of the plurality of fuel injection orifices are in fluid communication with the one or more fuel passages to generate a plurality of fuel jets and wherein the air-assist orifice is in fluid communication with the one or more air passages to form a co-annular air-assist jacket about at least one of the plurality of fuel jets.

4. The premixer as claimed in claim 3, wherein the plurality of fuel jets and the co-annular air-assist jacket are mixed with an incoming cross-flow of air to generate a fuel-air mixture.

5. The premixer as claimed in claim 1, wherein the plurality of swirler vanes are circumferentially spaced apart from each other about the inner hub, and define a plenum.

6. The premixer as claimed in claim 1, wherein the plurality of fuel injection orifices are configured having varying areas to address pressure drop and jet penetration across each of the plurality of fuel injection orifices.

7. The premixer as claimed in claim 1, wherein an area of the air-assist orifice formed about at least one of the plurality of fuel injection orifices and flow of air therethrough is optimized to provide improved jet penetration.

8. A premixer of a fuel nozzle assembly comprising: an annular outer ring;an annular inner hub, configured co-annular with the outer ring;a plurality of swirler vanes extending radially outward from the annular inner hub toward the annular outer ring, each of the plurality of swirler vanes including a fuel passage portion and an airfoil portion;a plurality of fuel injection orifices formed in each vane and through the airfoil portion;an air-assist orifice formed co-annular with and about at least one of the plurality of fuel injection orifices and through the airfoil portion;one or more fuel passages defined within each swirler vane and extending through the fuel passage portion and the airfoil portion, the one or more fuel passages terminating at one or more of the plurality of fuel injection orifices; andone or more air passages extending through the fuel passage portion and the airfoil portion, the one or more air passages terminating at one or more of the air-assist orifices.

9. The premixer as claimed in claim 8, wherein each of the plurality of fuel injection orifices are in fluid communication with the one or more fuel passages to generate a plurality of fuel jets and wherein the air-assist orifice is in fluid communication with the one or more air passages to form a co-annular air-assist jacket about at least one of the plurality of fuel jets.

10. The premixer as claimed in claim 8, wherein the plurality of fuel jets and the co-annular air-assist jacket are mixed with an incoming cross-flow of air to generate a fuel-air mixture.

11. A combustor section of a gas turbine engine, comprising: a combustion chamber;a fuel nozzle assembly associated with the combustion chamber, the fuel nozzle assembly comprising a plurality of swirler vanes configured to swirl airflow, each of the plurality of swirler vanes comprising: a plurality of fuel injection orifices configured to inject a plurality of fuel jets into the airflow; andan air-assist orifice configured about at least one of the plurality of fuel injection orifices and form an air-assist jacket about one or more of the fuel jets.

12. The combustor section as claimed in claim 11, wherein the fuel nozzle assembly further comprises: an annular outer ring;an annular inner hub, configured co-annular with the outer ring; andthe plurality of swirler vanes extending radially outward from the annular inner hub toward the annular outer ring.

13. The combustor section as claimed in claim 12, wherein each of the plurality of swirler vanes includes: a fuel passage portion;an airfoil portion; andone or more fuel passages extending through the fuel passage portion and the airfoil portion, the one or more fuel passages terminating at one or more of the plurality of fuel injection orifices formed through the airfoil portion; andone or more air passage extending through the fuel passage portion and the airfoil portion, the one or more air passages terminating at one or more of the air-assist orifices formed through the airfoil portion and about the fuel injection orifice.

14. The combustor section as claimed in claim 13, wherein each of the plurality of fuel injection orifices are in fluid communication with the one or more fuel passages to generate a plurality of fuel jets and wherein the air-assist orifice is in fluid communication with the one or more air passages to form a co-annular air-assist jacket about at least one of the plurality of fuel jets.

15. The combustor section as claimed in claim 14, wherein the plurality of fuel jets and the co-annular air-assist jacket are mixed with an incoming cross-flow of air to generate a fuel-air mixture.

16. The combustor section as claimed in claim 11, wherein the plurality of swirler vanes are circumferentially spaced apart from each other about the inner hub, and define a plenum.

17. The combustor section as claimed in claim 12, further comprising: an end cover that encloses the combustor;a fuel manifold formed in the end cover; anda plurality of connecting passages in fluid communication with the fuel manifold and the fuel nozzle assembly, the plurality of connecting passages configured to provide fuel to the plurality of swirler vanes.

18. A method of assembly a fuel nozzle assembly, comprising: providing an outer ring;positioning an inner hub coaxially within the outer ring such that a plenum is formed therebetween;coupling a swirler between the outer ring and the inner hub, the swirler including a plurality of swirler vanes configured to rotate fluid flowing therethrough in a downstream direction, each of the plurality of swirler vanes comprising a fuel passage portion and an airfoil portion;defining a plurality of fuel injection orifices in the airfoil portion of one or more of the plurality of swirler vanes, the plurality of fuel injection orifices configured to inject a plurality of fuel jets into the plenum; anddefining an air-assist orifice about at least one of the plurality of fuel injection orifices to form an air-assist jacket about one or more of the fuel jets, the air-assist orifice configured to provide increased momentum to the fuel jet during injection into the plenum.

19. The method as claimed in claim 18, further comprising: defining one or more fuel passages extending through the fuel passage portion and the airfoil portion, the one or more fuel passages terminating at one or more of the plurality of fuel injection orifices formed through the airfoil portion; anddefining one or more air passage extending through the fuel passage portion and the airfoil portion, the one or more air passages terminating at one or more of the air-assist orifices formed through the airfoil portion and about the fuel injection orifice.

20. The method as claimed in claim 19, wherein each of the plurality of fuel injection orifices are in fluid communication with the one or more fuel passages to generate the plurality of fuel jets and wherein the air-assist orifice is in fluid communication with the one or more air passages to form the co-annular air-assist jacket about at least one of the plurality of fuel jets.