AERODYNAMICALLY-SHAPED BODY

Statement Regarding Federally Sponsored Development

Development for this invention was supported in part by Contract No. DE- FE0023968, awarded by the United States Department of Energy. Accordingly, the United States Government may have certain rights in this invention.

1 . Field Disclosed embodiments are generally related to combustion turbine engines, such as gas turbine engines and, more particularly, to injector assemblies disposed in a secondary combustion stage in a distributed combustion system (DCS).

2. Description of the Related Art

In gas turbine engines, fuel is delivered from a fuel source to a combustion section where the fuel is mixed with air and ignited to generate hot combustion products that define working gases. The working gases are directed to a turbine section where they effect rotation of a turbine rotor. It is known that production of NOx emissions from the burning fuel in the combustion section may be reduced by providing a portion of the fuel to be ignited axiaily downstream from a main combustion stage. This approach is referred to in the art as a distributed combustion system (DCS). See, for example, US patents 8,375,726 and 8,752, 386. Each of the above-listed patents is herein

incorporated by reference.

Brief Description Of The Drawings

FIG. 1 is a simplified fragmentary schematic of a disclosed combustor system (e.g., a distributed combustion system (DCS)) for a combustion turbine engine, such as a gas turbine engine. FIG. 2 illustrates a schematic of one non-limiting embodiment of a disclosed injector assembly including a fuel injector disposed in a stream-lined reactant-guiding structure (e.g., an aerodynamically-shaped scoop) arranged to deliver a jet of reactants to a combustion stage (e.g., secondary or axiai combustion stage) of the DCS.

FIGs. 3-5 illustrate respective schematics of further non-limiting embodiments of disclosed reactant-guiding structures.

FIGs. 8-7 illustrate respective schematics of non-limiting embodiments of disclosed reactant-guiding structures as may embody respective airfoils defining respective cambers.

FIG. 8 is a schematic of a disclosed reactant-guiding structure embodying a bell- mouth structure.

FIG. 9 is a cross-sectional view of a disclosed injector assembly where one or more orifices may be used for generating a jet in cross-flow of reactants in lieu of a reactant-guiding structure protruding into the combustion zone (e.g., a scoop).

Detailed Description

The inventors of the present invention have recognized certain issues that can arise in known distributed combustion systems (DCSs) where injector assemblies disposed in a secondary combustion stage (zone) that may be arranged axiaily downstream from a main combustion stage generally have a circular cross-sectional profile. These injector assemblies may comprise an assembly of a fuel nozzle within an air scoop having a blunt (e.g., circular) cross-sectional profile. The secondary combustion stage may also be referred to as an axial combustion stage. For example, the local peak temperatures near these axial-stage injector assemblies (or simply axial injectors) can approach the adiabatic flame temperature of the fuel/air mixture being injected in the secondary combustion stage. This adiabatic temperature can be substantially higher than temperatures in the main combustion stage, resulting in increased localized NOx generation near the axial injectors having the circular cross- sectional profile.

The present inventors have cleverly recognized that one source of elevated local peak temperatures near the injectors with the blunt profile may be the formation of recirculation zones in the leeward side of such injectors where vortex shedding may allow the formation of fuel-rich zones that, for example, can result from low entrainment of primary zone gases in a relatively high combustion residence time. Another source of elevated local temperatures may be a limited opportunity for a head end fluid (e.g., combustion products from the primary combustion zone) to entrain with the axial stage reactants prior to ignition of the axial stage flame resulting from premature ignition of the axial stage reactants due to the flame stabilizing effect of recirculating products in the recirculation zone. Additionally, non-optimized shear generated mixing between the axial stage reactants and primary zone gases can result in elevated flame temperatures due to low dilution of the axial stage reactants prior to ignition of the axial stage flame.

At least in view of the foregoing considerations and without limiting disclosed embodiments to any particular theoretical principle of operation, the present inventors propose axial injectors structured to eliminate or at least reduce the size of such recirculation zones, and additionally structured to increase the amount of entraining which occurs prior to ignition of the axial stage reactants. In order to reduce

recirculation of axial stage reactants in the leeward side of the jet, the present inventors propose replacing the blunt (e.g., circular) cross section injectors with injectors appropriately (e.g., aerodynamically) configured so that low-pressure regions

responsible for the formation of recirculation zones may be replaced by additional axial stage reactants.

Patent application (Attorney Docket No. 201517741 ) titled "Combustion System With Injector Assembly Aerodynamically-shaped Body And/or Ejection Orifices", proposes to include an array of aerodynamically configured (e.g., shaped) ejection orifices on one or more side walls of such aerodynamically-shaped injector structures.

This application further proposes respective combinations, such as a combination of a reactant-guiding structure comprising a stream-lined body (e.g., airfoil-shaped scoop) with an array of aerodynamically-shaped ejection orifices and/or circular-shaped orifices; or a combination of a blunt reactant-guiding structure (e.g., cylindrical-shaped scoop) with an array of aerodynamical!y-shaped ejection orifices.. This application is filed concurrently with the present application and is herein incorporated by reference in its entirety. With the proposed injector structures, in certain non-limiting embodiments, it is now feasible to achieve a reduced combustion residence time, which is conducive to reduce NOx emissions to be within acceptable levels at turbine inlet temperatures of approximately 1700°C (3200 ) and above.

In the following detailed description, various specific details are set forth in order to provide a thorough understanding of such embodiments. However, those skilled in the art will understand that embodiments of the present invention may be practiced without these specific details, that the present invention is not limited to the depicted embodiments, and that the present invention may be practiced in a variety of alternative embodiments. In other instances, methods, procedures, and components, which would be well-understood by one skilled in the art have not been described in detail to avoid unnecessary and burdensome explanation.

Furthermore, various operations may be described as multiple discrete steps performed in a manner that is helpful for understanding embodiments of the present invention. However, the order of description should not be construed as to imply that these operations need be performed in the order they are presented, nor that they are even order dependent, unless otherwise indicated. Moreover, repeated usage of the phrase "in one embodiment" does not necessarily refer to the same embodiment, although it may. It is noted that disclosed embodiments need not be construed as mutually exclusive embodiments, since aspects of such disclosed embodiments may be appropriately combined by one skilled in the art depending on the needs of a given application.

The terms "comprising", "including", "having", and the like, as used in the present application, are intended to be synonymous unless otherwise indicated. Lastly, as used herein, the phrases "configured to" or "arranged to" embrace the concept that the feature preceding the phrases "configured to" or "arranged to" is intentionally and specifically designed or made to act or function in a specific way and should not be construed to mean that the feature just has a capability or suitability to act or function in the specified way, unless so indicated.

FIG. 1 is a simplified fragmentary schematic of a combustor system 10 (e.g., a DCS) for a combustion turbine engine, such as a gas turbine engine. In one non-limiting embodiment, a plurality of circumferentia!ly-arrariged injector assemblies 12 is disposed in a combustion stage (e.g., axial stage) downstream from a main combustion stage 18 of the combustor system. As shown in FIG. 2, in one non-limiting embodiment, each injector assembly 12 may include a fuel injector 15 fluidiy coupled to a reactant-guiding structure 16 arranged to deliver a jet of reactants (e.g., fuel and air, schematically represented by arrow 7) to the combustion stage. Fuel injector 15 need not be located within injector assembly 12. For example, fuel injector 15 could be disposed upstream from injector assembly 12.

In one non-limiting embodiment, reactant-guiding structure 18 may be configured to form a streamlined body (e.g., an airfoil or other similar aerodynamically-shaped structure) relative to a flow 20 of a fluid (e.g., head end flow) to be mixed with the reactants delivered to the axial combustion stage. In one non-limiting embodiment, reactant-guiding structure 16 is arranged to deliver to the combustion stage a jet of the reactants through, for example, an open top side (e.g., an open side disposed at one end along the longitudinal axis of the streamlined body, as diagrammatically illustrated in FIG. 2) of reactant-guiding structure 18. This streamlined body may be effective to eliminate or at least reduce the size of the above-described recirculation zones, (schematically represented by oval 22), which in turn avoids or reduces excessive NOx formation rates that can result in a relatively high combustion residence time, and may be further effective to increase the amount of entraining which occurs prior to ignition of the axial stage reactants. For example, in the case of reactant-guiding structure 18 having a stream-lined body, formation of the axial stage flame (schematically

represented by jagged line 24) is believed to occur incrementally downstream

(compared to flame formation involving known blunt (e.g., circular) scoops) and this is effective to promote entrainment of the head end fluid with the axial stage reactants (schematically represented by curved lines 26) prior to ignition of the axial stage flame. In one non-!imiting embodiment, as may be appreciated in FIG, 3, reactant- guiding structure 18 may comprise a curved leading edge 30 and a trailing edge comprising a tapering tail section 32. This embodiment is believed to reduce the likelihood of flow separation in the main flow and thus reduces the likelihood of recirculation zones being generated around the axial jet.

In another non-limiting embodiment, as may be appreciated in FIG. 4, reactant- guiding structure 18 may comprise a non-curved leading edge 34 and a trailing edge comprising a tapering tail section 36. This embodiment while still effective to provide the recirculation region with additional axial stage reactants, additionally increases the cross sectional area of the region of axial reactant fluid which engages the cross-flow in the windward side of the jet. As will be appreciated by one skilled in the art, the windward side is the side of relatively high shear, and this high shear is effective to promote fast mixing. The idea behind increasing this area is to increase the amount of axial stage fluid which mixes quickly with the cross flow prior to the occurrence of a stable flame, thus reducing the flame temperature and overall NOx emissions. These structures may be conceptualized as non-limiting examples of a streamlined scoop, such as having a non-circular cross-sectional shape in lieu of known blunt (e.g., circular) scoops.

In still a further non-limiting embodiment, as may be appreciated in FIG. 5, reactant-guiding structure 16 (e.g., airfoil 38) may optionally include a blockage body 40 to direct an incremental amount of axial stage fuel/air mixture to locations proximate the wall in the leading edge of the airfoil. The idea behind this option is to cause rapid mixing of an incremental amount of axial stage fuel/air mixture with the head end cross flow with relatively less fuel/air mixture in the center of the reactant-guiding structure 18 where less mixing with the cross flow would occur.

In yet a further non-limiting embodiment, as may be appreciated in FIG. 6, reactant-guiding structure 16 may comprise an airfoil 42 defining a respective camber 44. This camber configuration may serve to incrementally improve large-scale mixing behavior within the secondary combustion zone. In one non-limiting embodiment, adjacent airfoils, such as in the plurality of circumferentia!ly arranged injector assemblies, may comprise respective cambers extending along a common direction. If the camber for each reactant-guiding structure 16 is in the same direction, the result would be to create large scale rotation within the flow which can improve mixing behavior. In an alternative non-limiting embodiment, as may be appreciated in FIG. 7, adjacent airfoils may comprise respective cambers 44, 45 extending along alternately varying directions, where, for example, resulting large scale flow features may interact between adjacent axial stage injectors, which in turn may be conducive to promote pre- flame mixing.

!n one non-limiting embodiment, the profile of the reactant-guiding structure 16 can be tailored to maximize the strength of the shear layers between the axial stage jet and the cross flow, and thus improve mixing with the head end flow. For example, as may be appreciated in FIG. 8, a bell-mouth 50 structure can be incorporated into reactant-guiding structure 16 , which is effective to reduce gradients in axial stage velocity, and thus maximize the velocity near the wail of reactant-guiding structure 16 for a given axial stage volumetric flow rate. In this manner, the velocity gradient between the axial jet and the cross flow near reactant-guiding structure 16 may be effectively increased, thus promoting enhanced mixing between axial stage reactants and the cross flow. As would be now appreciated by those skilled in the art, the increased velocity near the wall of reactant-guiding structure 16 has an additional benefit of increasing the convection coefficient associated with the axial reactant flow thus improving the cooling effectiveness of this flow. This cooling improvement would reduce the temperature of the walls of reactant-guiding structure 16, which wails are heated on their exterior surfaces by the hot-temperature head end flow which surrounds reactant-guiding structure 16. Disclosed non-limiting embodiments (e.g., cross-sections and/or profiles) in connection with reactant-guiding structure 16, as described above in the context of FIGs. 2-8, can also be utilized equally effectively on injector assemblies where one or more orifices are used for generating jet in cross-flow conditions in lieu of structures (e.g., scoops) that protrude into the secondary combustion stage. In this non-limiting embodiment, as illustrated in FIG. 9, a wall 60 separates the combustion zone from a manifold or other source of axial stage reactants (not shown). A respective orifice cross section 62 would embody a stream-lined shape (e.g., airfoil, etc), as described above, and the respective orifice profile can be optionally tailored to increase near wall velocities --as discussed in the context of the bell mouth structure illustrated in FIG. 8- such as by tailoring the cross section of the orifice to vary with depth along the orifice. For example, the area of the orifice cross section 62 gradually decreases in size as the orifice extends towards an exit on the side wall of the reactant-guiding structure.

It will be appreciated that each of the disclosed axial stage injection

embodiments can be applied in traditional secondary combustion zones as well as in applications where the axial combustion stage operates subject to elevated Mach number cross-flows, such as in a flow-accelerating cone 17 (FIG 1 ). Based on the narrowing cross-sectional profile of cone 17, as the flow travels from a cone inlet 19 to cone outlet 21 , the flow of combustion gases may be accelerated to a relatively high subsonic Mach (M) number, such as without limitation may comprise a range from approximately 0.3 M to approximately a 0.8 M. Accordingly, the combustion gases may flow through cone 17 with an increasing flow speed, and as a result, this flow of combustion gases can experience a decreasing static temperature in cone 17. That is, in one non-limiting embodiment the secondary combustion stage may be located in flow-accelerating cone 17 and injector assemblies 12 may be disposed in flow- accelerating cone 17. The streamlined shaped scoop has the additional benefit in high subsonic Mach number cross-flows in that the amount of flow blockage in the cross flow path, which occurs as a result of a given volumetric flow of axial stage reactants is reduced over known blunt (e.g., circular) scoop designs. As a result, local unwanted Mach number increases that otherwise would develop due to blocking effects from the presence of such blunt scoops in the path of the flow are reduced. Additionally, the reduced blockage is believed to be effective in minimizing the generation of oblique shock waves in high subsonic Mach number environments.

Increasing the Mach number of the cross flow introduces an additional NOx reduction benefit associated with the reduction in static temperature which accompanies a corresponding increase in the Mach number of the flow. Such static temperature reductions further reduce NOx emissions due to the reduced Arrhenius rate of formation of NOx compounds. For readers desirous of background information in connection with one non-limiting application involving a high Mach number combustion system, see patent application PCT/US2015/041948 filed on July 24, 2015, titled "Combustion System Having A Reduced Combustion Residence Time In A Combustion Turbine Engine", which is herein incorporated by reference in its entirety.

In operation, disclosed embodiments are expected to be conducive to a combustion system capable of realizing approximately a 65% combined cycle efficiency or greater in a gas turbine engine. Disclosed embodiments are also expected to realize a combustion system capable of maintaining stable operation at turbine inlet

temperatures of approximately 1700°C and higher while maintaining a relatively low level of NOx emissions, and acceptable temperatures in components of the engine without an increase in cooling air consumption.

While embodiments of the present disclosure have been disclosed in exemplary forms, it will be apparent to those skilled in the art that many modifications, additions, and deletions can be made therein without departing from the spirit and scope of the invention and its equivalents, as set forth in the following claims.