Patent ID: 12215870

DETAILED DESCRIPTION

Various embodiments are discussed in detail below. While specific embodiments are discussed, this is done for illustration purposes only. A person skilled in the relevant art will recognize that other components and configurations may be used without departing from the spirit and scope of the present disclosure.

As noted above, it has been suggested that a hydrogen fuel may be utilized to achieve improvements in the emissions from commercial aircraft. Hydrogen fuel, however, poses a number of challenges as compared to combustible hydrocarbon liquid fuel. Hydrogen fuel, for example, is a reactive fuel that burns at higher temperatures than combustible hydrocarbon liquid fuel. When hydrogen fuel is used in current gas turbine engines with rich burn combustors, the higher combustion temperature requires additional water (or other diluent) additions to reduce the production of nitrogen oxides (“NOx”), as compared combustible hydrocarbon liquid fuel. Injecting additional water quenches flame resulting in higher production of carbon monoxide and reduces efficiency.

The present disclosure discusses ways to achieve NOx emissions targets with improved efficiency and less diluent, such as water, consumption than in conventional rich burn combustors. The combustors and methods discussed herein distribute the injection of fuel and diluent into a combustion chamber of a combustor with a portion of the fuel and/or a portion of the diluent being injected at the forward end of the combustion chamber and the remaining portion of the fuel and/or diluent being injected downstream in a bulk airflow direction. Such a combustor and method have advantages of reducing the average dwell time of the fuel and diluent in the combustor and reducing the amount of NOx that is produced. In addition, the diluent can be targeted towards certain regions of the combustor. Some regions of the combustor have be hot spots, where the temperature is locally elevated relative to the surrounding temperature. Hot spots are locations of increased NOx production. Diluent can be targeted towards these hot spots creating a more uniform temperature distribution throughout the combustor and reducing the total amount of diluent, such as water, consumed compared to a combustor where all of the water is introduced at the forward end of the combustor.

The combustors and methods discussed herein are particularly suited for use in combustors of gas turbine engines using a highly reactive fuel such as hydrogen fuel (diatomic hydrogen fuel) or hydrogen enriched fuel. As discussed above, such fuels burn hotter and thus may create higher amounts of NOx. The combustors and methods discussed herein enable these fuel systems to be used while achieving NOx emissions targets with improved efficiency and less diluent, such as water, consumption.

A particularly suitable application for the combustors and methods discussed herein is in gas turbine engine used on aircraft.FIG.1is a perspective view of an aircraft10that may implement various preferred embodiments. The aircraft10includes a fuselage12, wings14attached to the fuselage12, and an empennage16. The aircraft10also includes a propulsion system that produces a propulsive thrust required to propel the aircraft10in flight, during taxiing operations, and the like. The propulsion system for the aircraft10shown inFIG.1includes a pair of engines100. In this embodiment, each engine100is attached to one of the wings14by a pylon18in an under-wing configuration. Although the engines100are shown attached to the wing14in an under-wing configuration inFIG.1, in other embodiments, the engine100may have alternative configurations and be coupled to other portions of the aircraft10. For example, the engine100may additionally or alternatively include one or more aspects coupled to other parts of the aircraft10, such as, for example, the empennage16, and the fuselage12.

As will be described further below with reference toFIG.2, the engines100shown inFIG.1are gas turbine engines that are each capable of selectively generating a propulsive thrust for the aircraft10. The amount of propulsive thrust may be controlled at least in part based on a volume of fuel provided to the gas turbine engines100via a fuel system200. The fuel is stored in a fuel tank212of the fuel system200. As shown inFIG.1, at least a portion of the fuel tank212is located in each wing14and a portion of the fuel tank212is located in the fuselage12between the wings14. The fuel tank212, however, may be located at other suitable locations in the fuselage12or the wing14. The fuel tank212may also be located entirely within the fuselage12or the wing14. The fuel tank212may also be separate tanks instead of a single, unitary body, such as, for example, two tanks each located within a corresponding wing14. A diluent is also provided to the gas turbine engines100via the fuel system200. The diluent is stored in a diluent tank214. The diluent tank214may be located on the aircraft10in the same positions as the fuel tank212discussed above.

Although the aircraft10shown inFIG.1is an airplane, the embodiments described herein may also be applicable to other aircraft10, including, for example, helicopters. In addition, the embodiments described herein may also be applicable to other applications where hydrogen is used as a fuel. The engines described herein are gas turbine engines, but the embodiments described herein also may be applicable to other engines. The engine100may be used in various other applications including stationary power generation systems and other vehicles beyond the aircraft10explicitly described herein, such as boats, ships, cars, trucks, and the like.

For the embodiment depicted, the engine100is a high bypass turbofan engine. The engine100may also be referred to as a turbofan engine100herein.FIG.2is a schematic, cross-sectional view of one of the engines100used in the propulsion system for the aircraft10shown inFIG.1. The cross-sectional view ofFIG.2is taken along line2-2inFIG.1. The turbofan engine100has an axial direction A (extending parallel to a longitudinal centerline101, shown for reference inFIG.2), a radial direction R, and a circumferential direction. The circumferential direction (not depicted inFIG.2) extends in a direction rotating about the axial direction A. The turbofan engine100includes a fan section102and a turbomachine104disposed downstream from the fan section102.

The turbomachine104depicted inFIG.2includes a tubular outer casing106that defines an annular inlet108. The outer casing106encases, in a serial flow relationship, a compressor section including a booster or low-pressure (LP) compressor110and a high-pressure (HP) compressor112, a combustion section300(also referred to herein as a combustor300), a turbine section including a high-pressure (HP) turbine116and a low-pressure (LP) turbine118, and a jet exhaust nozzle section120. The compressor section, the combustor300, and the turbine section together define at least in part a core air flowpath121extending from the annular inlet108to the jet exhaust nozzle section120. The turbofan engine further includes one or more drive shafts. More specifically, the turbofan engine includes a high-pressure (HP) shaft or spool122drivingly connecting the HP turbine116to the HP compressor112, and a low-pressure (LP) shaft or spool124drivingly connecting the LP turbine118to the LP compressor110.

The fan section102shown inFIG.2includes a fan126having a plurality of fan blades128coupled to a disk130in a spaced-apart manner. The fan blades128and the disk130are rotatable, together, about the longitudinal centerline (axis)101by the LP shaft124. The disk130is covered by rotatable front hub132aerodynamically contoured to promote an airflow through the plurality of fan blades128. Further, an annular fan casing or outer nacelle134is provided, circumferentially surrounding the fan126and/or at least a portion of the turbomachine104. The nacelle134is supported relative to the turbomachine104by a plurality of circumferentially spaced outlet guide vanes136. A downstream section138of the nacelle134extends over an outer portion of the turbomachine104so as to define a bypass airflow passage140therebetween.

The turbofan engine100is operable with the fuel system200and receives a flow of fuel from the fuel system200. As will be described further below, the fuel system200includes a fuel delivery assembly202providing the fuel flow from the fuel tank212to the engine100, and more specifically to a plurality of primary fuel nozzles340and a plurality of secondary fuel nozzles400(not labeled inFIG.2; seeFIG.3) of the combustor300of the turbomachine104of the turbofan engine100.

The turbofan engine100also includes various accessory systems to aid in the operation of the turbofan engine100and/or an aircraft including the turbofan engine100. For example, the turbofan engine100may include a main lubrication system152, a compressor cooling air (CCA) system154, an active thermal clearance control (ATCC) system156, and generator lubrication system158, each of which is depicted schematically inFIG.2. The main lubrication system152is configured to provide a lubricant to, for example, various bearings and gear meshes in the compressor section, the turbine section, the HP spool122, and the LP shaft124. The lubricant provided by the main lubrication system152may increase the useful life of such components and may remove a certain amount of heat from such components. The compressor cooling air (CCA) system154provides air from one or both of the HP compressor112or LP compressor110to one or both of the HP turbine116or LP turbine118. The active thermal clearance control (ATCC) system156cools a casing of the turbine section to maintain a clearance between the various turbine rotor blades and the turbine casing within a desired range throughout various engine operating conditions. The generator lubrication system158provides lubrication to an electronic generator (not shown), as well as cooling/heat removal for the electronic generator. The electronic generator may provide electrical power to, for example, a startup electrical motor for the turbofan engine100and/or various other electronic components of the turbofan engine100and/or an aircraft including the turbofan engine100.

Heat from these accessory systems152,154,156,158, and other accessory systems, may be provided to various heat sinks as waste heat from the turbofan engine100during operation, such as to various vaporizers220, as discussed below. Additionally, the turbofan engine100may include one or more heat exchangers162within, for example, the turbine section or jet exhaust nozzle section120for extracting waste heat from an airflow therethrough to also provide heat to various heat sinks, such as the vaporizers220, discussed below.

It will be appreciated, however, that the turbofan engine100discussed herein is provided by way of example only. In other embodiments, any other suitable engine may be utilized with aspects of the present disclosure. For example, in other embodiments, the engine may be any other suitable gas turbine engine, such as a turboshaft engine, a turboprop engine, a turbojet engine, and the like. In such a manner, it will further be appreciated that, in other embodiments, the gas turbine engine may have other suitable configurations, such as other suitable numbers or arrangements of shafts, compressors, turbines, fans, etc. Further, although the turbofan engine100is shown as a direct drive, fixed-pitch turbofan engine100, in other embodiments, a gas turbine engine may be a geared gas turbine engine (i.e., including a gearbox between the fan126and shaft driving the fan, such as the LP shaft124), may be a variable pitch gas turbine engine (i.e., including a fan126having a plurality of fan blades128rotatable about their respective pitch axes), etc. Further, still, in alternative embodiments, aspects of the present disclosure may be incorporated into, or otherwise utilized with, any other type of engine, such as reciprocating engines. Additionally, in still other exemplary embodiments, the exemplary turbofan engine100may include or be operably connected to any other suitable accessory systems. Additionally, or alternatively, the exemplary turbofan engine100may not include or be operably connected to one or more of the accessory systems152,154,156,158, and162, discussed above.

The fuel system200of this embodiment is configured to store the fuel for the engine100in the fuel tank212and to deliver the fuel to the engine100via the fuel delivery assembly202. The fuel delivery assembly202includes tubes, pipes, and the like, to fluidly connect the various components of the fuel system200to the engine100. As discussed above, the engine100, and in particular the combustor300discussed herein may be particularly suited for use with highly reactive fuels such as hydrogen fuel (diatomic hydrogen) or hydrogen enriched fuels. In the embodiments discussed herein, the fuel is a hydrogen fuel.

The fuel tank212may be configured to hold the hydrogen fuel at least partially within the liquid phase, and may be configured to provide hydrogen fuel to the fuel delivery assembly202substantially completely in the liquid phase, such as completely in the liquid phase. For example, the fuel tank212may have a fixed volume and contain a volume of the hydrogen fuel in the liquid phase (liquid hydrogen fuel). As the fuel tank212provides hydrogen fuel to the fuel delivery assembly202substantially completely in the liquid phase, the volume of the liquid hydrogen fuel in the fuel tank212decreases and the remaining volume in the fuel tank212is made up by, for example, hydrogen in the gaseous phase (gaseous hydrogen). It will be appreciated that as used herein, the term “substantially completely” as used to describe a phase of the hydrogen fuel refers to at least 99% by mass of the described portion of the hydrogen fuel being in the stated phase, such as at least 97.5%, such as at least 95%, such as at least 92.5%, such as at least 90%, such as at least 85%, or such as at least 75% by mass of the described portion of the hydrogen fuel being in the stated phase.

To store the hydrogen fuel substantially completely in the liquid phase, the hydrogen fuel is stored in the fuel tank212at very low (cryogenic) temperatures. For example, the hydrogen fuel may be stored in the fuel tank212at about −253 Deg. Celsius or less at atmospheric pressure, or at other temperatures and pressures to maintain the hydrogen fuel substantially in the liquid phase. The fuel tank212may be made from known materials such as titanium, Inconel®, aluminum, or composite materials. The fuel tank212and the fuel system200may include a variety of supporting structures and components to facilitate storing the hydrogen fuel in such a manner.

The liquid hydrogen fuel is supplied from the fuel tank212to the fuel delivery assembly202. The fuel delivery assembly202may include one or more lines, conduits, etc., configured to carry the hydrogen fuel between the fuel tank212and the engine100. The fuel delivery assembly202thus provides a flow path of the hydrogen fuel from the fuel tank212to the engine100. The hydrogen fuel is delivered to the engine by the fuel delivery assembly202in the gaseous phase, the supercritical phase, or both (at least one of the gaseous phase and the supercritical phase). The fuel system200thus includes a vaporizer220in fluid communication with the fuel delivery assembly202to heat the liquid hydrogen fuel flowing through the fuel delivery assembly202. The vaporizer220is positioned in the flow path of the hydrogen fuel between the fuel tank212and the engine100. The vaporizer220may be positioned at least partially within the fuselage12or the wing14, such as at least partially within the wing14. The vaporizer220may, however, be positioned at other suitable locations in the flow path of the hydrogen between the fuel tank212and the engine100. For example, the vaporizer220may be positioned external to the fuselage12and the wing14and positioned at least partially within the pylon18or the engine100. When positioned in the engine100, the vaporizer may be located in the nacelle134, for example. Although only one vaporizer220is shown inFIG.2, the fuel system200may include multiple vaporizers220. For example, when a vaporizer220is positioned in the engine100or in the pylon18and functions as a primary vaporizer configured to operate once the engine100is in a thermally stable condition, another vaporizer220is positioned upstream of the primary vaporizer and proximate to the fuel tank212and functions as a primer vaporizer during start-up (or prior to start-up) of the engine100.

The vaporizer220is in thermal communication with at least one heat source222,224. In this embodiment, the vaporizer220is in thermal communication with a primary heat source222and an auxiliary heat source224. In this embodiment, primary heat source222is waste heat from the engine100, and the vaporizer220is thus thermally connected to at least one of the main lubrication system152, the compressor cooling air (CCA) system154, the active thermal clearance control (ATCC) system156, the generator lubrication system158, and the heat exchangers162to extract waste heat from the engine100to heat the hydrogen fuel. In such a manner, it will be appreciated that the vaporizer220is configured to operate by drawing heat from the primary heat source222once the engine100is capable of providing enough heat, via the auxiliary heat source224, to the vaporizer220, in order to facilitate operation of the vaporizer220.

The vaporizer220may be heated by any suitable heat source, and, in this embodiment, for example, the auxiliary heat source224is a heat source external to the engine100. The auxiliary heat source224may include, for example, an electrical power source, a catalytic heater or burner, and/or a bleed airflow from an auxiliary power unit. The auxiliary heat source224may be integral to the vaporizer220, such as when the vaporizer220includes one or more electrical resistance heaters, or the like, that are powered by the electrical power source. In this configuration the auxiliary heat source224may provide heat for the vaporizer220independent of whether or not the engine100is running and can be used, for example, during start-up (or prior to start-up) of the engine100.

As noted, the vaporizer220is in communication with the flow of the hydrogen fuel through the fuel delivery assembly202. The vaporizer220is configured to draw heat from at least one of the primary heat source222and the auxiliary heat source224to heat the flow of hydrogen fuel from a substantially completely liquid phase to a substantially completely gaseous phase or to a substantially completely supercritical phase.

The fuel delivery assembly202also includes a high-pressure pump232to induce the flow of the hydrogen fuel through the fuel delivery assembly202to the engine100. The high-pressure pump232may generally be the primary source of pressure rise in the fuel delivery assembly202between the fuel tank212and the engine100. The high-pressure pump232may be configured to increase a pressure in the fuel delivery assembly202to a pressure greater than a pressure within a combustion chamber330of the combustor300of the engine100. For example, the high-pressure pump232may be configured to increase a pressure in the fuel delivery assembly202to at least four hundred pounds per square inch (“psi”), such as to at least five hundred psi, such as to at least six hundred psi, such as to at least seven hundred psi, such as to at least seven hundred fifty psi, such as up to two thousand psi.

The high-pressure pump232is positioned within the flow of hydrogen fuel in the fuel delivery assembly202at a location downstream of the vaporizer220. In this embodiment, the high-pressure pump232is positioned external to the fuselage12and the wing14, and is positioned at least partially within the pylon18, or at least partially within the engine100. More specifically, the high-pressure pump232is positioned within the engine100. With the high-pressure pump232located in such a position, the high-pressure pump232may be any suitable pump configured to receive the flow of hydrogen fuel in substantially completely a gaseous phase or a supercritical phase. It will be appreciated, however, that, in other embodiments, the high-pressure pump232may be positioned at any other suitable locations, including other positions within the flow path of the hydrogen fuel. For example, the high-pressure pump232may be located upstream of the vaporizer220and may be configured to receive the flow of hydrogen fuel through the fuel delivery assembly202in a substantially completely liquid phase.

As will be discussed further below, a diluent is also used during combustion of the fuel. The diluent is stored in the diluent tank214, as discussed above, and delivered to the engine100via a diluent delivery assembly204. The diluent delivery assembly204includes tubes, pipes, and the like, to fluidly connect the various components used to deliver the diluent to the engine100. In this embodiment, the diluent is water, but any suitable diluent may be used, including, for example, nitrogen and carbon dioxide. The diluent delivery assembly204also includes a diluent pump234to induce the flow of the diluent through the diluent delivery assembly204to the engine100. The diluent pump234may generally be the primary source of pressure rise in the diluent delivery assembly204between the diluent tank214and the engine100. The diluent pump234may be configured to increase a pressure in the diluent delivery assembly204to a pressure greater than a pressure within a combustion chamber330of the combustor300of the engine100.

The diluent pump234is positioned within the flow of diluent in the diluent delivery assembly204at a location downstream of the diluent tank214. In this embodiment, the diluent pump234is positioned external to the fuselage12and the wing14, and is positioned at least partially within the pylon18, or at least partially within the engine100. More specifically, the diluent pump234is positioned within the engine100. It will be appreciated, however, that, in other embodiments, the diluent pump234may be positioned at any other suitable locations, including other positions within the flow path of the diluent.

The fuel system200also includes a metering system240in fluid communication with the fuel delivery assembly202and the diluent delivery assembly204. Any suitable metering system240may be used, such as a series of metering valves and proportioning valves. As shown inFIG.3, for example, a fuel metering valve242and a diluent metering valve244may be placed in fluid communication with the fuel delivery assembly202and the diluent delivery assembly204, respectively. The fuel delivery assembly202is configured to provide the fuel metering valve242, and the fuel metering valve242is configured to receive, hydrogen fuel, and, likewise, the diluent delivery assembly204is configured to provide the diluent metering valve244, and the diluent metering valve244is configured to receive, the diluent. The fuel metering valve242and the diluent metering valve244are further configured to provide the flow of fuel and diluent, respectively, to the engine100in a desired manner. The fuel metering valve242is configured to provide a desired volume of hydrogen fuel, at, for example, a desired flow rate, to a combustion chamber330of the combustor300, and the diluent metering valve244is configured to provide a desired volume of diluent, at, for example, a desired flow rate, to a combustion chamber330of the combustor300. Adjusting the fuel metering valve242changes the volume of fuel provided to the combustion chamber330of the combustor300and, thus, changes the amount of propulsive thrust produced by the engine100to propel the aircraft10. The diluent metering valve244may be adjusted in conjunction with the fuel metering valve242to change the volume of diluent provided to the combustion chamber330of the combustor300as the amount of fuel is changed. Additional details of the metering system240will be discussed further below.

FIGS.3and4show the combustor300of the engine100according to an embodiment of the present disclosure.FIG.3is a detail view showing detail3inFIG.2, and, asFIG.2is a cross-sectional view,FIG.3is also a cross-sectional view of the combustor300.FIG.4is a perspective view of the combustor300. The combustor300includes a combustor case310and a combustor liner320. The combustor liner320of this embodiment has a combustor inner liner320A and a combustor outer liner320B. A combustion chamber330is formed within the combustor liner320. The combustor liner320, and thus also the combustion chamber330, has a forward end322and an outlet324. A primary fuel nozzle340is positioned at the forward end322of the combustion chamber330. The primary fuel nozzle340of this embodiment is part of a swirler/fuel nozzle assembly342. In this embodiment, the combustor300is an annular combustor300and a plurality of primary fuel nozzles340is arranged in an annular configuration as shown inFIG.4with the plurality of primary fuel nozzles340(the swirler/fuel nozzle assemblies342) aligned in a circumferential direction of the combustor.

As discussed above, the compressor section, the combustor300, and the turbine section form, at least in part, the core air flowpath121extending from the annular inlet108to the jet exhaust nozzle section120. Air entering through the annular inlet108is compressed by blades of a plurality of fans of the LP compressor110and HP compressor112. A portion of the compressed air (primary air302) enters the forward end322of the combustion chamber330. Fuel is injected by the primary fuel nozzle340into the primary air302and mixed with the primary air302. As noted above, the primary fuel nozzle340of this embodiment is part of a swirler/fuel nozzle assembly342. The swirler/fuel nozzle assembly342includes a swirler344that is used to generate turbulence in the primary air302. The primary fuel nozzle340injects fuel into the turbulent airflow of the primary air302and the turbulence promotes rapid mixing of the fuel with the primary air302.

The mixture of fuel and compressed air is combusted in the combustion chamber330, generating combustion gases (combustion products), which accelerate as the combustion gases leave the combustion chamber330. The products of combustion are accelerated as the products are expelled through the outlet324to drive the engine100. The primary air302thus flows in a bulk airflow direction (indicated by the arrow B inFIG.3) from the forward end322of the combustion chamber330to the outlet324. The terms “downstream” and “upstream” may be used to describe the position of components in the combustor300or locations in the combustor300relative to the direction of the bulk airflow B. Much of the fuel injected by the primary fuel nozzle340is combusted in a primary combustion zone332in the region of the combustor300directly downstream of the primary fuel nozzle340. The combusted fuel air mixture is then accelerated through the outlet324to turn the turbines (e.g., drive the turbine blades) of the HP turbine116and the LP turbine118. As discussed above the HP turbine116and the LP turbine118, among other things, drive the LP compressor110and HP compressor112.

Another portion of the compressed air (inner liner dilution air304A) flows around the outside of the combustor liner320and is introduced into the combustion chamber330by dilution holes326formed in the combustor inner liner320A at positions downstream of the primary fuel nozzle340. The inner liner dilution air304A helps quench combustion gasses from primary zone before being introduced into the turbine section of the engine100. The inner liner dilution air304A bypasses the forward end322of the combustion chamber330and the primary combustion zone332. The inner liner dilution air304A is introduced into a secondary combustion zone334, which, in this embodiment, is the portion of the combustion chamber330downstream of the primary combustion zone332. The inner liner dilution air304A flows into the combustion chamber330through at least one dilution hole326in the combustor inner liner320A. The combustion products from the primary combustion zone332flow in a cross flowing direction with the inner liner dilution air304A. Similarly, outer liner dilution air304B flows into the combustion chamber330through at least one dilution hole326in the combustor outer liner320B. The combustion products from the primary combustion zone332flow in cross flowing direction with the outer liner dilution air304B. Collectively the inner liner dilution air304A and the outer liner dilution air304B may be referred to as dilution air304herein. Dilution air304introduced through dilution holes326reduces the temperature in a core region336of the combustion chamber330. More specifically, the inner liner dilution air304A and the outer liner dilution air304B penetrate inside the combustor and reduce the high temperature in the core region336of the combustion chamber330.

The combustor300of this embodiment also includes a plurality of secondary fuel nozzles400, and, as will be discussed in more detail below, the secondary fuel nozzles400are configured to inject a portion of the fuel into the combustion chamber330. Each secondary fuel nozzle400is positioned downstream of the primary fuel nozzle340. In this embodiment, the secondary fuel nozzle400projects into the combustion chamber330and through the combustor liner320. More specifically in this embodiment, the secondary fuel nozzle400projects through the combustor outer liner320B and is located within a corresponding dilution hole326such that outer liner dilution air304B flows around the outside of the secondary fuel nozzle400(see alsoFIGS.6and7). The secondary fuel nozzle400does not need to be located within the dilution hole326, however, and may project through the combustor liner320at other locations. Locating the secondary fuel nozzle400in the dilution hole326, however, has the advantage of using the outer liner dilution air304B to help cool the secondary fuel nozzle400and thus extend the life of the secondary fuel nozzle400. Additionally, the dilution air304that flows around secondary fuel nozzle400and along the secondary fuel nozzle400can have higher penetration into the core region336of the combustion chamber330to efficiently reduce temperature in the core region336of the combustion chamber330and further reduce NOx production. In the annular combustor300of this embodiment, the plurality of secondary fuel nozzles400are aligned in a direction transverse to the bulk airflow direction, as shown inFIG.4. When the combustor300is an annular combustor300, the plurality of secondary fuel nozzles400may be aligned in the circumferential direction of the combustor300. This direction is transverse to the bulk airflow direction B and may be referred to herein as a lateral direction.

A portion of the fuel is injected into the combustor300using the plurality of primary fuel nozzles340. This portion of the fuel is referred to herein as the primary portion of the fuel. The remainder of the fuel is injected into the combustor300using the plurality of secondary fuel nozzles400, and in this embodiment, the remainder of the fuel is injected into the combustor300in a transverse direction to the bulk airflow direction B using the plurality of secondary fuel nozzles400. This portion of the fuel is referred to herein as the secondary portion of the fuel. By this configuration and strategy, NOx production can be reduced as compared to the configuration where one hundred percent of the fuel is injected by the primary fuel nozzles340. One effect of injecting fuel downstream of the forward end322of the combustion chamber330with the plurality of the secondary fuel nozzles400is that the secondary portion of the fuel has a shorter dwell time in the combustion chamber330, which can contribute to the reduction in NOx production. Preferably, the primary portion of the fuel is from thirty percent to eighty percent of the fuel being injected into the combustion chamber over a given time interval, and the secondary portion of the fuel is the remaining portion of the fuel injected into the combustion chamber over the time interval (e.g., twenty to seventy percent). These percentages may also be determined for a segment of the combustor300. In the annular combustor300shown inFIG.4, for example, the segment is formed by the angular span formed by at least one primary fuel nozzle340. In this example, the segment includes one primary fuel nozzle340and two secondary fuel nozzles400(e.g., secondary fuel nozzle400A and secondary fuel nozzle400B). Likewise, there can also be a single secondary fuel nozzle400or a plurality of secondary fuel nozzles400within each segment. When a plurality of secondary fuel nozzles400is used, the secondary fuel nozzles400may be placed at the same axial location within the combustion chamber330and may be aligned in a direction transverse to the bulk airflow direction B, such as the lateral direction. Alternatively, the secondary fuel nozzles400may be placed at different axial locations and may be aligned in the in the axial direction (such as in the bulk airflow direction B) or may have different circumferential locations inside the segment of the combustion chamber330.

As discussed above, the fuel metering valve242is configured to provide a desired volume of fuel, at, for example, a desired flow rate, to the combustion chamber330. As shown inFIG.3, the fuel metering valve242is fluidly connected to a primary fuel manifold252and a secondary fuel manifold254. The primary fuel manifold252and the secondary fuel manifold254distribute (provide) the fuel received to the primary fuel nozzles340and secondary fuel nozzles400, respectively. After being metered to the desired flow rate by the fuel metering valve242, the fuel is then split based on the desired percentage distribution to the primary fuel manifold252and the secondary fuel manifold254to be subsequently injected into the combustion chamber330by the primary fuel nozzles340and the secondary fuel nozzles400. The ratio (percentage) of fuel injected by the primary fuel nozzles340and the secondary fuel nozzles400may be set (or static) for a given combustor300, but this ratio may also be adjustable across different operating conditions of engine100. For example, the metering system240may also include at least one fuel proportioning valve246that can be used to adjust the amount of fuel distributed to the primary fuel manifold252and the secondary fuel manifold254.

As noted above, a diluent, such as at least one of water, nitrogen, and carbon dioxide, may be injected into the combustion chamber330and used during the combustion of the fuel. The diluent may be injected into the forward end322of the combustion chamber330using the primary fuel nozzle340. In this embodiment, the primary fuel nozzle340is configured to inject both fuel and diluent into the primary combustion zone332, but other suitable configurations may be used including, for example, where the primary fuel nozzle340injects fuel and separate diluent nozzles configured to inject the diluent into the primary combustion zone332are used.

Although one hundred percent of the diluent may be injected into the primary combustion zone332by, for example, the primary fuel nozzle340, the combustor300disclosed herein injects diluent downstream of the primary fuel nozzle340, targeting high temperature regions (hot spots) in the combustion chamber330. Introducing diluent targeted to the location of the hot spots creates a more uniform temperature distribution throughout the combustor while reducing the total amount of diluent, such as water, consumed compared to a combustor300in which all of the diluent (e.g., water) is introduced to the forward end322of the combustion chamber330. In addition, reducing or eliminating these hot spots further reduces NOx production.

In this embodiment, the secondary fuel nozzle400is configured to inject both fuel and diluent into the combustion chamber330, and the secondary fuel nozzle400is configured to direct the diluent towards hot spots in the combustion chamber330. Details of the secondary fuel nozzle400of this embodiment are discussed below, and, in this configuration, both secondary fuel nozzles400(labeled400A and400B inFIG.4) in a segment inject both fuel and diluent. The combustor300, however, is not so limited and other suitable configurations may be used. In each segment, for example, one secondary fuel nozzle400A may be configured to inject the secondary portion of the fuel into the segment and the segment may also include a diluent nozzle400B configured to inject diluent into the segment. In such a configuration, the combustor300may thus include a plurality of diluent nozzles400B downstream of the plurality of primary fuel nozzles340in the bulk airflow direction, and the plurality of secondary fuel nozzles400A and the plurality of diluent nozzles400B may be aligned in a direction transverse to the bulk airflow direction B, such as the lateral direction.

In this embodiment, a portion of the diluent is injected into the combustor300using the plurality of primary fuel nozzles340. This portion of the diluent is referred to herein as the primary portion of the diluent. The remainder of the diluent is injected into the combustor300using the plurality of secondary fuel nozzles400(or in the alternative embodiment discussed above the plurality of diluent nozzles400B). This portion of the diluent is referred to herein as the secondary portion of the diluent. Preferably, the primary portion of the diluent is from twenty percent to eighty percent of the diluent being injected into the combustion chamber over a given time interval, and the secondary portion of the diluent is the remaining portion of the diluent injected into the combustion chamber over the time interval (e.g., twenty to eighty percent). These percentages may also be determined for a segment of the combustor300, such as the segment discussed above with reference toFIG.4.

As discussed above, the diluent metering valve244is configured to provide a desired volume of diluent, at, for example, a desired flow rate, to the combustion chamber330. As shown inFIG.3, the diluent metering valve244is fluidly connected to a primary diluent manifold256and a secondary diluent manifold258. The primary diluent manifold256and the secondary diluent manifold258distribute (provide) the diluent received to the primary fuel nozzles340and secondary fuel nozzles400, respectively. After being metered to the desired flow rate by the diluent metering valve244, the diluent is then split based on the desired percentage distribution to the primary diluent manifold256and the secondary diluent manifold258to be subsequently injected into the combustion chamber330by the primary fuel nozzles340and the secondary fuel nozzles400. The ratio (percentage) of diluent injected by the primary fuel nozzles340and the secondary fuel nozzles400may be set (or static) for a given combustor300, but this ratio may also be adjustable. For example, the metering system240may also include at least one diluent proportioning valve248that can be used to adjust the amount of diluent distributed to the primary diluent manifold256and the secondary diluent manifold258.

FIG.5is a flowchart showing a method of operating a gas turbine engine (such as engine100) according to an embodiment of the present disclosure. AlthoughFIG.5is a flowchart depicting a linear sequence, the method should be appreciated as having the various steps discussed below occur continuously or simultaneously with each other. In addition, the steps may have another order other than as depicted inFIG.5. In step S505, air is compressed, using, for example, the LP compressor110and the HP compressor112. Then in step S510, the compressed air is directed (flows) through the combustion chamber330. As discussed above, the primary air302flows in the bulk airflow direction B through the combustion chamber330, and dilution air304is introduced through the dilution holes326. The primary portion of the fuel and the primary portion of the diluent are injected into the forward end322of the combustion chamber330in steps S515and S525using, for example, the primary fuel nozzle340, as discussed above. The primary portion of the fuel and the primary portion of the diluent is flowing throughout different power conditions of the engine100(from 0% to 100% power of the engine100). As depicted in step S520ofFIG.5, the mixture of the primary portion of the fuel and compressed air is combusted in the primary combustion zone332.

The secondary portion of the fuel and the secondary portion of the diluent are injected into the combustion chamber330in steps S530and S535. In a preferred embodiment, the secondary portion of the fuel and the secondary portion of the diluent are injected into the combustion chamber330under higher power conditions. Herein, lower power conditions are 20% of the total engine power and below, and higher power conditions are powers of the engine100that are greater than 20%. In an embodiment, only the primary portion of the fuel is flowing at lower power conditions, and both the primary portion of the fuel and the secondary portion of the fuel is flowing at higher power conditions. At lower power conditions, one of the primary portion of the diluent and the secondary portion of the diluent may be flowing or both the primary portion and the secondary portion of the diluent may be flowing. At higher power conditions, both the primary portion and the secondary portion of the diluent is flowing. The primary portion and the secondary portion of the fuel and/or the diluent can be operated in a way to maximize benefit on emission and efficiency. In the combustion chamber330, the additional fuel not combusted in step S520(only a residual fuel of the primary portion at lower power conditions or both the residual fuel of the primary portion of the fuel and the secondary portion of the fuel at higher power conditions) and the compressed air are mixed and the fuel is combusted in step S540. The combusted products exit the combustion chamber330by the outlet324and then drives a turbine, such as the HP turbine116and the LP turbine118in step S545. Additional details of the method are described herein relative to, for example, the aforementioned components (e.g., the primary fuel nozzle340and the secondary fuel nozzle400).

FIGS.6and7are cross-sectional views of the secondary fuel nozzle400that is configured to inject both fuel and diluent.FIG.6is a cross-sectional view of the secondary fuel nozzle400taken along line6-6inFIG.4and, thus, the left side ofFIG.6is toward the forward end322of the combustion chamber330and the right side ofFIG.6is toward the outlet324of the combustion chamber330.FIG.7is a cross-sectional view of the secondary fuel nozzle400taken along line7-7inFIG.3and, thus, is a view looking toward the forward end322of the combustion chamber330.

The secondary fuel nozzle400of this embodiment is cylindrical and, thus, has a radial direction and an axial direction. The secondary fuel nozzle400of this embodiment has a cross section taken in a direction orthogonal to the axial direction of the secondary fuel nozzle400along line10-10inFIG.3. This cross section is shown inFIG.10A, and as can be seen inFIG.10A, this cross section is circular. Any suitable geometry may be used for the secondary fuel nozzle400, however. For example, the secondary fuel nozzle400may have an airfoil shape, such as a teardrop shape as shown inFIG.10B. The teardrop or other airfoil shape may be advantageous to reduce the wake in the airflow past the secondary fuel nozzle400. The secondary fuel nozzle400is mounted to the combustor case310and extends through the space between the combustor case310and combustor liner320before projecting into the combustion chamber330. As discussed above, the secondary fuel nozzle400projects through the dilution hole326, and, in this embodiment, the diameter of the dilution hole326is larger than the diameter of the secondary fuel nozzle400and sized to allow desired amount of dilution air304to flow around the secondary fuel nozzle400and into the combustion chamber330. As noted above, the secondary fuel nozzle400does not need to be inserted through the dilution hole326, but may be inserted through a separate opening on the combustor outer liner320B.

The combustor liner320has an inner surface328(a surface facing the combustor300) that surrounds the dilution hole326. In this embodiment, the secondary fuel nozzle400projects into the combustor300in a direction that is generally normal to the inner surface328of the combustor liner320. The term “generally normal” accounts for potential curvature in the inner surface328of the combustor liner320and contemplates small deviations from normal including, for example, five degrees. The secondary fuel nozzle400includes a central axis402, and the orientation of the secondary fuel nozzle400may be taken with reference to the central axis402. With the secondary fuel nozzle400projecting in a generally normal direction, the secondary fuel nozzle400projects in a direction to deliver the secondary portion of the fuel and diluent to the secondary combustion zone334, and more specifically in this embodiment to the core region336. The secondary fuel nozzle400thus preferably projects a sufficient distance, such as from zero to four times the diameter of the secondary fuel nozzle400, to deliver the secondary portion of the fuel and the diluent to the core region336.

The secondary fuel nozzle400may, however, have other orientations. The secondary fuel nozzle400may project into the combustion chamber330in a direction that forms an oblique angle with the inner surface328of the combustor liner320. For example, the secondary fuel nozzle400may project in a direction toward the primary combustion zone332, as shown inFIG.8. In another example, the secondary fuel nozzle400may have a cold side projection and project in a direction with the bulk airflow direction towards the outlet324of the combustion chamber330, as shown inFIG.9.

As shown inFIGS.6and7, the secondary fuel nozzle400may have a plurality of passages, such as a first passage410and a second passage420, each configured to convey one of the secondary portion of the fuel and the secondary portion of the diluent. In this embodiment, the first passage410is fluidly connected to the secondary diluent manifold258and is configured to convey diluent, and the second passage420is fluidly connected to the secondary fuel manifold254and is configured to convey fuel. The opposite configuration where the first passage410conveys the fuel and the second passage420conveys the diluent may also be a suitable configuration for the secondary fuel nozzle400. The plurality of passages may have any suitable geometry and configuration within the secondary fuel nozzle400. In this embodiment, the first passage410is cylindrical with the axis of the first passage410coincident with the central axis402of the secondary fuel nozzle400. The second passage420of this embodiment has an annular geometry and surrounds the first passage410in the radial direction of the secondary fuel nozzle400, which is also the radial direction of the first passage410. The axis of the second passage420is coincident with the central axis402of the secondary fuel nozzle400.

The secondary fuel nozzle400has a tip404and each of the first passage410and the second passage420extend to the tip404. The first passage410includes a plurality of orifices412configured to discharge the diluent from the first passage410(and, thus, the secondary fuel nozzle400) into the combustion chamber330. The plurality of orifices412of the first passage410may discharge the diluent in a direction parallel to the central axis402of the secondary fuel nozzle400, but they may also discharge the diluent radially outward from the central axis402. Likewise, the second passage420includes a plurality of orifices422configured to discharge the fuel from the second passage420(and, thus, the secondary fuel nozzle400) into the combustion chamber330. The plurality of orifices422of the second passage420may discharge the fuel in a direction parallel to the central axis402of the secondary fuel nozzle400, but they may also discharge the fuel radially outward from the central axis402.

With the secondary fuel nozzle400extending into the combustion chamber330the plurality of orifices412of the first passage410are thus configured to target the core region336, which may be a hot spot of the combustion chamber330. The secondary fuel nozzle400may also be configured to target other hot spots. One such hot spot is the wake or region behind (downstream in the bulk airflow direction B) of the secondary fuel nozzle400. The secondary fuel nozzle400is configured to inject diluent into the combustion chamber330towards a position behind the secondary fuel nozzle400in the bulk airflow direction B. The secondary fuel nozzle400may have at least one downstream orifice414on a downstream surface406of the secondary fuel nozzle400, as shown inFIG.6. The downstream orifice414is fluidly connected to the first passage410by a channel416. In this embodiment, two downstream orifices414are provided on the secondary fuel nozzle400with one downstream orifice414above the other downstream orifice414in the axial direction of the secondary fuel nozzle400. Each downstream orifice414is configured to direct (inject) diluent into the region behind (downstream in the bulk airflow direction B) the secondary fuel nozzle400and target the wake of the secondary fuel nozzle400and the dilution jet formed by dilution air304flowing around fuel nozzle. In this embodiment, the downstream orifices414direct diluent in the bulk airflow direction and in a direction generally parallel to the inner surface328of the combustor outer liner320B. Other configurations, such as multiple orifices414clustered around the location behind secondary fuel nozzle400or dilution hole326to reduce high temperature regions and hence NOx emission may be used. In addition, various different shapes of the orifices414may be used to effectively spread the secondary portion of the diluent in a region of high temperature.

Another hot spot may be the portions of the combustion chamber330between the dilution holes326. The secondary fuel nozzle400may configured to inject a diluent towards this hot spot and into the combustion chamber300in a lateral direction from the secondary fuel nozzle400. Each secondary fuel nozzle400may also include a lateral orifice418on at least one of the lateral side surfaces408(sides of the secondary fuel nozzle400in the lateral direction), as shown inFIG.7. The lateral orifice418is fluidly connected to the first passage410by a channel416. In this embodiment, two lateral orifices418are provided on each lateral side surface408of the secondary fuel nozzle400with one lateral orifice418above the other lateral orifice418in the axial direction of the secondary fuel nozzle400. The lateral orifices418are configured to direct (inject) diluent into the region between the dilution holes326. The lateral orifices418are thus configured to direct diluent laterally. The lateral orifices418may be configured to direct diluent in a lateral direction that is generally parallel to the inner surface328of the combustor outer liner320B.

Further aspects of the present disclosure are provided by the subject matter of the following clauses.

1. A gas turbine engine comprising: (A) a compressor section including a plurality of compressor fan blades configured to compress air flowing therethrough and to provide the air as compressed air; (B) a combustor for combusting a fuel, the combustor configured to receive the compressed air from the compressor section, the combustor including: (a) a combustor liner having (i) a combustion chamber formed therein, (ii) a forward end, and (iii) an outlet, the combustor liner being configured to have the compressed air flow therethrough in a bulk airflow direction from the forward end of the combustion chamber to the outlet of the combustion chamber, the combustion chamber having a primary combustion zone and a secondary combustion zone located downstream of the primary combustion zone in the bulk airflow direction; (b) at least one primary fuel nozzle at the forward end of the combustor liner, the at least one primary fuel nozzle being configured to inject a primary portion of the fuel into the primary combustion zone; and (c) at least one secondary fuel nozzle downstream of the at least one primary fuel nozzle in the bulk airflow direction, the at least one secondary fuel nozzle being configured to inject a secondary portion of the fuel into the secondary combustion zone, wherein the combustor is configured to mix the compressed air with the primary portion of the fuel and the secondary portion of the fuel to form a fuel and air mixture, to combust the fuel and air mixture forming combustion products, and to discharge the combustion products through the outlet of the combustion chamber; and (C) a turbine configured to receive the combustion products and be driven by the combustion products, wherein the turbine is configured to rotate the plurality of compressor fan blades of the compressor section.

2. The gas turbine engine of any preceding clause, further comprising a plurality of primary fuel nozzles, the plurality of primary fuel nozzles being configured to inject the primary portion of the fuel into the primary combustion zone; and a plurality of secondary fuel nozzles, the plurality of secondary fuel nozzles being configured to inject a secondary portion of the fuel into the secondary combustion zone.

3. The gas turbine engine of any preceding clause, wherein the primary portion of the fuel is from thirty percent to eighty percent of the fuel being injected into the combustion chamber over a time interval, and the secondary portion of the fuel is the remaining portion of the fuel injected into the combustion chamber over the time interval.

4. The gas turbine engine of any preceding clause, further comprising a plurality of diluent nozzles downstream of the plurality of primary fuel nozzles in the bulk airflow direction, the plurality of diluent nozzles being configured to inject a diluent into the combustion chamber.

5. The gas turbine engine of any preceding clause, wherein the plurality of secondary fuel nozzles and the plurality of diluent nozzles are aligned in a direction transverse to the bulk airflow direction.

6. The gas turbine engine of any preceding clause, wherein at least a portion of the plurality of secondary fuel nozzles are further configured to inject a diluent into the combustion chamber.

7. The gas turbine engine of any preceding clause, wherein the combustor includes a segment having the at least one primary fuel nozzle and the at least one secondary fuel nozzle, wherein the primary portion of the fuel is from thirty percent to eighty percent of the fuel being injected into the segment of the combustion chamber over a time interval, and the secondary portion of the fuel is the remaining portion of the fuel injected into the segment of the combustion chamber over the time interval.

8. The gas turbine engine of any preceding clause, wherein the at least one secondary fuel nozzle is configured to inject the secondary portion of the fuel into the combustion chamber in a transverse direction, the transverse direction being a direction transverse to the bulk airflow direction.

9. The gas turbine engine of any preceding clause, wherein the at least one secondary fuel nozzle projects into the combustion chamber from the combustor liner.

10. The gas turbine engine of any preceding clause, wherein the combustion liner has an inner surface surrounding the at least one secondary fuel nozzle, and wherein the at least one secondary fuel nozzle projects into the combustion chamber in a direction that forms an oblique angle with the inner surface of the combustion liner.

11. The gas turbine engine of any preceding clause, wherein the at least one secondary fuel nozzle projects into the combustion chamber in a direction toward the primary combustion zone.

12. The gas turbine engine of any preceding clause, wherein the combustion liner has an inner surface surrounding the at least one secondary fuel nozzle, and wherein the at least one secondary fuel nozzle projects into the combustion chamber in a direction generally normal to the inner surface of the combustion liner.

13. The gas turbine engine of any preceding clause, wherein the at least one secondary fuel nozzle is further configured to inject a diluent into the combustion chamber towards a position behind the at least one secondary fuel nozzle in the bulk airflow direction.

14. The gas turbine engine of any preceding clause, wherein the combustion liner has an inner surface surrounding the at least one secondary fuel nozzle, and wherein the at least one secondary fuel nozzle is further configured to inject a diluent into the combustion chamber in a lateral direction from the at least one secondary fuel nozzle.

15. The gas turbine engine of any preceding clause, wherein the lateral direction is a direction generally parallel to the inner surface of the combustion liner.

16. The gas turbine engine of any preceding clause, wherein the at least one secondary fuel nozzle is further configured to inject a diluent into the combustion chamber.

17. The gas turbine engine of any preceding clause, wherein the at least one secondary fuel nozzle includes (i) a first passage configured to convey one of the diluent and the secondary portion of the fuel and (ii) a second passage configured to convey the other one of the diluent and the secondary portion of the fuel.

18. The gas turbine engine of any preceding clause, wherein the first passage is cylindrical having an axis and a radial direction relative to the axis of the first passage, and wherein the second passage is annular, the second passage surrounding the first passage in the radial direction.

19. The gas turbine engine of any preceding clause, wherein the combustor includes a segment having the at least one primary fuel nozzle and the at least one secondary fuel nozzle, wherein the at least one primary fuel nozzle is further configured to inject a primary portion of the diluent into the combustion chamber, and the at least one secondary fuel nozzle is configured to inject a secondary portion of the diluent into the combustion chamber, the primary portion of the diluent being from twenty percent to eighty percent of the diluent injected into the segment of the combustion chamber over a time interval, and the secondary portion of the diluent is the remaining portion of the fuel injected into the segment of the combustion chamber over the time interval.

20. The gas turbine engine of any preceding clause, further comprising: (D) a fuel system including: (a) a fuel tank configured to hold the fuel; (b) a fuel delivery assembly fluidly connecting the fuel tank to the at least one primary fuel nozzle and the at least one secondary fuel nozzle; and (c) a fuel metering valve in communication with the fuel delivery assembly and configured to provide a desired volume of fuel to the at least one primary fuel nozzle and the at least one secondary fuel nozzle; and (E) a diluent system including: (a) a diluent tank configured to hold the diluent; (b) a diluent delivery assembly fluidly connecting the diluent tank to the at least one primary fuel nozzle and the at least one secondary fuel nozzle; and (c) a diluent metering valve in communication with the diluent delivery assembly and configured to provide a desired volume of diluent to the at least one primary fuel nozzle and the at least one secondary fuel nozzle.

21. The gas turbine engine of any preceding clause, wherein the fuel is hydrogen fuel, wherein the fuel tank is configured to hold the hydrogen fuel in a liquid phase, and wherein the fuel system further includes (d) a vaporizer in communication with the fuel delivery assembly for heating the hydrogen fuel in the liquid phase to at least one of a gaseous phase and a supercritical phase, the vaporizer being located between the fuel tank and the combustor.

22. The gas turbine engine of any preceding clause, wherein the fuel system further includes (d) at least one fuel proportioning valve configured to adjust the amount of fuel supplied to the at least one primary fuel nozzle and the amount of fuel supplied to the at least one secondary fuel nozzle, and wherein the diluent system further includes (d) at least one diluent proportioning valve configured to adjust the amount of diluent supplied to the at least one primary fuel nozzle and the amount of diluent supplied to the at least one secondary fuel nozzle.

23. The gas turbine engine of any preceding clause, wherein the diluent is at least one of water, nitrogen, and carbon dioxide.

24. The gas turbine engine of any preceding clause, wherein the combustor liner further includes a at least one dilution hole, the at least one dilution hole being configured to introduce dilution air into the combustion chamber.

25. The gas turbine engine of any preceding clause, wherein the at least one secondary fuel nozzle projects through the at least one dilution hole into the combustion chamber.

26. The gas turbine engine of any preceding clause, wherein the at least one secondary fuel nozzle has a cross section, the cross section being one of a circle and a teardrop shape.

27. The gas turbine engine of any preceding clause, wherein the fuel is one of diatomic hydrogen fuel and a hydrogen enriched fuel.

28. A gas turbine engine comprising: (A) a compressor section including a plurality of compressor fan blades configured to compress air flowing therethrough and to provide the air as compressed air; (B) an annular combustor for combusting a fuel, the combustor having a circumferential direction and configured to receive the compressed air from the compressor section, the combustor including: (a) a combustor liner having (i) a combustion chamber formed therein, (ii) a forward end, and (iii) an outlet, the combustor liner being configured to have the compressed air flow therethrough in a bulk airflow direction from the forward end of the combustion chamber to the outlet of the combustion chamber, the combustion chamber having a primary combustion zone and a secondary combustion zone located downstream of the primary combustion zone in the bulk airflow direction; (b) a plurality of primary fuel nozzles at the forward end of the combustor liner and aligned in the circumferential direction of the combustor, the plurality of primary fuel nozzles being configured to inject a primary portion of the fuel into the primary combustion zone and a primary portion of a diluent into the primary combustion zone; (c) a plurality of secondary fuel nozzles downstream of the primary fuel nozzle in the bulk airflow direction and aligned in the circumferential direction of the combustor, the plurality of secondary fuel nozzles being configured to inject a secondary portion of the fuel into the secondary combustion zone and a secondary portion of the diluent into the secondary combustion zone, wherein the primary portion of the fuel is from thirty percent to eighty percent of the fuel being injected into the combustion chamber over a time interval, and the secondary portion of the fuel is the remaining portion of the fuel injected into the combustion chamber over the time interval, wherein the primary portion of the diluent is from twenty percent to eighty percent of the diluent injected the combustion chamber over a time interval, and the secondary portion of the diluent is the remaining portion of the fuel injected into the combustion chamber over the time interval, and wherein the combustor is configured to mix the compressed air with the primary portion of the fuel and the secondary portion of the fuel to form a fuel and air mixture, to combust the fuel and air mixture forming combustion products, and to discharge the combustion products through the outlet of the combustion chamber; and (C) a turbine configured to receive the combustion products and be driven by the combustion products, wherein the turbine is configured to rotate the plurality of compressor fan blades of the compressor section.

Although the foregoing description is directed to the preferred embodiments, it is noted that other variations and modifications will be apparent to those skilled in the art, and may be made without departing from the spirit or scope of the disclosure Moreover, features described in connection with one embodiment may be used in conjunction with other embodiments, even if not explicitly stated above.