Axially staged rich quench lean combustion system

A combustion system and a method of combustion for a gas turbine engine includes a combustor liner defining a combustion chamber. A plurality of fuel nozzle sets extend into, and supply fuel flow to, the combustion chamber. A pilot fuel nozzle injects a first fuel spray in a tangential direction relative to the combustion liner and toward the upstream end of the combustion chamber. A main fuel nozzle injects a second fuel spray toward the exit end of the combustion chamber. At ignition conditions, a majority of the fuel flow is injected through the pilot fuel nozzles, and at high power conditions a majority of the fuel flow is injected through the main fuel nozzles. At high power conditions, a fuel rich mixture is supplied to the combustion chamber, and a row of quench jets are configured to supply air to the combustion chamber, providing rich-quench-lean combustion.

TECHNICAL FIELD

The present invention generally relates to gas turbine engines, and more particularly relates to rich burn, quick quench, lean burn axially staged combustion systems that provide reduced emissions.

BACKGROUND

A gas turbine engine may be used to power aircraft or various other types of vehicles or systems. Such engines typically include a compressor that receives and compresses incoming gas such as air; a combustor in which the compressed gas is mixed with fuel and burned to produce high-pressure, high-velocity exhaust gas; and one or more turbines that extract energy from the exhaust gas exiting the combustor. Combustion typically takes place in one combustion zone where fuel is burned at all power conditions

There is an increasing desire to reduce combustion by-product emissions, particularly oxides of nitrogen (NOx), carbon monoxide (CO), and particulates, which may form during the combustion process. Combustion is typically achieved in a combustion chamber over a range of operating conditions. As a result, combustors operate under a variety of pressures, temperatures, and mass flows. These factors change with power requirements and environmental conditions. For example, engine idle conditions are considerably different than full power conditions. Controlling the various forms of combustion by-products over the range of operating conditions is challenging.

Accordingly, it is desirable to provide improved combustion systems in gas turbine engines with effective combustion for the control of NOx and other emissions under varying conditions. It is also desirable to provide effective control in a cost-effective manner. Furthermore, other desirable features and characteristics of the present invention will become apparent from the subsequent detailed description of the invention and the appended claims, taken in conjunction with the accompanying drawings and this background of the invention.

BRIEF SUMMARY

In accordance with an exemplary embodiment, a combustion system for a gas turbine engine is provided. The combustion system includes a combustor liner defining a combustion chamber. A plurality of fuel nozzles extend into, and supply fuel flow to, the combustion chamber. A pilot fuel nozzle injects a first fuel spray in a tangential direction relative to the combustion liner and toward the upstream end of the combustion chamber. A main fuel nozzle injects a second fuel spray toward the exit end of the combustion chamber. At ignition conditions, a majority of the fuel flow is injected through the pilot fuel nozzles, and at high power conditions a majority of the fuel flow is injected through the main fuel nozzles. At high power conditions, a fuel rich mixture is supplied to the combustion chamber, and a row of quench jets are configured to supply air to the combustion chamber, providing rich-quench-lean combustion.

In accordance with another exemplary embodiment, a method of combustion in a gas turbine engine includes defining a combustion chamber by a combustor liner that extends from an upstream end to an exit end. Air flows through the combustor liner generally from the upstream end to the exit end. A fuel flow is supplied to the combustion chamber through a pilot fuel nozzle, which injects a first fuel spray in a tangential direction relative to the combustion liner and toward the upstream end of the combustion chamber. The fuel flow is also provided through a main fuel nozzle, which injects a second fuel spray that is directed toward the exit end of the combustion chamber. A shroud extends into the combustion chamber and houses the pilot fuel nozzle and the main fuel nozzle. The first fuel spray supplies substantially all of the fuel flow at idle operating conditions, and at full power conditions the second fuel spray accounts for more than fifty percent of the fuel flow.

In accordance with another exemplary embodiment, a combustion system for a gas turbine engine includes a combustor liner defining a combustion chamber extending from an upstream end to an exit end. Air flows into and through the combustor liner, generally from the upstream end to the exit end. An axial direction is defined from the upstream end to the exit end, and the combustor liner encircles the combustion chamber. A fuel circuit includes a plurality of fuel nozzle sets extending into the combustion chamber, and supplies a fuel flow to the combustion chamber. A pilot fuel nozzle in each fuel nozzle set injects a first fuel spray in a tangential direction relative to the combustion liner and toward the upstream end of the combustion chamber. A main fuel nozzle in each fuel nozzle set injects a second fuel spray toward the exit end of the combustion chamber. At ignition conditions of the gas turbine engine the pilot fuel nozzles supply more than fifty percent of the fuel flow and at high power conditions more than fifty percent of the fuel flow is injected through the main fuel nozzles. The main fuel nozzles supply a fuel rich mixture to the combustion chamber, and a row of quench jets supply air to the combustion chamber at a location downstream from the main fuel nozzles in the axial direction, providing rich-quench-lean combustion in the combustion chamber.

DETAILED DESCRIPTION

Broadly, the exemplary embodiments discussed herein provide a combustion system and methods of combustion for a gas turbine engine that achieve combustion for emission control over a range of operating conditions. The combustion system generally includes a combustor with a pilot zone that consumes fuel at low power conditions resulting in effective emission control of carbon monoxide (CO) and unburned hydrocarbons (UHC's), and a main zone where effective emission control of Oxides of Nitrogen (NOx) and particulates is achieved at high power conditions. The pilot zone of the combustor consumes fuel at low power conditions with high combustion efficiency provided through the use of a unique pilot injector arrangement. The main zone consumes fuel that is primarily supplied through main injectors, using rich burn, quick quench, lean burn (rich-quench-lean or RQL), combustion. The pilot zone provides a long residence time for consuming the fuel through complete combustion at low power conditions. The main zone provides a short combustion time to consume the fuel under conditions that minimize unwanted emissions, as further described below.

FIG. 1is a cross-sectional view of a combustor area of an engine20showing a combustion system22in accordance with an exemplary embodiment. For purposes of the current example, the engine20is a turbofan type of gas turbine engine, with a turbine24that achieves mechanical energy from combustion of air and fuel in a combustor26. The combustor26mixes admitted air with fuel and ignites the resulting mixture to generate high energy combustion gases that are then directed into the turbine24. The mechanical energy from the turbine24is used to drive a compressor28, which may be embodied in the form of a ducted fan driven by the turbine24through a shaft30. The shaft30defines an axis31around which the compressor28and the turbine24rotate. The compressor28pressurizes air for use in the combustor26and generally accelerates air through the engine20, which may contribute to thrust. Air is delivered from the compressor28to the combustor26through an air discharge32. Compressed air is discharged into the combustor's case34. Combustion is contained within a liner36which is disposed in the case34. The combustor26is an annular type, with the liner36and the case34encircling the shaft30. Accordingly, the liner36defines a single annular shaped combustion chamber40that extends around the axis31. An injector module42extends into the combustion chamber40for the introduction of air and fuel. Given the combustion chamber40is annular in shape extending around the shaft30as defined by the liner36, it will be appreciated that a number of injector modules are disposed around its circumference. In a number of embodiments, a circumferential array of injector modules42is disposed around the combustion chamber40. These injector modules42may be equally angularly spaced about the annular combustion chamber40. In the current embodiment, the number of injector modules42is sixteen, although the number will vary with the application. In general, air from the compressor28is supplied into the case34, and air and fuel are delivered into the combustion chamber40and ignited by the igniter44. In addition to the air supplied to the combustion chamber40through the injector module42, air passes from within the case34into the combustion chamber40through various openings in the liner36as further described below.

In the exemplary embodiment ofFIG. 1, the liner36includes a dome46at its upstream end48located toward the compressor28, and an exhaust opening50at its exit end52located toward the turbine24. An axial direction49is defined between the upstream end48and the exit end52, and is parallel to the axis31. Between the dome46and the exhaust opening50, the liner36includes a ring shaped outer liner wall54and a ring shaped inner liner wall56, each of which encircles the axis31. The outer liner wall54extends around the inner liner wall56to define the annular combustion chamber40. Notable is that no fuel injector is disposed through the dome46, but rather the injector module42extends through the outer liner wall54into the combustion chamber40, at a downstream location from the dome46. Compressed air enters the case34, which defines an expanding diffuser58leading to the liner36. The air flows between the case34and the liner36through an outer air plenum60and an inner air plenum62. As will be appreciated, air enters the combustion chamber from the air plenums60,62through the injector module42, through effusion cooling holes64distributed in the dome46, the outer wall54and the inner wall56, and through quench jets68as further described below. In one example, the relative contributions of admitted air include a majority of the air entering through the injector modules42, with the quench jets68providing the second largest inflow of air to the combustion chamber40, followed by the effusion cooling holes64.

Referring toFIGS. 2 and 3, the injector module42includes a pilot injector70and a main injector72. Each of the injectors70,72includes a fuel nozzle set69with a pilot fuel nozzle74and a main fuel nozzle76, and air jets78,80respectively. Accordingly, fuel and air for combustion is introduced into the combustion chamber40from the pilot injector70and/or from the main injector72. While all the fuel enters the combustion chamber from the injector module(s)42, air is introduced through the injector module(s)42, the effusion cooling holes64and the quench jets68. In the exemplary embodiment, the pilot injector70and the main injector72are surrounded by a common shroud82. The shroud82houses the fuel nozzles74,76and provides protection from flames in the combustion chamber40.

In the exemplary embodiment, the pilot injector70includes a face84through which the pilot fuel nozzles74and the air jets78open into the combustion chamber40. With additional reference toFIG. 4, which shows a sector of the annular combustion chamber40as viewed from the exit end52, the injector module42is shown extending through the liner36. The face84is oriented to face into the combustion chamber40directing the injected fuel and air in a direction88that has a radial direction component90, a tangential direction component92, and a forward direction component94. The radial direction component90is directed from the face84toward the axis31(shown inFIG. 1). The forward direction component94is directed from the face84toward the upstream end48of the liner36in the axial direction49. The tangential direction component92is directed from the face84tangentially, relative to the annular combustion chamber40and the curvature of the liner36. Accordingly, the pilot injector70injects in a compound angular direction that is a combination of radial, tangential and axial direction components. With additional reference toFIG. 5, this orientation of the pilot injector face84injects air and fuel into a pilot zone96of the combustion chamber40from the position of the pilot injector70in the outer liner wall54at a position spaced away from the upstream end48. In the current example, the pilot zone96has a relatively large volume, for example, 492 cubic inches. In low power conditions of the engine20, substantially complete combustion of the fuel delivered through the pilot fuel nozzle74occurs in the pilot zone96, aided by a longer residence time due to the large volume.

The main injector72includes a face98through which the main fuel nozzles76and the air jets80open into the combustion chamber40. The face98is oriented to face into the combustion chamber40directing the injected fuel and air in a direction100that has a radial direction component102and a rearward direction component104. The radial direction component102is directed from the face98toward the axis31(shown inFIG. 1). The rearward direction component104is directed from the face98toward the exit end52of the liner36, in the axial direction49. Accordingly, the main injector72injects in a compound angular direction that is a combination of radial and axial components. This orientation of the main injector face98injects air and fuel into a main zone106of the combustion chamber40from the position of the main injector72in the outer liner wall54. In the current example, the main zone106has a relatively small volume, for example, 281 cubic inches. This size is smaller than the pilot zone96and the small size reduces the residence time for fuel introduced through the main injector72. NOx emissions are determined in part, by residence times, and shorter residence times are preferred.

The combustion chamber has a third axially staged zone, referred to as quench zone110. The quench zone110is so named due to a number of quench jets68that admit air from the outer air plenum60and/or the inner air plenum62into the combustion chamber40to rapidly dilute the fuel to air ratio present. In this example the quench zone110has a volume of 171 cubic inches. In comparison, the quench zone110is smaller than each of the pilot zone96and the main zone106. The remainder of the fuel introduced through the main injectors72that isn't combusted in the main zone106is completely combusted in quench zone110. The small volume helps reduce residence time for the combustion.

The location of combustion within the three zones of the combustion chamber40is determined by the supply of fuel between the pilot injectors70and the main injectors72, by their orientation, and by air admission. At ignition of the engine20, fuel is supplied through the pilot injectors70with no fuel through the main injectors72. Combustion takes place in the pilot zone96. Air is admitted directly into the pilot zone96through the air jets78of the pilot injectors70, and also through the effusion cooling holes64in the dome46, the outer liner wall54and the inner liner wall56. Air flow in the combustion chamber40is generally directed from the upstream end48toward the exit end52for exit through the exhaust opening50. The igniter44is located on the upstream side of the injector module42, which is closer to the upstream end48. The pilot injector70directs fuel with the forward direction component94to ensure ignition. The pilot injector70also directs fuel with the tangential direction component92to support complete combustion in the pilot zone96. At low power conditions, such as at idle of the engine20, as noted above, a long residence time in the combustion chamber40is provided to assist in achieving complete combustion. Fuel is supplied to the pilot zone96of the combustion chamber40through the pilot injectors70. The pilot zone96is operated at near stoichiometric, with an equivalence ratio near 1.0, so that complete combustion is achieved. The equivalence ratio is the ratio of the actual fuel/air ratio to the stoichiometric fuel/air ratio. Accordingly, stoichiometric combustion has an equivalency ratio of 1.0. Stoichiometric combustion occurs when all the oxygen is consumed in the reaction, and there is no molecular oxygen in the exhaust products. The compound angular direction at which the pilot injector70injects air and fuel helps with mixing, residence time, and combustion. Accordingly, high combustion efficiency is provided with complete combustion that results in low carbon monoxide (CO) production, which is one of the goals in emission control. Little or no fuel is supplied through the main injectors72at low power conditions. In some embodiments, a minimal amount of fuel is supplied through the main injectors72at low power conditions to avoid coking.

In the exemplary embodiment, the combustor26is an annular RQL gas turbine engine combustor. In other embodiments, the combustor26may be another type of combustor. A “rich-burn” condition in the main zone106enhances the stability of the combustion reaction by producing and sustaining a high concentration of energetic hydrogen and hydrocarbon radical species. Rich burn means more fuel is introduced than a stoichiometric amount. Rich burn conditions minimize the production of NOx due to the relative low temperatures and low population of oxygen containing intermediate species. From the main zone106to the quench zone110, the combustion regime shifts rapidly from rich to lean without dwelling at a high NOx production state. This is accomplished by moving from a rich burn state where more fuel is present than a stoichiometric amount to a lean burn state where less fuel is present than a stoichiometric amount. During operation at a high power state, such as during aircraft takeoff where the engine20operates at or near full power, a portion of the pressurized air from the compressor28enters the main zone106of the combustion chamber106through the injector module42and the effusion cooling holes64. As described above, the pilot and main injectors70,72are arranged to supply fuel in different compound angular directions, which in combination include the radial direction component90, the tangential direction component92, the forward direction component94, the radial direction component102and the rearward direction component104. The admitted air mixes with a quantity of fuel sufficient to achieve the rich burn condition in the main zone106. Fuel is supplied through both the pilot injectors70and the main injectors72. At high power conditions, the pilot zone96is operated fuel lean, such as at an equivalence ratio of 0.4. The main zone is operated fuel rich, such as at an equivalence ratio of 2.0. Operating the pilot zone96fuel lean results in low NOx production, while providing pre-heating for combustion in the main zone106. Operating the main zone106under fuel rich conditions and with a low residence time due to the relatively small volume and air flow, also results in low NOx production. The rich stoichiometry of the fuel-air mixture in the main zone106produces a relatively cool, oxygen-deprived flame, which also reduces NOx formation.

Moving from the main zone106to the quench zone110, combustion conditions transition from fuel rich to fuel lean rapidly. Air is admitted through the quench jets68to make the conversion. The combustion gases from the rich burn main zone106, which include unburned fuel, enter the quench zone110. Air is admitted from the inner and outer air plenums60,62and into the quench zone110through the quench jets68in the outer and inner liner walls54,56, respectively. Flow through the quench jets68is referred to as quench air because it rapidly mixes with the combustion gases to move them from a stoichiometrically rich state to a stoichiometrically lean state. This supports further combustion and releases additional energy from the fuel. The combustion gases passing through the quench zone110are mixed rapidly to change from a fuel-rich state to a fuel-lean state. In one example, the quench jets68admit 38 percent of the compressor discharge air. Combined with the low volume of the quench zone110, the admitted air rapidly creates the fuel lean conditions. In addition, the quench jets68are configured to produce radially penetrating air jets that project into the combustion chamber40for rapid mixing within the quench zone110. Some of the air jets admitted through the quench jets68may radially penetrate the combustion chamber40more than 50% of its radial depth. Following quench, the combustion products are consumed under fuel-lean conditions, where the combustion process concludes. The exhaust gases move through the exhaust opening50and on the turbine24.

FIG. 6illustrates combustion in the combustion chamber40of the engine20at various operating conditions graphically. Fuel pressure is indicated along the vertical axis120in pounds per square inch, and fuel flow is indicated along the horizontal axis122in pounds per hour. Fuel flow is split between the pilot injectors70and the main injectors72. For example, a flow divider valve (not shown) may sequence operation to divide flow between the two injector sets. Curve124represents the pilot injectors70and curve126represents the main injectors72. At start of the engine20, fuel flow is introduced through the pilot injectors70as indicate by curve124. The igniters44will be turned on and light off occurs at point128. The introduced fuel is lit by the igniters44causing ignition, and combustion occurs in the pilot zone96. Speed of the engine20increases and fuel flow increases to a stable idle state at point130. During this phase, more than fifty percent of the fuel flow is through the pilot injectors70, and in the exemplary embodiment, the fuel flow is entirely through the pilot injectors70. As noted above, in some embodiments, a small amount of trickle flow is delivered through the main injectors72during this phase to prevent coking. The igniters44are turned off at a threshold speed of the engine20. Above the idle point130, fuel flow is initiated and/or increased through the main injectors72as shown by the curve126, while it is decreased through the pilot injectors70as shown by a downward slope of the curve124. Combustion now occurs in the main zone106, and in the quench zone110. Fuel flow through the pilot injectors70is decreased to a minimum level132sufficient to keep the flame initiated from the pilot injectors70from blowing out. The minimum level132is maintained from point134onward. For example, at full power of the engine20, more than fifty percent, and in the exemplary embodiment, approximately ninety percent of the fuel flow is through the main injectors72. Also at full power less than fifty percent of the total fuel flow is through the pilot injectors70, and in the exemplary embodiment, approximately ten percent is through the pilot injectors70.

Referring toFIG. 7, an exemplary embodiment of the injector module42is illustrated. The injector module42is coupled to the case34, extends through the liner36, and is configured to inject fuel and air into the combustion chamber40through a pilot injector70and a main injector72. The pilot and main injectors70,72are disposed in a shroud82together. The shroud82is formed from a hollow cylindrical section130with an open end132and an end134that is closed by a domed section136, which projects outwardly from the cylindrical section130. The domed section136defines the face84of the pilot injector70and the face98of the main injector72. Accordingly, the orientation of the faces84,98is determined by the shape of the domed section136. Orientation of the faces84,98through formation of the dome136provides the compound angular directions of the fuel and air leaving the pilot injectors70and the main injectors72. The open end132is movably secured in an opening140defined in the liner36. The shroud82includes an open interior142that is open to the outer air plenum60through the open end132. A fuel rail144extends through an opening146in the case34and into the shroud82, connecting therewith at the domed section136. The fuel rail144includes a pilot leg146feeding fuel to the pilot fuel nozzle74(shown inFIGS. 2-3), and a main leg148feeding fuel to the main fuel nozzle76(shown inFIGS. 2-3). The fuel nozzles74,76are supplied with fuel from a supply system150through a fuel circuit151with separate supply lines152,154for the pilot and main supply respectively, which extend through the fuel rail144. The domed section136defines the air jets78through the face84, which are distributed around the pilot leg146. Similarly, the domed section136defines the air jets80through the face98, which are distributed around the main leg148. Air delivered from the compressor28to the outer air plenum60, is delivered to the combustion chamber40through the open interior142and the air jets78,80. The fuel rail144is maintained in position relative to the case34by a mounting flange156, which closes the opening146. Mounting brackets158, and160extend from the fuel rail144to the shroud82and are connected therewith by pins164,166. A seal ring168is trapped in a groove170formed around the opening140and engages the shroud82adjacent the open end132. The seal ring168allows the shroud82to move relative to the liner36in a direction172that allows expansion and contraction as the circumference of the shroud82changes with heating and cooling at a rate different than the liner36. The seal ring168also allows the shroud82to move in a direction174relative to the liner36, allowing it to move toward and away from the case34at a rate different than the liner36.

With reference toFIG. 8, a process200of combustion in the gas turbine engine20is initiated at step202, such as when a design and construction operation is started. At step204the engine20is constructed with a combustor, such as the combustor26described above, that is defined by a combustor liner36extending from an upstream end48to an exit end52. Air flow is moved through the combustor liner36generally from the upstream end48to the exit end52. At step206air is supplied by a compressor, such as the compressor28, and is admitted into the combustion chamber40through the liner36via the pilot and main injectors70,72, the effusion cooling holes64and the quench jets68. Fuel to the combustion chamber is supplied through the pilot fuel nozzle74and the main fuel nozzle76. A shroud extends into the combustion chamber, and houses the pilot fuel nozzle74and the main fuel nozzle76. At step208, fuel is injected, such as through the pilot fuel nozzle74via a fuel spray, such as in the direction88that has the radial direction component90, the tangential direction component92and the forward direction component94as described above. At starting of the engine20, fuel flow is introduced through the pilot injectors70. The introduced fuel is lit by the igniters44causing ignition, and combustion occurs in the pilot zone96. At step210, speed of the engine20increases and fuel flow increases to an idle state. During this phase, more than fifty percent of the fuel flow is through the pilot injectors70, and in an exemplary embodiment, the fuel flow is entirely through the pilot injectors70. The introduced fuel is completely combusted at stoichiometric conditions producing low levels of CO. When additional power is needed from the engine20, fuel is injected through the main fuel nozzles76at step212. A fuel spray is directed into the main zone106in a direction100that has a radial direction component102and a rearward direction component104as described above. At step214, more than fifty percent of the fuel flow is delivered through the main fuel nozzles76at full power conditions of the engine20. In an exemplary embodiment, ninety percent of the fuel is supplied through the main fuel nozzles76into the main zone106. The main zone is operated fuel rich and some non-combusted fuel moves to the quench zone110at step216. Moving from the main zone106to the quench zone110, combustion conditions transition from fuel rich to fuel lean rapidly. Air is admitted through the quench jets68to make the conversion. Following quench, the combustion products are consumed under fuel-lean conditions, where the combustion process concludes. Combustion is complete at step218and the exhaust gases move through the exhaust opening50and on the turbine24. The RQL combustion process results in low NOx and particulate generation. The process200may include additional steps including those described above.

Accordingly, exemplary embodiments discussed herein provide improved CO, NOx and particulate emission and temperature characteristics through a pilot and main injector arrangement, axially staged combustion, and by maintaining a desired stoichiometry in an RQL operated combustor40. Exemplary embodiments may find beneficial use in many applications, including electricity generation, propulsion, fluid transmission, and others.