Circumferentially and axially staged annular combustor for gas turbine engine

A combustor section for a gas turbine engine includes an outer wall assembly and an inner wall assembly inboard of the outer wall assembly to define an annular combustion chamber therebetween. A forward fuel injection system is in communication with the combustion chamber. A downstream fuel injection system is in communication with the combustion chamber through the outer wall assembly and a swirl mixer system in communication with the combustion chamber through the inner wall assembly.

BACKGROUND

The present disclosure relates to a gas turbine engine and, more particularly, to a combustor section therefor.

Gas turbine engines such as those that power modern commercial and military aircraft, generally include a compressor section to pressurize an airflow, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases.

The combustor section generally includes radially spaced apart inner and outer liners that define an annular combustion chamber therebetween. The inner and outer liners each typically includes an outer support shell lined with heat shields, often referred to as floatwall liner panels, to line an annular combustion chamber.

Combustion of the hydrocarbon fuel in the presence of pressurized air may produce nitrogen oxide (NOX) emissions that are subjected to excessively stringent controls by regulatory authorities, and thus may be sought to be minimized. NOx formation is not only a function of temperature, but also of flame residence time and Oxygen concentration in the reaction zone. Increasing the flame strain tends to reduce the residence time in the flame, but may also increase the Oxygen concentration in the flame. These intermediate effects of strain rates tend to increase the production rate of NOx. At high strain rates, however, the reduction in flame temperature overcomes the influence of the Oxygen concentration, and NOx production rates are reduced.

SUMMARY

A combustor section for a gas turbine engine, according to one disclosed non-limiting embodiment of the present disclosure, includes an outer wall assembly and an inner wall assembly inboard of the outer wall assembly to define an annular combustion chamber therebetween. A forward fuel injection system is in communication with the annular combustion chamber. A downstream fuel injection system is in communication with the annular combustion chamber through the outer wall assembly. A swirl mixer system is in communication with the annular combustion chamber through the inner wall assembly.

In a further embodiment of the present disclosure, the downstream fuel injection system includes a multiple of first downstream fuel nozzles that alternate with a multiple of second downstream fuel nozzles.

In a further embodiment of the present disclosure, the multiple of first downstream fuel nozzles and the multiple of second downstream fuel nozzles are fueled in pairs.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the multiple of first downstream fuel nozzles are fueled through the multiple of second downstream fuel nozzles such that the multiple of first downstream fuel nozzles are each downstream to a respective one of the multiple of second downstream fuel nozzles.

In a further embodiment of any of the foregoing embodiments of the present disclosure, a valve is included in each of the multiple of second downstream fuel nozzles which selectively communicate fuel to a respective one of the multiple of first downstream fuel nozzles.

In a further embodiment of any of the foregoing embodiments of the present disclosure, each of the multiple of first downstream fuel nozzles is operable to generate a first main zone within the annular combustion chamber. Each of the multiple of second downstream fuel nozzles is operable to generate a second main zone within the annular combustion chamber.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the swirl mixer system includes a multiple of swirlers. Each of the swirlers defines a quench zone directed circumferentially between each first main zone and each associated second main zone.

In a further embodiment of any of the foregoing embodiments of the present disclosure, each quench zone overlaps with a respectively adjacent quench zone to define a shear region.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the forward fuel injection system includes a multiple of forward fuel injectors. Each of the forward fuel injectors operable to generate a pilot zone.

In a further embodiment of any of the foregoing embodiments of the present disclosure, each of the pilot zones are circumferentially located between each first main zone and each associated second main zone.

In a further embodiment of any of the foregoing embodiments of the present disclosure, sixteen (16) pilot zones; sixteen (16) quench zones; eight first main zones; and eight (8) second main zones are included.

In a further embodiment of any of the foregoing embodiments of the present disclosure, each of the pilot zones is circumferentially located in line with each first main zone.

In a further embodiment of any of the foregoing embodiments of the present disclosure eight (8) pilot zones; sixteen (16) quench zones; eight first main zones; and eight (8) second main zones are included.

A method of communicating fuel to a combustor section of a gas turbine engine, according to another disclosed non-limiting embodiment of the present disclosure, includes communicating pilot fuel axially into an annular combustion chamber; communicating fuel radially inboard into the annular combustion chamber; and communicating a multiple of quench zones radially outboard into the annular combustion chamber.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the method includes circumferentially varying the fuel communicating radially inboard into the combustion chamber to control combustion dynamics.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the method includes selectively communicating the fuel radially inboard into the combustion chamber as circumferentially alternating first and second main zones.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the method includes selectively activating the second main zones.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the method includes selectively communicating the fuel radially inboard into the annular combustion chamber through a multiple of first downstream fuel nozzles each operable to generate a first main zone within the annular combustion chamber. A multiple of second downstream fuel nozzles each operable to generate a second main zone within the annular combustion chamber.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the multiple of first downstream fuel nozzles are each downstream to a respective one of the multiple of second downstream fuel nozzles to circulate fuel through the multiple of second downstream fuel nozzles when the multiple of second downstream fuel nozzles are inactive.

In a further embodiment of any of the foregoing embodiments of the present disclosure, each quench zone overlaps with a respectively adjacent quench zone to define a shear region.

DETAILED DESCRIPTION

FIG. 1schematically illustrates a gas turbine engine20. The gas turbine engine20is disclosed herein as a two-spool turbo fan that generally incorporates a fan section22, a compressor section24, a combustor section26and a turbine section28. Alternative engine architectures200might include an augmentor section12, an exhaust duct section14and a nozzle section16in addition to the fan section22′, compressor section24′, combustor section26′ and turbine section28′ (FIG. 2) among other systems or features. The fan section22drives air along a bypass flowpath and into the compressor section24to drive core air along a core flowpath. The core air is compressed then communicated into the combustor section26for downstream expansion through the turbine section28. Although depicted as a turbofan in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines such as a turbojets, turboshafts, and three-spool (plus fan) turbofans with an intermediate spool.

The engine20generally includes a low spool30and a high spool32mounted for rotation about an engine central longitudinal axis A relative to an engine static structure36via several bearing structures38. The low spool30generally includes an inner shaft40that interconnects a fan42, a low pressure compressor (“LPC”)44and a low pressure turbine (“LPT”)46. The inner shaft40may drive the fan42directly or through a geared architecture48(seeFIG. 1) to drive the fan42at a lower speed than the low spool30. An example reduction transmission is an epicyclic transmission, namely a planetary or star gear system.

The high spool32includes an outer shaft50that interconnects a high pressure compressor (“HPC”)52and high pressure turbine (“HPT”)54. A combustor56is arranged between the HPC52and the HPT54. The inner shaft40and the outer shaft50are concentric and rotate about the engine central longitudinal axis A which is collinear with their longitudinal axes.

Core airflow is compressed by the LPC44then the HPC52, mixed with the fuel and burned in the combustor56, then expanded over the HPT54and the LPT46. The LPT46and HPT54rotationally drive the respective low spool30and high spool32in response to the expansion.

With reference toFIG. 3, the combustor section26generally includes a combustor56with an outer combustor wall assembly60, an inner combustor wall assembly62and a diffuser case module64therearound. The outer combustor wall assembly60and the inner combustor wall assembly62are spaced apart such that an annular combustion chamber66is defined therebetween.

The outer combustor wall assembly60is spaced radially inward from an outer diffuser case64A of the diffuser case module64to define an outer annular plenum76. The inner combustor wall assembly62is spaced radially outward from an inner diffuser case64B of the diffuser case module64to define an inner annular plenum78. It should be understood that although a particular combustor is illustrated, other combustor types with various combustor liner arrangements will also benefit herefrom. It should be further understood that the disclosed cooling flow paths are but an illustrated embodiment and should not be limited only thereto.

The combustor wall assemblies60,62contain the combustion products for direction toward the turbine section28. Each combustor wall assembly60,62generally includes a respective support shell68,70which supports one or more liner panels72,74mounted thereto. Each of the liner panels72,74may be generally rectilinear and manufactured of, for example, a nickel based super alloy, ceramic or other temperature resistant material and are arranged to form a liner array. In the liner array, a multiple of forward liner panels72A and a multiple of aft liner panels72B are circumferentially staggered to line the outer shell68. A multiple of forward liner panels74A and a multiple of aft liner panels74B are circumferentially staggered to also line the inner shell70.

The combustor56further includes a forward assembly80immediately downstream of the compressor section24to receive compressed airflow therefrom. The forward assembly80generally includes an annular hood82and a bulkhead assembly84that supports a multiple of swirlers90(one shown). Each of the swirlers90is circumferentially aligned with one of a multiple of fuel nozzles86(one shown) and a respective hood port94. The multiple of fuel nozzles86, swirlers90and associated fuel communication structure defines a forward fuel injection system95that supports combustion in the combustion chamber66.

The bulkhead assembly84includes a bulkhead support shell96secured to the combustor walls60,62, and a multiple of circumferentially distributed bulkhead liner panels98secured to the bulkhead support shell96around each respective swirler opening92. The bulkhead support shell96is generally annular and the multiple of circumferentially distributed bulkhead liner panels98are segmented, typically one to each fuel nozzle86and swirler90.

The annular hood82extends radially between, and is secured to, the forwardmost ends of the combustor wall assemblies60,62. The multiple of circumferentially distributed hood ports94facilitate the direction of compressed air into the forward end of the combustion chamber66through each respective swirler90. Each fuel nozzle86may be secured to the diffuser case module64to project into the respective swirler90along an axis F.

The forward assembly80introduces core combustion air into the forward section of the combustion chamber66while the remainder enters the outer annular plenum76and the inner annular plenum78. The multiple of fuel nozzles86and adjacent structure generate a blended fuel-air mixture that supports stable combustion in the combustion chamber66.

Opposite the forward assembly80, the outer and inner support shells68,70are mounted adjacent to a first row of Nozzle Guide Vanes (NGVs)54A in the HPT54. The NGVs54A are static engine components which direct core airflow combustion gases onto the turbine blades of the first turbine rotor in the turbine section28to facilitate the conversion of pressure energy into kinetic energy. The core airflow combustion gases are also accelerated by the NGVs54A because of their convergent shape and are typically given a “spin” or a “swirl” in the direction of turbine rotor rotation. The turbine rotor blades absorb this energy to drive the turbine rotor at high speed.

Each combustor wall assembly60,62(only the inner wall assembly62shown inFIG. 4) includes a multiple of studs100that extend from the liner panels72,74so as to permit the liner panels72,74to be mounted to their respective support shells68,70with fasteners102such as nuts. The studs100project rigidly from the liner panels72,74through the respective support shells68,70to receive the fasteners102at a threaded distal end section thereof to define one or more impingement cavities106. The liner panels72,74typically include one or more rails114(shown partially) that extend from a cold side110thereof. The rail114extends around the periphery of the cold side110(also shown inFIG. 5) to interface with their respective support shells68,70when mounted thereto to define the one or more impingement cavities106; however, other internal rails may alternatively or additionally be provided to define additional compartments.

A multiple of cooling impingement passages104penetrate through the support shells68,70to allow air from the respective annular plenums76,78to enter impingement cavities106formed within the combustor wall assemblies60,62between the respective support shells68,70and liner panels72,74. The cooling impingement passages104are generally normal to the surface of the liner panels72,74. The air in the cavities106provides cold side impingement cooling of the respective liner panels72,74that is generally defined herein as heat removal via internal convection.

A multiple of effusion passages108penetrate through each of the liner panels72,74. The geometry of the passages (e.g., diameter, shape, density, surface angle, incidence angle, etc.) as well as the location of the passages with respect to the high temperature combustion gas flow also contributes to effusion film cooling. The combination of impingement passages104and effusion passages108may be referred to as an Impingement Film Floatwall (IFF) assembly.

The effusion passages108allow the air to pass from each cavity106defined in part by the cold side110of the liner panels72,74to a hot side112thereof and facilitate the formation of a relatively cool insulating blanket of air along the hot side112. The effusion passages108are generally more numerous than the impingement passages104to promote the development of a sheath of film cooling along the hot side112. Film cooling as defined herein is the introduction of a relatively cooler air at one or more discrete locations along a surface exposed to a high temperature environment to protect that surface in the region of the air injection as well as downstream thereof.

A swirl mixer system115generally includes a multiple of air admission passages116that penetrate through the respective support shell70and liner panels74each along a common axis D. The air admission passages116quench the hot combustion gases within the combustion chamber66by direct supply of cooling air from the inner annular plenum78.

With continued reference toFIG. 3, a downstream fuel injection system120communicates with the combustion chamber66downstream of the forward fuel injection system95. The downstream fuel injection system120includes a multiple of downstream fuel nozzles122(illustrated schematically) located around the outer wall assembly60to introduce a portion of the fuel required for desired combustion performance; e.g., emissions, operability, durability as well as to lean-out the fuel contribution provided by the multiple of axial fuel nozzles86. Each of the multiple of downstream fuel nozzles122are located along an axis R generally transverse to the axis F of the axial fuel nozzles86. The downstream fuel injection system120generally includes a radially outer fuel manifold124(illustrated schematically) that communicates fuel to the multiple of downstream fuel nozzles122. It should be appreciated that various mount arrangements may alternatively or additionally provided inside or outside the diffuser case64.

Each of the multiple of downstream fuel nozzles122direct fuel through a downstream swirler126located coaxially with a radial outer port134in the outer wall assembly60for communication of an air-fuel mixture into the combustion chamber66. Each downstream swirler126includes an inner swirler128and an outer swirler130located within a convergent passageway132that converge inward to each port134that accelerates airflow from the inner swirler128and the outer swirler130. The inner swirler128and the outer swirler130initially counter swirl the airflow from the outer plenum78which then cancel each other out within the convergent passageway132to form an essentially axial flow fuel-air mixture jet into the combustion chamber66.

The multiple of air admission passages116are located in the inner wall assembly62to generally oppose the multiple of downstream fuel nozzles122. A swirler140is disposed in each of the multiple of air admission passages116for imparting a swirl to the air directed into the combustion chamber. In legacy gas turbine combustors, this air is admitted to quench the partially-combusted combustion product gases as solid column-like jets designed to penetrate deeply into the flow of partially-combusted combustion products passing through the zone.

Each of the swirlers140includes a central hub142, an outer shroud144disposed radially outwardly of and circumscribing the central hub142, and a plurality of circumferentially distributed vanes146that extend outwardly from the central hub142to the inner wall of the outer shroud144(FIG. 5). The vanes146are disposed at a desired vane angle so as to impart a desired degree of swirl to the combustion air admitted into the combustion chamber66.

The degree of swirl imparted to the combustion air is directly proportional to the magnitude of the vane angle, a smaller vane angle imparting a lesser degree of swirl to the combustion air and a larger vane angle imparting a greater degree of swirl to the combustion air. The vanes146may be disposed within the swirler140at a vane angle as measured in degrees departure from the vertical, ranging from as small as about 10 degrees to as high as about 50 degrees, and nominally in the range from about 15 degrees to about 25 degrees, in one example.

The swirl strength desired, and therefore the vane angle selected for a given application, generally depends on the ratio of the momentum of the quench air jets admitted through the air admission passages116to the momentum of the main flow of the combustion product gases. For a given combustor design, the greater that the magnitude of this momentum ratio is, the higher the swirl strength that can be applied while still achieving sufficient penetration and mixing, and therefore the larger the vane angle desired. Conversely, the lower that the magnitude of this momentum ratio is, the smaller the swirl strength that can be applied while still achieving sufficient penetration and mixing, and therefore the smaller the vane angle desired.

The number of vanes in a given swirler will depend upon the particular application. In general, the central hub142must be sufficiently large to package and support the number vanes146that extend between the central hub142and the outer shroud144but small enough that the central hub142, in combination with the vanes146that extend therefrom, to leave at least 70%, by line of sight, of the area circumscribed by the outer shroud144open to air flow therethrough (FIG. 5). For purposes of example, but not limitation, the swirler140may have a central hub142having a diameter of about 0.1875 inches (4.7625 millimeters), an outer shroud144having an inner diameter of about 0.65 inches (16.5 millimeters), and have ten (10) vanes, each disposed at a vane angle of twenty (20) degrees.

The forward fuel injection system95generates a multiple of pilot zones P from the forward assembly80. Each of the pilot zones P from each of the fuel nozzles86provide a swirled fuel-air mixture directed into the combustion chamber66generally along the axis F. The downstream fuel injection system120generates a main zone M fuel-air mixture directed into the combustion chamber66from the outer wall assembly60generally transverse to the axis F (also shown inFIG. 6). The swirl mixer system115generates a main zone M fuel-air mixture directed into the combustion chamber66from the inner wall assembly62generally transverse to the axis F (also shown inFIG. 6).

The downstream fuel injection system120includes a multiple of first downstream fuel nozzles122A that alternate with a multiple of second downstream fuel nozzles122B that receive fuel from the radially outer fuel manifold124. The radially outer fuel manifold124communicates with the multiple of first and second downstream fuel nozzles122A,122B in pairs. In this disclosed non-limiting embodiment, a fuel stem150from the radially outer fuel manifold124communicates fuel to one of the first multiple of downstream fuel nozzle122A first through an adjacent one of the multiple of second downstream fuel nozzles122B. That is, each of the multiple of downstream fuel nozzle122A are downstream an associated one of the multiple of second downstream fuel nozzles122B with respect to fuel flow.

A valve152(illustrated schematically) is associated with each of the multiple of second downstream fuel nozzles122B such that under an example low power condition and partial power condition, the valve152is closed to direct fuel to the one of the first multiple of downstream fuel nozzle122A yet circulate fuel through the multiple of second downstream fuel nozzles122B to avoid fuel coking therein. That is, each fuel stem150feeds one of the multiple of first downstream fuel nozzles122A and thru the valve152, one of the multiple of second downstream fuel nozzles122B of each associated pair fueled by the fuel stem150.

In one disclosed non-limiting operational embodiment (FIG. 7), under a low power condition such as idle, the forward fuel injection system95receives 100% of the fuel while the first and multiple of second downstream fuel nozzles122A,122B receive 0% of the fuel. Under a partial power condition such as cruise, the forward fuel injection system95receives about 20%-40% of the fuel, the multiple of first downstream fuel nozzles122A receive the balance of about 80%-60% of the fuel and the multiple of second downstream fuel nozzles122B receive 0% of the fuel as the valve152is closed. Notably, the fuel first circulates thru the multiple of second downstream fuel nozzles122B when the valve152is closed prior to communication to the respective multiple of first downstream fuel nozzles122A in each pair. Under a high power condition such as takeoff, the forward fuel injection system95receives about 20% of the fuel, the multiple of first downstream fuel nozzles122A receive about 30%-40% of the fuel and the multiple of second downstream fuel nozzles122B receive about 30%-40% of the fuel as the valve152is open.

Advantageously, other fuel distributions may alternatively or additionally provided for these as well as other operational conditions. For example, the fuel distribution between the first and multiple of second downstream fuel nozzles122A,122B may be readily circumferentially varied to control combustion dynamics. Such control of combustion dynamics may be additionally be utilized to vary the acoustic field within the combustor56.

With reference toFIG. 6, in one disclosed non-limiting embodiment, each of the multiple of first downstream fuel nozzles122A generates a first main zone M1fuel-air mixture directed into the combustion chamber66from the outer wall assembly60generally transverse to the axis F while the multiple of second downstream fuel nozzles122B generates a second main zone M2fuel-air mixture directed into the combustion chamber66from the outer wall assembly60generally transverse to the axis F. Each of the multiple of air admission passages116of the swirl mixer system115generates a swirled quench zone Q directed opposite to and circumferentially between respective first and second main zones M1, M2. Each of the multiple of pilot zones P centered along its respective axis F are located circumferentially in line with one of the multiple of air admission passages116(FIG. 8). That is, the number of pilot zones P, the number of quench zones Q and the total number of first and second main zones M1, M2are equal. For purposes of example, but not limitation, sixteen (16) pilot zones P; sixteen (16) quench zones Q; eight (8) first main zones M1; and eight (8) second main zones M2are provided.

Each pair of swirled mix zones Q intersect in a shear region S that is circumferentially aligned with one of the respective multiple of first main zones M1. That is, each of the swirlers140swirl in a consistent direction and overlap such that the overlap therebetween defines the shear region S. Each of the swirled mix zones Q also span respective pairs of first and second main zones M1, M2. The swirled mix zones Q thereby facilitate mixing of all three stages to, for example, reduce NOx emissions. The quench swirl strength may be readily optimized through the swirler140to achieve a desired mixture of stages and penetration.

With reference toFIG. 9, in another disclosed non-limiting embodiment, each of the multiple of pilot zones P, centered along its respective axis F, are located circumferentially in line with one of the respective multiple of first main zones M1. For purposes of example, but not limitation, eight (8) pilot zones P; sixteen (16) quench zones Q; eight (8) first main zones M1; and eight (8) second main zones M2are provided which is relatively less complicated system.

The circumferentially staging and opposed swirled quench zones tailor the mixing of multiple stages, reduce cost, address coking potentialities, and facilitate control of combustion dynamics.