COMBUSTOR HAVING A PLURALITY OF OPENINGS TO CONTROL A VORTEX DRIVER JET FOR A GAS TURBINE ENGINE

Description

TECHNICAL FIELD

The present disclosure relates generally to a combustor having a plurality of openings to control a vortex driver jet for a gas turbine engine.

BACKGROUND

Turbine engines, for example, for aircraft, generally include a fan and a turbo-engine section arranged in flow communication with one another. The turbo-engine includes, in serial flow relationship, a compressor section, a combustion section, and a turbine section. The combustion section includes a combustor where the compressed airflow is mixed with fuel and ignited to generate combustion gases.

BRIEF DESCRIPTION OF THE DRAWINGS

Features and advantages will be apparent from the following, more particular, description of various exemplary embodiments, as illustrated in the accompanying drawings, wherein like reference numbers generally indicate identical, functionally similar, and/or structurally similar elements.

FIG. 1 is a schematic cross-sectional view of a turbine engine, according to an embodiment of the present disclosure.

FIG. 2 is a cross-sectional side view of a combustor of the turbine engine shown in FIG. 1, according to an embodiment of the present disclosure.

FIG. 3A is a cross-sectional side view of a combustor of the turbine engine, the combustor having a plurality of openings provided in a casing of the combustor, the openings located near an upstream end of the combustor being in an open position, according to an embodiment of the present disclosure.

FIG. 3B is a cross-sectional side view of the combustor shown in FIG. 3A, the combustor having a plurality of openings provided in a casing of the combustor, the openings located near a downstream end of the combustor being in an open position, according to another embodiment of the present disclosure.

FIG. 4A is a cross-sectional side view of a combustor of the turbine engine, the combustor having a plurality of openings provided in a casing of the combustor, the openings located near a downstream end of the combustor being in an open position for introducing airflow into the combustor, according to another embodiment of the present disclosure.

FIG. 4B is a cross-sectional side view of a combustor of the turbine engine, the combustor having a plurality of openings provided in a casing of the combustor, the openings located near a downstream end of the combustor are in communication with a sealed chamber, according to another embodiment of the present disclosure.

FIG. 5 is a cross-sectional side view of a combustor of the turbine engine, the combustor having a back pressure plate, according to another embodiment of the present disclosure.

FIGS. 6A to 6C are schematic examples of the outer annular ring of the back pressure plate shown in FIG. 5 in various configurations, according to an embodiment of the present disclosure.

FIGS. 7A to 7C are schematic examples of the outer annular ring of the back pressure plate of FIG. 5 in various configurations, according to another embodiment of the present disclosure.

DETAILED DESCRIPTION

Features, advantages, and embodiments of the present disclosure are set forth or apparent from a consideration of the following detailed description, drawings, and claims. Moreover, the following detailed description is exemplary and intended to provide further explanation without limiting the scope of the disclosure as claimed.

Various embodiments of the present disclosure 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 present disclosure.

As used herein, the terms “first” and “second” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components.

The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows, and “downstream”refers to the direction to which the fluid flows.

The terms “forward” and “aft” refer to relative positions within a turbine engine or vehicle, and refer to the normal operational attitude of the turbine engine or vehicle. For example, with regard to a turbine engine, forward refers to a position on the turbine engine that is closer to the propeller or the fan and aft refers to a position on the turbine engine that is further away from the propeller or the fan.

As used herein, the term “axial” refers to directions and orientations that extend substantially parallel to a centerline of the turbine engine. As used herein, the term “radial” refers to directions and orientations that extend substantially perpendicular to the centerline of the turbine engine. In addition, as used herein, the terms “circumferential” and “circumferentially” refer to directions and orientations that extend arcuately about the centerline of the turbine engine.

As used herein, “top” refers to a highest or uppermost point, portion, or surface of a component in the orientations shown in the figures.

As used herein, the terms “low,” “mid” (or “mid-level”), and “high,” or their respective comparative degrees (e.g., “lower” and “higher”, where applicable), when used with compressor, combustor, turbine, shaft, fan, or turbine engine components, each refers to relative pressures, relative speeds, relative temperatures, or relative power outputs within an engine unless otherwise specified. For example, a “low-power” setting defines the engine or the combustor configured to operate at a power output lower than a “high-power” setting of the engine or the combustor, and a “mid-level power” setting defines the engine or the combustor configured to operate at a power output higher than a “low-power” setting and lower than a “high-power” setting. The terms “low,” “mid” (or “mid-level”) or “high” in such aforementioned terms may additionally, or alternatively, be understood as relative to minimum allowable speeds, pressures, or temperatures, or minimum or maximum allowable speeds, pressures, or temperatures relative to normal, desired, steady state, etc., operation of the engine. A mission cycle for a turbine engine includes, for example, a low-power operation, a mid-level power operation, and a high-power operation. Low-power operation includes, for example, engine start, idle, taxiing, and approach. Mid-level power operation includes, for example, cruise. High-power operation includes, for example, takeoff and climb.

The various power levels of the turbofan engine are defined as a percentage of a sea level static (SLS) maximum engine rated thrust. Low power operation includes, for example, less than thirty percent (30%) of the SLS maximum engine rated thrust of the turbofan engine. Mid-level power operation includes, for example, thirty percent (30%) to eighty-five percent (85%) of the SLS maximum engine rated thrust of the turbofan engine. High power operation includes, for example, greater than eighty-five percent (85%) of the SLS maximum engine rated thrust of the turbofan engine. The values of the thrust for each of the low power operation, the mid-level power operation, and the high power operation of the turbofan engine are exemplary only, and other values of the thrust can be used to define the low power operation, the mid-level power operation, and the high power operation.

The terms “coupled,” “fixed,” “attached,” “connected,” and the like, refer to both direct coupling, fixing, attaching, or connecting, as well as indirect coupling, fixing, attaching, or connecting through one or more intermediate components or features, unless otherwise specified herein.

The singular forms “a,” “an,” and “the” include plural references unless the context clearly dictates otherwise.

As used herein, a “turbo-engine” includes a compressor section, a combustion section, and a turbine section.

As used herein, a “turbofan engine” includes a turbo-engine and a fan that directs air into the turbo-engine, and rated for use in a regional aircraft, a narrow body aircraft, or a wide body aircraft. A turbofan/turboprop engine rated for small aircraft applications will have a maximum takeoff thrust range between two thousand and nine thousand pound-force (2,000 lbf to 9,000 lbf). A turbofan engine rated for use on a regional aircraft will have a maximum takeoff thrust in a range of ten thousand pound-force to twenty thousand pound-force (10,000 lbf to 20,000 lbf). A turbofan engine rated for use on a narrow body aircraft will have a maximum takeoff thrust in a range of fifteen thousand pound-force to thirty thousand pound-force (15,000 lbf to 30,000 lbf). A turbofan engine rated for use on a wide body aircraft will have a maximum takeoff thrust in a range of forty thousand pound-force to one hundred ten thousand pound-force (40,000 lbf to 110,000 lbf).

As used herein, the term “ducted engine” means a turbofan engine with a fan casing or a nacelle that circumferentially surrounds the fan.

Hereafter, the term “turbofan engine” will refer to either a “ducted engine” or an “open fan engine.”

As used herein, a Mach number is a ratio of the speed of the turbofan engine (of the aircraft) to the speed of sound in the surrounding airflow.

Approximating language, as used herein throughout the specification and claims, is applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, a value modified by a term or terms, such as “about,” “approximately,” “generally,” and “substantially” is not to be limited to the precise value specified. In at least some instances, the approximating language may correspond to the precision of an instrument for measuring the value, or the precision of the methods or the machines for constructing the components and/or the systems or manufacturing the components and/or the systems. For example, the approximating language may refer to being within a one, two, four, ten, fifteen, or twenty percent margin in either individual values, range(s) of values and/or endpoints defining range(s) of values.

Providing gas turbine engines with reduced length and weight while maintaining or improving combustion efficiency is a significant challenge. Reducing length or combustion volume typically leads to lower efficiency and incomplete combustion in the combustor of the turbine engine.

The use of vortex combustors allows for reducing combustor length and volume while maintaining or improving combustion efficiency. For example, conventional trapped vortex combustors have a reduced length or volume while maintaining, or even improving, combustion efficiency compared to legacy swirl stabilized combustion systems. However, conventional trapped vortex combustors have a limitation. The trap in a trapped vortex combustor becomes less effective at small combustor sizes. Alternatively, an un-trapped or partially trapped vortex combustor can be implemented. However, maintaining a stable toroidal vortex without a full trap may be difficult to accomplish.

For a combustor having a geometrically un-trapped or a geometrically partially trapped vortex for flame stabilization, stabilization of the vortex is predominately driven by aerodynamic jets. The momentum and the penetration of these jets into the combustor is implemented for aerodynamically closing the toroidal vortex/vortices.

Toroidally stabilized vortex combustors are smaller and lighter than traditional rich or lean burn combustors. A geometrically partially trapped or un-trapped combustor is more conducive to smaller engine applications due to geometric constraints and scaling of aerodynamic structures.

Embodiments of the present disclosure improve the stability of the partially trapped or un-trapped vortex by increasing pressure drop (momentum) of the vortex driver jets in order to ensure aerodynamic closure of the vortex. However, there is only so much that can be done through the geometry of driver openings (e.g., driver holes, driver slots, or driver vanes). Embodiments of the present disclosure provide additional features to the combustor-diffuser-nozzle (CDN) to improve feed pressure and therefore increase local pressure drop and momentum of the vortex drivers. This feature can be primarily implemented by changing a position of bleed flows, locally adding flow or a synthetic jet to back pressure vortex driver jets, or adding physical blockage downstream of the jets to improve pressure drop.

The use of variation in bleed flows, flow inlets, synthetic jets, or physical blockages (static or variable) to increase or otherwise to alter momentum of vortex driver jets in vortex combustor can be used to increase local pressure drop and momentum of the driver vortex drivers.

In an embodiment, the CDN out-take bleeds in the combustor can have different axial positions in which the bleeds can be varied by location to change the local pressure at the vortex driver jets to strengthen or to weaken the vortex.

In an embodiment, the CDN can be provided with flow introduction near to generate vortex driver jets and back-pressure jets at driver openings to increase momentum for strengthening vortex stability.

In an embodiment, a perforated plate or a screen at an exit of inner combustor passages and/or outer combustor passages before stage-one-nozzle (S1N) to back-pressure passages can be provided to increase pressure supply to driver jets, increasing jet momentum for strengthening vortex stability. For example, a single static plate, or two plates that move relative to one another can be used to vary back-pressure to optimize combustor performance.

Referring now to the drawings, FIG. 1 is a schematic cross-sectional view of a turbine engine 10, taken along a longitudinal centerline axis 12 of the turbine engine 10, according to an embodiment of the present disclosure. As shown in FIG. 1, the turbine engine 10 defines an axial direction A (extending parallel to the longitudinal centerline axis 12 provided for reference) and a radial direction R that is normal to the axial direction A. In general, the turbine engine 10 includes a fan section 14 and a turbo-engine 16 disposed downstream from the fan section 14.

The turbo-engine 16 includes, in serial flow relationship, a compressor section 21, a combustion section 26, and a turbine section 27. The turbo-engine 16 is substantially enclosed within an outer casing 18 that is substantially tubular and defines a turbo-engine inlet 20 that is annular about the longitudinal centerline axis 12. As schematically shown in FIG. 1, the compressor section 21 includes a booster or a low pressure (LP) compressor 22 followed downstream by a high pressure (HP) compressor 24. The combustion section 26 is downstream of the compressor section 21. The turbine section 27 is downstream of the combustion section 26 and includes a high pressure (HP) turbine 28 followed downstream by a low pressure (LP) turbine 30. The turbo-engine 16 further includes a jet exhaust nozzle section 32 that is downstream of the turbine section 27, a high-pressure (HP) shaft 34 or a spool, and a low-pressure (LP) shaft 36. The HP shaft 34 drivingly connects the HP turbine 28 to the HP compressor 24. The HP turbine 28 and the HP compressor 24 rotate in unison through the HP shaft 34. The LP shaft 36 drivingly connects the LP turbine 30 to the LP compressor 22. The LP turbine 30 and the LP compressor 22 rotate in unison through the LP shaft 36. The compressor section 21, the combustion section 26, the turbine section 27, and the jet exhaust nozzle section 32 together define a turbo-engine air flow path.

For the embodiment depicted in FIG. 1, the fan section 14 includes a fan 38 (e.g., a variable pitch fan) having a plurality of fan blades 40 coupled to a disk 42 in a spaced apart manner. As depicted in FIG. 1, the fan blades 40 extend outwardly from the disk 42 generally along the radial direction R. In the case of a variable pitch fan, the plurality of fan blades 40 are rotatable relative to the disk 42 about a pitch axis P by virtue of the fan blades 40 being operatively coupled to an actuation member 44 configured to collectively vary the pitch of the fan blades 40 in unison. The fan blades 40, the disk 42, and the actuation member 44 are together rotatable about the longitudinal centerline axis 12 via a fan shaft 45 that is powered by the LP shaft 36 across a power gearbox, also referred to as a gearbox assembly 46. In this way, the fan 38 is drivingly coupled to, and powered by, the turbo-engine 16, and the turbine engine 10 is an indirect drive engine. The gearbox assembly 46 is shown schematically in FIG. 1. The gearbox assembly 46 is a reduction gearbox assembly for adjusting the rotational speed of the fan shaft 45 and, thus, the fan 38 relative to the LP shaft 36 when power is transferred from the LP shaft 36 to the fan shaft 45.

Referring still to the exemplary embodiment of FIG. 1, the disk 42 is covered by a fan hub 48 that is aerodynamically contoured to promote an airflow through the plurality of fan blades 40. In addition, the fan section 14 includes an annular fan casing or a nacelle 50 that circumferentially surrounds the fan 38 and at least a portion of the turbo-engine 16. The nacelle 50 is supported relative to the turbo-engine 16 by a plurality of outlet guide vanes 52 that are circumferentially spaced about the nacelle 50 and the turbo-engine 16. Moreover, a downstream section 54 of the nacelle 50 extends over an outer portion of the turbo-engine 16, and, with the outer casing 18, defines a bypass airflow passage 56 therebetween.

During operation of the turbine engine 10, a volume of air 58 enters the turbine engine 10 through an inlet 60 of the nacelle 50 or the fan section 14. As the volume of air 58 passes across the fan blades 40, a first portion of air, also referred to as bypass air 62, is routed into the bypass airflow passage 56, and a second portion of air, also referred to as turbo-engine air 64, is routed into the upstream section of the turbo-engine air flow path through the turbo-engine inlet 20 of the LP compressor 22. The pressure of the turbo-engine air 64 is then increased, generating compressed airflow 65. The compressed airflow 65 is routed through the HP compressor 24 and into the combustion section 26, where the compressed airflow 65 is mixed with fuel and ignited to generate combustion gases 66.

The combustion gases 66 are routed into the HP turbine 28 and expanded through the HP turbine 28 where a portion of thermal energy or kinetic energy from the combustion gases 66 is extracted via one or more stages of HP turbine stator vanes 68 and HP turbine rotor blades 70 that are coupled to the HP shaft 34. This causes the HP shaft 34 to rotate, thereby supporting operation of the HP compressor 24 (self-sustaining cycle). In this way, the combustion gases 66 do work on the HP turbine 28. The combustion gases 66 are then routed into the LP turbine 30 and expanded through the LP turbine 30. Here, a second portion of the thermal energy or the kinetic energy is extracted from the combustion gases 66 via one or more stages of LP turbine stator vanes 72 and LP turbine rotor blades 74 that are coupled to the LP shaft 36. This causes the LP shaft 36 to rotate, thereby supporting operation of the LP compressor 22 (self-sustaining cycle) and rotation of the fan 38 via the gearbox assembly 46. In this way, the combustion gases 66 do work on the LP turbine 30.

The combustion gases 66 are subsequently routed through the jet exhaust nozzle section 32 of the turbo-engine 16 to provide propulsive thrust. Simultaneously, the bypass air 62 is routed through the bypass airflow passage 56 before being exhausted from a fan nozzle exhaust section 76 of the turbine engine 10, also providing propulsive thrust. The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section 32 at least partially define a hot gas path 78 for routing the combustion gases 66 through the turbo-engine 16.

A controller 102 is in communication with the turbine engine 10 for controlling aspects of the turbine engine 10. For example, the controller 102 is in two-way communication with the turbine engine 10 for receiving signals from various sensors and control systems of the turbine engine 10, and for controlling components of the turbine engine 10, as detailed further below. The controller 102, or components thereof, may be located onboard the turbine engine 10, onboard the aircraft, or can be located remote from each of the turbine engine 10 and the aircraft. The controller 102 can be a Full Authority Digital Engine Control (FADEC) that controls aspects of the turbine engine 10.

The controller 102 may be a standalone controller or may be part of an engine controller to operate various systems of the turbine engine 10. In this embodiment, the controller 102 is a computing device having one or more processors and a memory. The one or more processors can be any suitable processing device, including, but not limited to, a microprocessor, a microcontroller, an integrated circuit, a logic device, a programmable logic controller (PLC), an application specific integrated circuit (ASIC), or a Field Programmable Gate Array (FPGA). The memory can include one or more computer-readable media, including, but not limited to, non-transitory computer-readable media, a computer readable non-volatile medium (e.g., a flash memory), a RAM, a ROM, hard drives, flash drives, or other memory devices.

The memory can store information accessible by the one or more processors, including computer-readable instructions that can be executed by the one or more processors. The instructions can be any set of instructions or a sequence of instructions that, when executed by the one or more processors, cause the one or more processors and the controller 102 to perform operations. The controller 102 and, more specifically, the one or more processors are programmed or configured to perform these operations, such as the operations discussed further below. In some embodiments, the instructions can be executed by the one or more processors to cause the one or more processors to complete any of the operations and functions for which the controller 102 is configured, as will be described further below. The instructions can be software written in any suitable programming language or can be implemented in hardware. Additionally, or alternatively, the instructions can be executed in logically or virtually separate threads on the processors. The memory can further store data that can be accessed by the one or more processors.

The technology discussed herein makes reference to computer-based systems and actions taken by, and information sent to and from, computer-based systems. One of ordinary skill in the art will recognize that the inherent flexibility of computer-based systems allows for a great variety of possible configurations, combinations, and divisions of tasks and functionality between and among components. For instance, processes discussed herein can be implemented using a single computing device or multiple computing devices working in combination. Databases, memory, instructions, and applications can be implemented on a single system or distributed across multiple systems. Distributed components can operate sequentially or in parallel.

The turbine engine 10 depicted in FIG. 1 is by way of example only. In other exemplary embodiments, the turbine engine 10 may have any other suitable configuration. For example, in other exemplary embodiments, the fan 38 may be configured in any other suitable manner (e.g., as a fixed pitch fan) and further may be supported using any other suitable fan frame configuration. The turbine engine 10 may also be a direct drive engine, which does not have a power gearbox. The fan speed is the same as the LP shaft speed for a direct drive engine. Moreover, in other exemplary embodiments, any other suitable number or configuration of compressors, turbines, shafts, or a combination thereof may be provided. In still other exemplary embodiments, aspects of the present disclosure may be incorporated into any other suitable turbine engine, such as, for example, turbofan engines, propfan engines, turbojet engines, turboprop, turboshaft engines, or aeroderivative ground based engines.

FIG. 2 is a cross-sectional side view of a combustor 206 in the combustion section 26 of the turbo-engine 16 shown in FIG. 1, according to an embodiment of the present disclosure. FIG. 2 depicts a longitudinal combustor centerline axis 112 that may generally correspond to the longitudinal centerline axis 12 (FIG. 1). The combustor 206, shown in FIG. 2, defines a combustor longitudinal direction (L) corresponding to the longitudinal combustor centerline axis 112, a combustor radial direction (R) extending outward from the longitudinal combustor centerline axis 112, and a combustor circumferential direction (C) extending circumferentially about the longitudinal combustor centerline axis 112. The combustor 206 extends from an upstream end 253 of the combustor 206 to a downstream end 255 of the combustor 206. The upstream end 253 of the combustor 206 may be in airflow communication with a diffuser 257 that is in airflow communication with the HP compressor 24 (shown in FIG. 1). The diffuser 257, in one aspect, may be arranged between an upstream end 263 of an outer casing 264 of the combustor 206 and an upstream end 265 of an inner casing 266 of the combustor 206. The downstream end 255 of the combustor 206 is in airflow communication with a turbine nozzle 272 that is in airflow communication with the HP turbine 28 (shown in FIG. 1). The outer casing 264 may extend circumferentially about the longitudinal combustor centerline axis 112 and may extend longitudinally from the upstream end 253 of the combustor 206 to the downstream end 255 of the combustor 206. The inner casing 266 may also extend circumferentially about the longitudinal combustor centerline axis 112 and may extend longitudinally from the upstream end 253 of the combustor 206 to the downstream end 255 of the combustor 206. The outer casing 264 is located at a first radial distance R1 from the longitudinal combustor centerline axis 112. The inner casing 266 is located at a second radial distance R2 from the longitudinal combustor centerline axis 112. The first radial distance is greater than the second radial distance.

The combustor 206 further includes a dome structure 256 that may include an inner dome portion 258 and an outer dome portion 260. Together, the inner dome portion 258 and the outer dome portion 260 define a turn portion 262 that may be generally a convex-shaped curve that forms a flow path for directing the flow of combustion products within the combustor 206. In addition, the convex-shaped curve of the dome structure 256 may assist in guiding a portion of compressed airflow 282A to flow to an outer flow passage 288, and assist in guiding another portion of compressed airflow 282B to flow to an inner flow passage 290. The convex-shaped curve of the dome structure 256 may therefore help to reduce pressure losses in the flow of the compressed airflow 282 that may otherwise occur when a more blunt-shaped dome structure is implemented.

An outer liner 252 may extend downstream from an outer end 261 of the dome structure 256, and an inner liner 254 may extend downstream from an inner end 259 of the dome structure 256. The outer liner 252 and the inner liner 254 may be formed integral with the dome structure 256 (e.g., formed of a continuous structural material, such as a formed ceramic matrix composite (CMC) structure), or may be joined to the dome structure 256 via, for example, a mechanical connection (e.g., a bolted connection), or bonded via, for example, brazing. The inner liner 254, the outer liner 252, and the dome structure 256 form a liner 269 and define a combustion chamber 267. The inner liner 254 and the outer liner 252 extend longitudinally along the longitudinal combustor centerline axis 112 and circumferentially around the longitudinal combustor centerline axis 112. The inner liner 254 and the outer liner 252 are radially spaced apart from each other to define therebetween the combustion chamber 267. The outer liner 252 is spaced apart from the outer casing 264 to define therebetween the outer flow passage 288. The inner liner 254 is spaced apart from the inner casing 266 to define therebetween the inner flow passage 290.

As used herein, CMC refers to a class of materials with reinforcing fibers in a ceramic matrix. Generally, the reinforcing fibers provide structural integrity to the ceramic matrix. Some examples of reinforcing fibers can include, but are not limited to, non-oxide silicon-based materials (e.g., silicon carbide, silicon nitride, or mixtures thereof), non-oxide carbon-based materials (e.g., carbon), oxide ceramics (e.g., silicon oxycarbides, silicon oxynitrides, aluminum oxide (Al₂O₃), silicon dioxide (SiO₂), aluminosilicates such as mullite, or mixtures thereof), or mixtures thereof.

Some examples of ceramic matrix materials can include, but are not limited to, non-oxide silicon-based materials (e.g., silicon carbide, silicon nitride, or mixtures thereof), oxide ceramics (e.g., silicon oxycarbides, silicon oxynitrides, aluminum oxide (Al₂O₃), silicon dioxide (SiO₂), aluminosilicates, or mixtures thereof), or mixtures thereof. Optionally, ceramic particles (e.g., oxides of Si, Al, Zr, Y, and combinations thereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica, talc, kyanite, and montmorillonite) can also be included within the ceramic matrix.

Generally, particular CMCs can be referred to as their combination of type of fiber/type of matrix. For example, C/SiC for carbon-fiber-reinforced silicon carbide, SiC/SiC for silicon carbide-fiber-reinforced silicon carbide, SiC/SiN for silicon carbide fiber-reinforced silicon nitride, SiC/SiC—SiN for silicon carbide fiber-reinforced silicon carbide/silicon nitride matrix mixture, etc. In other examples, the CMCs can be comprised of a matrix and reinforcing fibers comprising oxide-based materials such as aluminum oxide (Al₂O₃), silicon dioxide (SiO₂), aluminosilicates, and mixtures thereof. Aluminosilicates can include crystalline materials such as mullite (3Al₂O₃·2SiO₂), as well as glassy aluminosilicates.

In certain non-limiting examples, the reinforcing fibers may be bundled (e.g., form fiber tows) and/or coated prior to inclusion within the matrix. The bundles of fibers may be impregnated with a slurry composition prior to forming the preform or after formation of the preform. The preform may then undergo thermal processing, and subsequent chemical processing to arrive at a component formed of a CMC material having a desired chemical composition. For example, the preform may undergo a cure or a burn-out to yield a high char residue in the preform, and subsequent melt-infiltration with silicon, or a cure or a pyrolysis to yield a silicon carbide matrix in the preform, and subsequent chemical vapor infiltration with silicon carbide. Additional steps may be taken to improve densification of the preform, either before or after chemical vapor infiltration, by injecting the preform with a liquid resin or a polymer followed by a thermal processing step to fill the voids with the silicon carbide. A CMC material as used herein may be formed using any known or hereafter developed methods including, but not limited to, melt infiltration, chemical vapor infiltration, polymer impregnation pyrolysis (PIP), or any combination thereof.

A fuel nozzle assembly 274 may be connected to the outer casing 264, and may extend through the outer casing 264 and through the outer liner 252 to provide a flow of fuel 275 to the combustion chamber 267. The compressed airflow 282 enters the combustor 206 via the diffuser 257 as an inlet airflow. A portion of the compressed airflow 282, shown schematically as compressed airflow 282A, flows into the outer flow passage 288, while another portion of the compressed airflow 282, shown schematically as compressed airflow 282B, flows into the inner flow passage 290.

In an embodiment, the compressed airflow 282A in the outer flow passage 288 passes through cooling openings 232 and forward driver openings 242 (e.g., holes, slots, or vanes) in the dome structure 256 and the outer liner 252 into the combustion chamber 267. The outer flow 288 also passes through aft driver openings 242A (e.g., holes, slots, or vanes) in the dome structure 256 and the outer liner 252 into the combustion chamber 267. The compressed airflow 282B in the inner flow passage 290 passes through cooling openings 234 and driver openings 244 (e.g., holes, slots, or vanes) in the dome structure 256 and the inner liner 254 into the combustion chamber 267 to generate the vortex 295 and a vortex 297 with the compressed airflow 282A and to mix with the fuel 275.

The compressed airflow 282B in the inner flow passage 290 passes through cooling openings 234 (e.g., holes, slots, or vanes) and driver openings 244 (e.g., holes, slots, or vanes) in the dome structure 256 and the inner liner 254 into the combustion chamber 267 to generate a vortex 295 with the compressed airflow 282A and to mix with the fuel 275.

In an embodiment, as will be explained further in detail in the following paragraphs, the holes, slots and/or vanes in the dome structure 256, the outer liner 252, and/or in the inner liner 254 can be configured to orient a direction of flow of the compressed airflow 282A and a direction of flow of the compressed airflow 282B so as to create the vortex 295 and/or the vortex 297.

FIG. 3A is a cross-sectional side view of a combustor 306 suitable for use in the combustion section 26 of the turbo-engine 16 shown in FIG. 1, the combustor 306 having a plurality of openings provided in a casing of the combustor 306, the openings located near an upstream end of the combustor 306 being in an open position, according to another embodiment of the present disclosure. The combustor 306 is similar in many aspects to the combustor 206 (FIG. 2). Therefore, similar features in combustor 206 and combustor 306 are not described in detail in the following paragraphs. Similar features are referred to herein with the same reference numbers. The combustor 306 includes a plurality of outer openings 308A and 308B provided in the outer casing 264 and a plurality of inner openings 310A and 310B are provided in the inner casing 266. The plurality of outer openings 308A and the plurality of inner openings 310A can be provided at the upstream end 263 of the outer casing 264 and the upstream end 265 of the inner casing 266, respectively of the combustor 306. The plurality of outer openings 308B and the plurality of inner openings 310B can be provided at the downstream end 255 of the combustor 306. The plurality of outer openings 308A and 308B in the outer casing 264 and the plurality of inner openings 310A and 310B in the inner casing 266 can be holes, slots, and/or vanes. The holes, slots, and/or vanes can have any shape, including, but not limited to, a circular shape, an oval shape, an egg shape, a polygonal shape, an “I” shape, a wavy shape, etc. In addition, although one example of each opening 308A, 308B, 310A and 310B is depicted in FIG. 3A, one or more outer openings 308A, one or more outer openings 308B, one or more inner openings 310A, and one or more inner openings 310B can be provided. For example, the plurality of outer openings 308A and a plurality of outer openings 308B can be distributed circumferentially around the outer casing 264. For example, a plurality of inner openings 310A and the plurality of inner openings 310B can be distributed circumferentially around the inner casing 266.

In an embodiment, the combustor 306 further includes a plurality of first doors 309A and 309B configured to cover the plurality of outer openings 308A and 308B, respectively. In an embodiment, the combustor 306 also includes a plurality of second doors 311A and 311B configured to cover the plurality of inner openings 310A and 310B, respectively. The plurality of first doors 309A and 309B are configured to open and to close, to control an airflow through the plurality of outer openings 308A and 308B. The plurality of second doors 311A and 311B are configured to open and to close, to control an airflow through the plurality of inner openings 310A and 310B.

The plurality of first doors 309A and 309B are mounted to the outer casing 264. The plurality of second doors 311A and 311B are mounted to the inner casing 266. The plurality of first doors 309A and 309B can be mounted to the outer casing 264 via hinges (not shown). The plurality of first doors 309A and 309B can be slidably mounted to the outer casing 264. The plurality of first doors 309A and 309B can also be iris-shutters mounted to the outer casing 264. Similarly, the plurality of second doors 311A and 311B can be mounted to the inner casing 266 via hinges (not shown). The plurality of second doors 311A and 311B can be slidably mounted to the inner casing 266. The plurality of second doors 311A and 311B also can be iris-shutters mounted to the outer casing 264. Although the doors are shown as being used to cover the various openings, other mechanisms also can be used to open and to close the openings such as, for example, by using controllable valves, or the like.

In an embodiment, the combustor 306 includes an actuator 314 configured to actuate the plurality of first doors 309A and/or 309B and to actuate the plurality of second doors 311A and/or 311B. The actuator 314 is in communication with the controller 102 (shown in FIG. 3A and FIG. 1). The controller 102 is configured to send one or more signals to open and/or to close any of the plurality of first doors 309A, 309B and/or to open and/or to close any of the plurality of second doors 311A and/or 311B.

For example, as shown in FIG. 3A, the plurality of first doors 309A and the plurality of second doors 311A that are configured to cover the plurality of outer openings 308A and the plurality of inner openings 310A, respectively, provided near the upstream end 263 of the outer casing 264 of the combustor 306 are shown in an open position. The plurality of first doors 309A and the plurality of second doors 311A may be referred to as a plurality of upstream doors. In addition, as shown in FIG. 3A, the plurality of first doors 309B and the plurality of second doors 311B that are configured to cover the plurality of outer openings 308B and the plurality of inner openings 310B, respectively, provided near the downstream end 255 of the combustor 306 are shown in a closed position. The plurality of first doors 309B and the plurality of second doors 311B may be referred to as a plurality of downstream doors.

A portion of the compressed airflow 282, shown schematically as compressed airflow 282A, flows into the outer flow passage 288, while another portion of the compressed airflow 282, shown schematically as compressed airflow 282B, flows into the inner flow passage 290. In an embodiment, the compressed airflow 282A in the outer flow passage 288 passes through the cooling openings 232 (e.g., holes, slots, or vanes) and through driver openings 242 (e.g., holes, slots, or vanes) and aft driver openings 242A in the dome structure 256 and the outer liner 252 into the combustion chamber 267 to generate vortex driver jets 243A and 243B. The compressed airflow 282B in the inner flow passage 290 passes through cooling openings 234 (e.g., holes, slots, or vanes) and the driver openings 244 (e.g., holes, slots, or vanes) in the dome structure 256 and the inner liner 254 into the combustion chamber 267 to generate vortex driver jets 245. The air passing through cooling openings 232 and 232 together with the vortex driver jets 243A, 243B, and 245 help generate and drive the vortex 295 and the vortex 297.

By opening the plurality of first doors 309A and the plurality of second doors 311A, a portion of the compressed airflow 282 escapes or bleeds through the plurality of outer openings 308A and the plurality of inner openings 310A, respectively, to outside of the combustor 306. The plurality of outer openings 308A and the plurality of inner openings 310A are provided near the upstream end 263 of the outer casing 264 and upstream of the driver openings 242 and 244. As a result, less airflow enters the passages 280 and 290 decreasing passage velocity and increasing static pressure. Airflow pressure available at the driver openings 242 in the outer liner 252 is increased and airflow pressure at the driver openings 244 in the inner liner 254 is increased. Hence, more airflow passes through the driver openings 242 in the outer liner 252 and through the driver openings 244 in the inner liner 254 near the dome structure 256, thus strengthening the vortex 295 and/or the vortex 297.

FIG. 3B is a cross-sectional side view of a combustor 306 in the combustion section 26 of the turbo-engine 16 shown in FIG. 1, the combustor 306 having a plurality of openings provided in a casing of the combustor, the openings located near a downstream end of the combustor being in an open position, according to another embodiment of the present disclosure. In this embodiment, the plurality of first doors 309A and the plurality of second doors 311A that are configured to cover the plurality of outer openings 308A and the plurality of inner openings 310A, respectively, provided near the upstream end 263 of the outer casing 264 of the combustor 306 are shown in a closed position. In addition, as shown in FIG. 3B, the plurality of first doors 309B and the plurality of second doors 311B that are configured to cover the plurality of outer openings 308B and the plurality of inner openings 310B, respectively, provided near the downstream end 255 of the combustor 306 are shown in an open position.

By closing the plurality of first doors 309A and the plurality of second doors 311A and opening the plurality of first doors 309B and the plurality of second doors 311B, a portion of the compressed airflow 282 escapes or does not bleed through the plurality of outer openings 308A and the plurality of inner openings 310A provided near the upstream end 263 of the outer casing 264 and instead bleeds through the plurality of outer openings 308B and the plurality of inner openings 310B provided near the downstream end 255 of the combustor 306. As a result, more airflow moves through the passages 288 and 290 increasing velocity and decreasing static pressure available at the driver openings 242 in the outer liner 252 and less airflow pressure is available at the driver openings 244 in the inner liner 254. Hence, less airflow passes through the driver openings 242 and aft driver openings 242A in the outer liner 252 and through the driver openings 244 in the inner liner 254 near the dome structure 256. As a result, less airflow is available and, thus, the jet momentum through the driver openings 242 and aft driver openings 242A and through the driver openings 244 is decreased, which, in turn, weakens the vortex 295 and/or the vortex 297.

Therefore, by opening and/or closing casing take-out bleeds (e.g., openings with doors or valves) in the casing of the combustor located at various axial positions, a local pressure at the vortex driver jets 243A, 243B, and 245 can be changed to strengthen or to weaken the vortex. Therefore, the casing of the combustor can be provided with airflow bleeds to increase driver jet momentum for strengthening or weakening vortex stability, as desired.

FIG. 4A is a cross-sectional side view of a combustor 406 in the combustion section 26 of the turbo-engine 16 shown in FIG. 1, the combustor 406 having a plurality of openings provided in a casing of the combustor, the openings located near a downstream end of the combustor being in an open position for introducing airflow into the combustor 406, according to another embodiment of the present disclosure. The embodiment shown in FIG. 4A is similar in many aspects to the embodiment shown in FIG. 3B. As shown in FIG. 4A, the outer casing 264 and the inner casing 266 of the combustor 406 are provided with, respectively, a plurality of outer openings 408 and a plurality of inner openings 410 near the downstream end 255 of the combustor 406. In FIG. 3B, the plurality of outer openings 308B and the plurality of inner openings 310B are used as bleed openings to allow a portion of compressed airflow 282A in the outer flow passage 288 to escape or to bleed from the outer flow passage 288 through the plurality of outer openings 308B and a portion of compressed airflow 282 in the inner flow passage 290 to escape or to bleed through the plurality of inner openings 310B. In the embodiment shown in FIG. 4A, the plurality of outer openings 408 and the plurality of inner openings 410 are provided in the outer casing 264 and the inner casing 266, respectively, and are used to introduce airflow 408A and 410A into the combustor 406 (e.g., into the outer flow passage 288 and inner flow passage 290, respectively) near or in the vicinity of the driver openings 242 and 244, respectively. The introduction of airflow 408A through the plurality of outer openings 408 and the introduction of airflow 410A through the plurality of inner openings 410 increases back-pressure at the driver openings 242 and 244 leading to an increase in vortex driver jet momentum of the vortex driver jets 243B and 245 to strengthen vortex stability without adding or subtracting airflow from the combustor 406. Therefore, the introduction of airflow through the plurality of outer openings 408 and through the plurality of inner openings 410 provides the ability to control airflow pressure within the driver openings 242 and 244, to increase vortex driver jet momentum to strengthen vortex stability inside the combustion chamber 267.

FIG. 4B is a cross-sectional side view of a combustor 407 in the combustion section 26 of the turbo-engine 16 shown in FIG. 1, the combustor 407 having a plurality of openings provided in a casing of the combustor, the openings located near a downstream end of the combustor being in communication with a sealed chamber, according to another embodiment of the present disclosure. The combustor 407 shown in FIG. 4B is similar in many aspects to the combustor 406 shown in FIG. 4A. For example, in the combustor 407, a plurality of outer openings 409 and a plurality of inner openings 411 are provided in the outer casing 264 and inner casing 266, respectively, near or in the vicinity of the driver openings 242 and 244.

An outer sealed chamber 419 having a cavity 419A is coupled (e.g., mounted) to the outer casing 264 near the plurality of outer openings 409 such that the plurality of outer openings 409 communicate with the cavity 419A of the outer sealed chamber 419. A membrane (or piston) 419B can be provided inside the cavity 419A of the outer sealed chamber 419. A portion of the compressed airflow 282A can enter the cavity 419A of the outer sealed chamber 419 through the plurality of outer openings 409. The portion of compressed airflow 282A induces oscillations of the membrane 419B (or piston). As a result, a throttling synthetic jet 429 is generated by the oscillations of the oscillating membrane 419B without adding additional flow into the combustor 407. The throttling synthetic jet 429 adds airflow energy through the driver opening 242 and the aft driver opening 242A to generate vortex driver jets 243A and 243B within the combustion chamber 267 to enhance the formation of the vortex 295 and/or vortex 297 within the combustion chamber 267.

Similarly, an inner sealed chamber 421 having a cavity 421A is coupled (e.g., mounted) to the inner casing 266 near the plurality of inner openings 411 such that the plurality of inner openings 411 communicate with the cavity 421A of the inner sealed chamber 421. A membrane (or piston) 421B can be provided inside the cavity 421A of the inner sealed chamber 421. A portion of the compressed airflow 282B can enter the cavity 421A of the inner sealed chamber 421 through the plurality of inner openings 411. The portion of compressed airflow 282B induces oscillations of the membrane (or piston) 421B. As a result, a throttling synthetic jet 431 is generated by the oscillation of the oscillating membrane 421B without adding additional flow into the combustor 407. The throttling synthetic jet 431 adds airflow energy through the driver opening 244 to generate the vortex driver jet 245 within the combustion chamber 267 to enhance the formation of the vortex 295 and/or the vortex 297 within the combustion chamber 267.

Therefore, by using the outer sealed chamber 419 and/or the inner sealed chamber 421, a throttling synthetic jet 429 and/or 431 can be generated without using any additional mechanical actuator, large valve, or piping. The throttling synthetic jet 429 and/or 431 can be created by using the oscillations of the membrane 419B and/or the membrane 421B.

Although the outer sealed chamber 419 and the inner sealed chamber 421 are shown in FIG. 4B as having a rectangular shape, any shape can be used including, but not limited to, a polygonal shape, a cylindrical shape, a spherical shape, or a more complex shape. In addition, although one membrane 419B and one membrane 421B are shown in FIG. 4B, more than one membrane also can be used. In addition, a plurality of chambers may be used or a single three hundred sixty degrees (360 deg.) circumferential chamber with a single membrane or multiple membranes may be used with a singular or multiple outlets. The membrane 419B and the membrane 421B can be, for example, metallic. The membrane 419B and the membrane 421B can be positioned at 50% to 90% of a height of the cavity from the opening of the outer sealed chamber 419 and the inner sealed chamber 421, respectively. In addition, each of the plurality of outer openings 409 and each of the plurality of inner openings 411 can be provided with an individual or a separate outer sealed chamber 419 and individual or separate inner sealed inner chamber 412, respectively.

FIG. 5 is a cross-sectional side view of a combustor 506 in the combustion section 26 of the turbo-engine 16 shown in FIG. 1, the combustor 506 having a back pressure plate, according to another embodiment of the present disclosure. The combustor 506 can be similar in many aspects to the combustor 306 shown in FIG. 3A, for example. As shown in FIG. 5, the combustor 506 includes a plurality of outer openings 508 provided in the outer casing 264 and a plurality of inner openings 510 provided in the inner casing 266. For example, the plurality of outer openings 508 and the plurality of inner openings 510 can be similar to the plurality of outer openings 308A and the plurality of inner openings 310A of the combustor 306 shown in FIG. 3A. The combustor 506 further includes a back pressure plate 520 that can be provided in front of the stage-one-nozzle (S1N) feed. The back pressure plate can be located near the turbine nozzle 272. The back pressure plate 520 can be mounted at a downstream end of the combustor 506. The back pressure plate 520 is configured to raise total pressure loss back to legacy levels to increase pressure across the outer liner 252 and the inner liner 254. Increasing pressure across the outer liner 252 and the inner liner 254 increases a pressure of the airflow at the driver openings 242 and 244, thus, providing increased driver jet momentum and penetration to enhance the vortex 295 and/or vortex 297 within the combustion chamber 267 of the combustor 506.

In an embodiment, the back pressure plate 520 includes an outer annular ring 522 and an inner annular ring 524. The outer annular ring 522 is coupled (e.g., mounted) to the outer casing 264 and the inner annular ring 524 is coupled (e.g., mounted) to the inner casing 266. The outer annular ring 522 includes a first concentric outer annular ring 522A and a second concentric outer annular ring 522B. Similarly, the inner annular ring 524 includes a first concentric inner annular ring 524A and a second concentric inner annular ring 524B.

FIGS. 6A to 6C are schematic examples of the outer annular ring 522 of the back pressure plate 520 in various configurations, according to an embodiment of the present disclosure. Although not shown in FIGS. 6A to 6C, the inner annular ring 524 of the back pressure plate 520 has a same configuration as the outer annular ring 522. The inner annular ring 524, however, has a smaller diameter than a diameter of the outer annular ring 522. Therefore, in the following paragraphs, the structure of the outer annular ring 522 is described in detail and the same description applies to the inner annular ring 524. The first concentric outer annular ring 522A and the second concentric outer annular ring 522B have a plurality of holes 522C. The first concentric outer annular ring 522A and the second concentric outer annular ring 522B are configured to rotate with respect to each other around their common center. For example, while the second concentric outer annular ring 522B is fixed, the first concentric outer annular ring 522A can be rotated with respect to the second concentric outer annular ring 522B so as to align the plurality of holes 522C of the first concentric outer annular ring 522A and the second concentric outer annular ring 522B. FIGS. 6A to 6C show examples of positions of the plurality of holes 522C. When the plurality of holes 522C of the first concentric outer annular ring 522A are fully aligned with the plurality of holes 522C of the second concentric outer annular ring 522B, the back pressure plate 520 is said to be fully open, as shown in FIG. 6A. When the plurality of holes 522C of the first concentric outer annular ring 522A are partially aligned with the plurality of holes 522C of the second concentric outer annular ring 522B, the back pressure plate 520 is said to be partially open, as shown in FIG. 6B. When the plurality of holes 522C of the first concentric outer annular ring 522A are not aligned with the plurality of holes 522C of the second concentric outer annular ring 522B, the back pressure plate 520 is said to be closed.

FIGS. 7A to 7C are schematic examples of the outer annular ring 522 of the back pressure plate 520 in various configurations, according to another embodiment of the present disclosure. As shown in FIG. 7A to 7C, the first concentric outer annular ring 522A and the second concentric outer annular ring 522B are configured to move (e.g., translate) with respect to each other. For example, while the second concentric outer annular ring 522B is fixed, the first concentric outer annular ring 522A can be translated with respect to the second concentric outer annular ring 522B so as to align the plurality of holes 522C of the first concentric outer annular ring 522A and the second concentric outer annular ring 522B. FIGS. 7A to 7C show examples of positions of the plurality of holes 522C. When the plurality of holes 522C of the first concentric outer annular ring 522A are fully aligned with the plurality of holes 522C of the second concentric outer annular ring 522B, the back pressure plate 520 is said to be fully open, as shown in FIG. 7A. When the plurality of holes 522C of the first concentric outer annular ring 522A are partially aligned with the plurality of holes 522C of the second concentric outer annular ring 522B, the back pressure plate 520 is said to be partially open, as shown in FIG. 7B. When the plurality of holes 522C of the first concentric outer annular ring 522A are not aligned with the plurality of holes 522C of the second concentric outer annular ring 522B, the back pressure plate 520 is said to be closed.

In an embodiment, a mechanical actuator can be used to move (e.g., rotate or translate) the first concentric outer annular ring 522A and the second concentric outer annular ring 522B with respect to each other. In another embodiment, instead of using an actuator, the first concentric outer annular ring 522A and the second concentric outer annular ring 522B can be moved with respect to each other using thermal differences (i.e., thermal expansion).

For example, by using back pressure plate 520 as a single static plate or two plates that move relative to one another, the back-pressure can be varied to regulate the airflow pressure through the driver openings to optimize combustor performance.

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

A combustor having a plurality of openings to control a vortex driver jet within the combustor. The combustor includes an outer casing and an inner casing, the outer casing and the inner casing extending circumferentially about a longitudinal combustor centerline axis and extending longitudinally from an upstream end of the combustor to a downstream end of the combustor, an outer liner spaced apart from the outer casing to define therebetween an outer flow passage and an inner liner spaced apart from the inner casing to define therebetween an inner flow passage, the inner liner and the outer liner being joined to a dome structure, the dome structure, the outer liner and the inner liner defining a combustion chamber, the outer liner or the inner liner, or both, having a plurality of driver openings configured to deliver a vortex driver jet inside the combustion chamber to drive a vortex, and a plurality of outer openings provided in the outer casing and a plurality of inner openings provided in the inner casing. The plurality of outer openings are configured to bleed airflow from the outer flow passage through the plurality of outer openings to outside of the combustor or to introduce airflow through the plurality of outer openings into the outer flow passage, and the plurality of inner openings are configured to bleed airflow from the inner flow passage through the plurality of inner openings to outside of the combustor or to introduce airflow through the plurality of inner openings into the inner flow passage, to control the vortex driver jet within the driver openings to drive the vortex inside the combustion chamber.

The combustor of the preceding clause, wherein the plurality of outer openings are provided near an upstream end of the outer casing and the plurality of inner openings are provided are provided near an upstream end of the inner casing.

The combustor of any preceding clause, wherein the plurality of outer openings are provided near a downstream end of the outer casing and the plurality of inner openings are provided near a downstream end of the inner casing.

The combustor of any preceding clause, wherein the plurality of outer openings in the outer casing are distributed circumferentially around the outer casing.

The combustor of any preceding clause, wherein the plurality of inner openings in the inner casing are distributed circumferentially around the inner casing.

The combustor of any preceding clause, wherein the plurality of outer openings are located upstream of the driver openings and the plurality of inner openings are located upstream of the driver openings, the plurality of outer openings and the plurality of inner openings being configured to introduce airflow through the plurality of outer openings into the outer flow passage and through the plurality of inner openings into the inner flow passage, to increase an airflow pressure within the plurality of driver openings to increase vortex driver jet momentum to strengthen the vortex inside the combustion chamber.

The combustor of any preceding clause, further comprising a plurality of first doors configured to cover the plurality of outer openings in the outer casing and a plurality of second doors configured to cover the plurality of inner openings in the inner casing, wherein the plurality of first doors and the plurality of second doors are configured to open and to close to control airflow through the plurality of outer openings in the outer casing and through the plurality of inner openings in the inner casing.

The combustor of any preceding clause, wherein the plurality of first doors or the plurality of second doors, or both, are mounted to the outer casing and the inner casing, respectively, using hinges.

The combustor of any preceding clause, wherein the plurality of first doors or the plurality of second doors, or both, are slidably mounted to the outer casing and the inner casing, respectively.

The combustor of any preceding clause, further comprising an actuator configured to actuate the plurality of first doors to actuate the plurality of second doors, wherein the actuator is configured to receive one or more signals to open or to close any of the plurality of first doors and to open or to close any the plurality of second doors.

The combustor of any preceding clause, wherein the first plurality of doors comprises a plurality of upstream doors and a plurality of downstream doors, and the second plurality of doors comprise a plurality of upstream doors and a plurality of downstream doors.

The combustor of any preceding clause, wherein the plurality of upstream doors of the first plurality of doors are opened to bleed airflow from the outer flow passage through the plurality of outer openings downstream of the driver openings and the plurality of upstream doors of the second plurality of doors are opened to bleed airflow from the inner flow passage through the plurality of inner openings downstream of the driver openings, to decrease airflow pressure at the driver openings.

The combustor of any preceding clause, wherein the plurality of upstream doors of the first plurality of doors and the plurality of upstream doors of the second plurality of doors are closed, and the plurality of downstream doors of the first plurality of doors are opened, to bleed airflow from the outer flow passage through the plurality of outer openings upstream of the driver openings, and the plurality of downstream doors of the second plurality of doors are opened to bleed airflow from the inner flow passage through the plurality of inner openings to increase airflow pressure at the driver openings.

The combustor of any preceding clause, further comprising an outer sealed chamber coupled to the outer casing and an inner sealed chamber coupled to the inner casing, wherein a cavity of the outer sealed chamber is configured to communicate with the plurality of outer openings and a cavity of the inner sealed chamber is configured to communicate with the plurality of inner openings.

The combustor of any preceding clause, wherein the plurality of outer openings are located downstream of the driver openings and the plurality of inner openings are located downstream of the driver openings.

The combustor of any preceding clause, further comprising a membrane provided inside the cavity of the outer sealed chamber and a membrane provided inside the cavity of the inner sealed chamber, wherein the membrane inside the cavity of the outer sealed chamber and the membrane inside the cavity of the inner sealed chamber are configured to oscillate to generate a throttling synthetic jet, to add energy to the vortex driver jet within the plurality of driver openings to enhance the vortex within the combustion chamber.

The combustor of any preceding clause, further comprising a back pressure plate mounted at a downstream end of the combustor, wherein the back pressure plate is configured to increase pressure across the outer liner and the inner liner, to increase a pressure of airflow at the driver openings to increase vortex driver jet momentum.

The combustor of any preceding clause, wherein the back pressure plate includes an outer annular ring and an inner annular ring, the outer annular ring is coupled to the outer casing and the inner annular ring is coupled to the inner casing.

The combustor of any preceding clause, wherein the outer annular ring includes a first concentric outer annular ring and a second concentric outer annular ring, the first concentric annular ring and the second concentric outer annular ring having a plurality of holes, the first concentric outer annular ring and the second concentric outer annular ring are configured to move with respect to each other to open or to close the plurality of holes.

The combustor of any preceding clause, wherein the inner annular ring includes a first concentric inner annular ring and a second concentric inner annular ring, the first concentric annular ring and the second concentric inner annular ring having a plurality of holes, and the first concentric inner annular ring and the second concentric inner annular ring are configured to move with respect to each other to open or to close the plurality of holes.

A turbine engine including a combustor having a plurality of openings to control a vortex driver jet within the combustor. The combustor includes an outer casing and an inner casing, the outer casing and the inner casing extending circumferentially about a longitudinal combustor centerline axis and extending longitudinally from an upstream end of the combustor to a downstream end of the combustor, an outer liner spaced apart from the outer casing to define therebetween an outer flow passage and an inner liner spaced apart from the inner casing to define therebetween an inner flow passage, the inner liner and the outer liner being joined to a dome structure, the dome structure, the outer liner and the inner liner defining a combustion chamber, the outer liner or the inner liner, or both, having a plurality of driver openings configured to deliver a vortex driver jet inside the combustion chamber to drive a vortex, and a plurality of outer openings provided in the outer casing and a plurality of inner openings provided in the inner casing. The plurality of outer openings are configured to bleed airflow from the outer flow passage through the plurality of outer openings to outside of the combustor or to introduce airflow through the plurality of outer openings into the outer flow passage, and the plurality of inner openings are configured to bleed airflow from the inner flow passage through the plurality of inner openings to outside of the combustor or to introduce airflow through the plurality of inner openings into the inner flow passage, to control the vortex driver jet within the driver openings to drive the vortex inside the combustion chamber.

The turbine engine of the preceding clause, wherein the plurality of outer openings are provided near an upstream end of the outer casing and the plurality of inner openings are provided are provided near an upstream end of the inner casing.

The turbine engine of any preceding clause, wherein the plurality of outer openings are provided near a downstream end of the outer casing and the plurality of inner openings are provided near a downstream end of the inner casing.

The turbine engine of any preceding clause, wherein the plurality of outer openings in the outer casing are distributed circumferentially around the outer casing.

The turbine engine of any preceding clause, wherein the plurality of inner openings in the inner casing are distributed circumferentially around the inner casing.

The turbine engine of any preceding clause, wherein the plurality of outer openings are located upstream of the driver openings and the plurality of inner openings are located upstream of the driver openings, the plurality of outer openings and the plurality of inner openings being configured to introduce airflow through the plurality of outer openings into the outer flow passage and through the plurality of inner openings into the inner flow passage, to increase an airflow pressure within the plurality of driver openings to increase vortex driver jet momentum to strengthen the vortex inside the combustion chamber.

The turbine engine of any preceding clause, further comprising a plurality of first doors configured to cover the plurality of outer openings in the outer casing and a plurality of second doors configured to cover the plurality of inner openings in the inner casing, wherein the plurality of first doors and the plurality of second doors are configured to open and to close to control airflow through the plurality of outer openings in the outer casing and through the plurality of inner openings in the inner casing.

The turbine engine of any preceding clause, wherein the plurality of first doors or the plurality of second doors, or both, are mounted to the outer casing and the inner casing, respectively, using hinges.

The turbine engine of any preceding clause, wherein the plurality of first doors or the plurality of second doors, or both, are slidably mounted to the outer casing and the inner casing, respectively.

The turbine engine of any preceding clause, further comprising an actuator configured to actuate the plurality of first doors to actuate the plurality of second doors, wherein the actuator is configured to receive one or more signals to open or to close any of the plurality of first doors and to open or to close any the plurality of second doors.

The turbine engine of any preceding clause, wherein the first plurality of doors comprises a plurality of upstream doors and a plurality of downstream doors, and the second plurality of doors comprise a plurality of upstream doors and a plurality of downstream doors.

The turbine engine of any preceding clause, wherein the plurality of upstream doors of the first plurality of doors are opened to bleed airflow from the outer flow passage through the plurality of outer openings downstream of the driver openings and the plurality of upstream doors of the second plurality of doors are opened to bleed airflow from the inner flow passage through the plurality of inner openings downstream of the driver openings, to decrease airflow pressure at the driver openings.

The turbine engine of any preceding clause, wherein the plurality of upstream doors of the first plurality of doors and the plurality of upstream doors of the second plurality of doors are closed, and the plurality of downstream doors of the first plurality of doors are opened, to bleed airflow from the outer flow passage through the plurality of outer openings upstream of the driver openings, and the plurality of downstream doors of the second plurality of doors are opened to bleed airflow from the inner flow passage through the plurality of inner openings to increase airflow pressure at the driver openings.

The turbine engine of any preceding clause, further comprising an outer sealed chamber coupled to the outer casing and an inner sealed chamber coupled to the inner casing, wherein a cavity of the outer sealed chamber is configured to communicate with the plurality of outer openings and a cavity of the inner sealed chamber is configured to communicate with the plurality of inner openings.

The turbine engine of any preceding clause, wherein the plurality of outer openings are located downstream of the driver openings and the plurality of inner openings are located downstream of the driver openings.

The turbine engine of any preceding clause, further comprising a membrane provided inside the cavity of the outer sealed chamber and a membrane provided inside the cavity of the inner sealed chamber, wherein the membrane inside the cavity of the outer sealed chamber and the membrane inside the cavity of the inner sealed chamber are configured to oscillate to generate a throttling synthetic jet, to add energy to the vortex driver jet within the plurality of driver openings to enhance the vortex within the combustion chamber.

The turbine engine of any preceding clause, further comprising a back pressure plate mounted at a downstream end of the combustor, wherein the back pressure plate is configured to increase pressure across the outer liner and the inner liner, to increase a pressure of airflow at the driver openings to increase vortex driver jet momentum.

The turbine engine of any preceding clause, wherein the back pressure plate includes an outer annular ring and an inner annular ring, the outer annular ring is coupled to the outer casing and the inner annular ring is coupled to the inner casing.

The turbine engine of any preceding clause, wherein the outer annular ring includes a first concentric outer annular ring and a second concentric outer annular ring, the first concentric annular ring and the second concentric outer annular ring having a plurality of holes, the first concentric outer annular ring and the second concentric outer annular ring are configured to move with respect to each other to open or to close the plurality of holes.

The turbine engine of any preceding clause, wherein the inner annular ring includes a first concentric inner annular ring and a second concentric inner annular ring, the first concentric annular ring and the second concentric inner annular ring having a plurality of holes, and the first concentric inner annular ring and the second concentric inner annular ring are configured to move with respect to each other to open or to close the plurality of holes.

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