Combustor with baffled liners in a turbine engine

Abstract

A combustor for a turbine engine includes a dome, an outer liner having a plurality of outer liner holes and one or more outer liner dilution holes, an inner liner having a plurality of inner liner holes and one or more inner liner dilution holes, the dome, the outer liner and the inner liner defining a combustion chamber, an outer baffle connected to the outer liner to define an outer cavity therebetween, the outer baffle including a plurality of outer baffle holes, and an inner baffle connected to the inner liner to define an inner cavity therebetween, the inner baffle including a plurality of inner baffle holes. The one or more outer liner dilution holes and the one or more inner liner dilution holes are configured to provide pressure drop across the outer liner and the inner liner to enhance penetration of a jet airflow into the combustion chamber.

Description

TECHNICAL FIELD

[0001] The present disclosure relates generally to a combustor with baffled liners in a turbine engine.BACKGROUND

[0002] Gas turbine engines generally include a fan and a turbo-engine arranged in flow communication with one another, with the turbo-engine disposed downstream of the fan in a direction of flow through the gas turbine engine. The turbo-engine generally includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. With multi-shaft gas turbine engines, the compressor section can include a high pressure (HP) compressor disposed downstream of a low pressure (LP) compressor, and the turbine section can similarly include a low pressure (LP) turbine disposed downstream of a high pressure (HP) turbine. The combustion section includes a combustor that generates combustion gases discharged into the turbine section. With such a configuration, the HP compressor is coupled with the HP turbine via a high pressure (HP) shaft, and the LP compressor is coupled with the LP turbine via a low pressure (LP) shaft.BRIEF DESCRIPTION OF THE DRAWINGS

[0003] Features and advantages will be apparent from the following, more particular, description of various embodiments, as illustrated in the accompanying drawings, wherein like reference numbers generally indicate identical, functionally similar, or structurally similar elements, or both.

[0004] FIG. 1 is a schematic, cross-sectional diagram of a turbine engine, taken along a longitudinal centerline axis of the turbine engine, according to an embodiment of the present disclosure.

[0005] FIG. 2 is a schematic, cross-sectional view of a combustion section of the turbine engine of FIG. 1, according to an embodiment of the present disclosure.

[0006] FIG. 3 is a schematic, cross-sectional view of a combustor that can be used in the combustor section of the turbine engine of FIG. 1, according to another embodiment of the present disclosure.

[0007] FIG. 4 is a schematic, cross-sectional view of a combustor that can be used in the combustor section of the turbine engine of FIG. 1, according to yet another embodiment of the present disclosure.

[0008] FIG. 5 is an elevational view of a portion of an outer liner of the combustor shown in FIG. 3, the outer liner having a plurality of outer liner holes and one or more outer liner dilution holes, according to an embodiment of the present disclosure.DETAILED DESCRIPTION

[0009] 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, both the foregoing summary of the present disclosure and the following detailed description are exemplary and intended to provide further explanation without limiting the scope of the disclosure as claimed.

[0010] 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.

[0011] 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.

[0012] 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.

[0013] The terms “low” and “high,” or their respective comparative degrees (e.g., “lower” and “higher,” where applicable), when used with the compressor, turbine, shaft, or spool components, each refers to relative pressures or relative speeds, or both within an engine unless otherwise specified. For example, a “low-speed” component defines a component configured to operate at a rotational speed, such as a maximum allowable rotational speed, which is lower than that of a “high-speed” component of the engine. Alternatively, unless otherwise specified, the aforementioned terms may be understood in their superlative degree. For example, a “low-pressure turbine” may refer to the lowest maximum pressure within a turbine section, and a “high-pressure turbine” may refer to the highest maximum pressure within the turbine section. The terms “low” or “high” in such aforementioned regards may additionally, or alternatively, be understood as relative to minimum allowable speeds or pressures, or minimum or maximum allowable speeds or pressures relative to normal, desired, steady state, etc., operation of the engine, or any combination thereof.

[0014] 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.

[0015] The singular forms “a,”“an,” and “the” include plural references unless the context clearly dictates otherwise.

[0016] As used herein, the term “axial” refers to directions and orientations that extend substantially parallel to a longitudinal centerline of the turbine engine. Moreover, the terms “radial” and “radially” refer to directions and orientations that extend substantially perpendicular to the longitudinal centerline of the turbine engine. In addition, as used herein, the term “circumferentially” refers to directions and orientations that extend arcuately about the longitudinal centerline of the turbine engine.

[0017] For a rich burn combustor, there is a desire to increase thermal capability of metallic liners or additive liners by adding a baffle, but allowing a full air pressure drop through a dilution hole to maintain the jet penetration and subsequent emissions characteristics. Conventional techniques use individual ‘tower and cap’ dilutions holes. However, this may be cumbersome and expensive to implement on one hundred or more dilution holes. In an embodiment, in order to achieve desired pressure characteristics, a forward baffle and an aft baffle are provided. The use of the forward baffle and the aft baffle allows for full-air pressure drop dilution annularly and, thus, avoid using individual tower / cap assemblies.

[0018] In an embodiment, a combustor includes inner and outer baffles connected to inner and outer liners, respectively. In an embodiment, the combustor includes two inner baffles and two outer baffles connected to the inner and outer liner, respectively. For example, a first baffle (forward baffle) is connected to a forward part of a liner (inner liner or outer liner) and a second baffle (an aft baffle) is connected to an aft part of the liner (inner liner or outer liner).

[0019] The use of the forward baffle and the aft baffle provides the ability to maintain double air pressure drop characteristics to the cooling liners (durability), while allowing for a full air pressure drop to the dilution holes (emissions). This increases the durability of metallic or additively enabled combustor liners (inner liner and outer liner).

[0020] Referring now to the drawings, FIG. 1 is a schematic, cross-sectional diagram 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.

[0021] The turbo-engine 16 depicted in FIG. 1 generally includes an outer casing 18 that is substantially tubular and defines an annular inlet 20. As schematically shown in FIG. 1, the outer casing 18 encases, in serial flow relationship, a compressor section 21 including a booster or a low-pressure (LP) compressor 22 followed downstream by a high-pressure (HP) compressor 24, a combustion section 26, a turbine section 27 including a high-pressure (HP) turbine 28 followed downstream by a low-pressure (LP) turbine 30, and a jet exhaust nozzle section 32. A high-pressure (HP) shaft 34 or a spool drivingly connects the HP turbine 28 to the HP compressor 24, to rotate the HP turbine 28 and the HP compressor 24 in unison. A low-pressure (LP) shaft 36 or a spool drivingly connects the LP turbine 30 to the LP compressor 22, to rotate the LP turbine 30 and the LP compressor 22 in unison. The compressor section 21, the combustion section 26, the turbine section 27, and the jet exhaust nozzle section 32 together define a core air flowpath.

[0022] 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 (airfoils) 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. Each fan blade 40 is 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. The gearbox assembly 46 is shown schematically in FIG. 1. The gearbox assembly 46 includes a plurality of gears for adjusting the rotational speed of the fan shaft 45 and, thus, the fan 38 relative to the LP shaft 36.

[0023] Referring still to the embodiment of FIG. 1, the disk 42 is covered by a rotatable fan hub 48 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 or at least a portion of the turbo-engine 16, or both. The nacelle 50 is supported relative to the turbo-engine 16 by a plurality of circumferentially spaced outlet guide vanes 52 (stator airfoils). Moreover, a downstream section 54 of the nacelle 50 extends over an outer portion of the turbo-engine 16 to define a bypass airflow passage 56 therebetween.

[0024] 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, or both. As the volume of air 58 passes across the fan blades 40 (rotating airfoils), a first portion of air 62 is directed or routed into the bypass airflow passage 56, and a second portion of air 64 is directed or is routed into the upstream section of the core air flowpath, or, more specifically, into the annular inlet 20 of the LP compressor 22. The ratio between the first portion of air 62 and the second portion of air 64 is commonly known as a bypass ratio. The pressure of the second portion of air 64 is then increased, forming compressed air 65, and the compressed air 65 is routed through the HP compressor 24 and into the combustion section 26, where the compressed air 65 is mixed with fuel and burned to generate combustion gases 66. The fuel can be a kerosene-based fuel, such as Jet A-1, or a hydrogen fuel. The fuel can be transported from a fuel reservoir 79 using a fuel pipe 79A to the combustion section 26.

[0025] 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, or both, from the combustion gases 66 is extracted via sequential stages of HP turbine stator vanes 68 (stator airfoils) that are coupled to the outer casing 18 and HP turbine rotor blades 70 (rotating airfoils) that are coupled to the HP shaft 34, thus, causing the HP shaft 34 to rotate, thereby supporting operation of the HP compressor 24. 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 and the kinetic energy is extracted from the combustion gases 66 via sequential stages of LP turbine stator vanes 72 (stator airfoils) that are coupled to the outer casing 18 and LP turbine rotor blades 74 (rotating airfoils) that are coupled to the LP shaft 36, thus, causing the LP shaft 36 to rotate, thereby supporting operation of the LP compressor 22 and rotation of the fan 38 via the gearbox assembly 46.

[0026] 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 pressure of the first portion of air 62 is substantially increased as the first portion of 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.

[0027] The turbine engine 10 depicted in FIG. 1 is by way of example only. In other embodiments, the turbine engine 10 may have any other suitable configuration. For example, in other 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. Moreover, in other embodiments, any other suitable number or configuration of compressors, turbines, shafts, or a combination thereof may be provided. In still other 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, or turboshaft engines.

[0028] The turbine engine 10 discussed herein is suitable for use on aircraft. Suitable aircraft include, for example, airplanes and unmanned aerial vehicles (UAV). In other embodiments, the turbine engine can be any other turbine engine, such as an industrial turbine engine incorporated into a power generation system, or a nautical turbine engine on a ship or other vessel.

[0029] FIG. 2 is a schematic, cross-sectional view of the combustion section 26 of the turbine engine 10, shown in FIG. 1, according to an embodiment of the present disclosure. The combustion section 26 generally includes a combustor 80 that generates the combustion gases discharged into the turbine section, or, more particularly, into the HP turbine 28 (shown in FIG. 1). The combustor 80 includes an outer liner 82, an inner liner 84, and a dome 86. The outer liner 82, the inner liner 84, and the dome 86 together define a combustion chamber 88. In addition, a diffuser 90 is positioned upstream of the combustion chamber 88. The diffuser 90 has an outer diffuser wall 90A and an inner diffuser wall 90B. The inner diffuser wall 90B is closer to the longitudinal centerline axis 12. The diffuser 90 receives an air flow from the compressor section 21 (shown in FIG. 1) and provides a flow of compressed air to the combustor 80. In an embodiment, the diffuser 90 provides the flow of compressed air to a single circumferential row of fuel-air mixers 92. However, a plurality of circumferential rows of fuel-air mixer 92 can also be used. In an embodiment, the dome 86 of the combustor 80 is configured as a single annular dome, and the circumferential row of fuel-air mixers 92 are provided within openings formed in the dome 86 (air feeding dome or combustor dome). However, in other embodiments, a multiple annular dome can also be used. In general, other types of combustors can also be used.

[0030] In an embodiment, the diffuser 90 can be used to slow the high speed, highly compressed air from the HP compressor 24 (shown in FIG. 1) to a velocity optimal for the combustor 80. Furthermore, the diffuser 90 can also be configured to limit flow distortion as much as possible by avoiding undesirable flow effects such as boundary layer separation. Similar to most other turbine engine components, the diffuser 90 is generally designed to be as light as possible to reduce weight of the overall engine.

[0031] A fuel nozzle (not shown) provides fuel transported by fuel pipe 79A from fuel reservoir 79 (shown in FIG. 1) to fuel-air mixers 92 depending upon a desired performance of the combustor 80 at various engine operating states. In the embodiment shown in FIG. 2, an outer cowl 94 (e.g., an annular cowl) and an inner cowl 96 (e.g., an annular cowl) are located upstream of the combustion chamber 88 so as to direct air flow into fuel-air mixers 92. The outer cowl 94 and the inner cowl 96 may also direct a portion of the flow of air from the diffuser 90 to an outer passage 98 defined between the outer liner 82 and an outer casing 100, and an inner passage 102 defined between the inner liner 84 and an inner casing 104. In addition, an inner support cone 106 is further shown as being connected to a nozzle support 108 using a plurality of bolts 110 and nuts 112. Other combustion sections, however, may include any other suitable structural configurations.

[0032] The combustor 80 also has an igniter 114. The igniter 114 is provided to ignite the fuel-air mixture supplied to combustion chamber 88 of the combustor 80. The igniter 114 is attached to the outer casing 100 of the combustor 80 in a substantially fixed manner. Additionally, the igniter 114 extends generally along an axial direction A2, defining a distal end 116 that is positioned proximate to an opening in a combustor member of the combustor 80. The distal end 116 is positioned proximate to an opening 118 within the outer liner 82 of the combustor 80.

[0033] In an embodiment, the dome 86 of the combustor 80, together with the outer liner 82, the inner liner 84, and fuel-air mixers 92, form the combustion chamber 88 and define a swirling flow 130. The air flows through the fuel-air mixers 92 as the air enters the combustion chamber 88. The role of the dome 86 and the fuel-air mixers 92 is to generate turbulence in the air flow to rapidly mix the air with the fuel to create a fuel-air mixture. The fuel-air mixers 92 (also called a swirlers or mixers) establish a local low pressure zone that forces some of the combustion products to recirculate, as illustrated in FIG. 2, creating needed high turbulence.

[0034] The outer liner 82 includes a plurality of outer liner cooling holes 202. The inner liner 84 includes a plurality of inner liner cooling holes 204. The plurality of outer liner cooling holes 202 and the plurality of inner liner cooling holes 204 are provided to cool down the outer liner 82 and the inner liner 84, respectively, as the outer liner 82 and the inner liner 84 can be exposed to the hot environment of the combustion chamber 88.

[0035] FIG. 3 is a schematic, cross-sectional view of a combustor 300 that can be used in the combustor section 26 of the turbine engine 10, shown in FIG. 1, according to another embodiment of the present disclosure. As shown in FIG. 3, the combustor 300 includes an outer liner 302, an inner liner 304, and a dome 306. The outer liner 302, the inner liner 304, and the dome 306 together define a combustion chamber 308.

[0036] In an embodiment, the dome 306 of the combustor 300 is configured as a single annular dome. In other embodiments, multiple annular domes can also be used. In general, other types of combustors can also be used.

[0037] The outer liner 302 includes a plurality of outer liner holes 302A. The inner liner 304 includes a plurality of inner liner holes 304A. The plurality of outer liner holes 302A and the plurality of inner liner holes 304A are configured to provide a partial air pressure drop across the outer liner 302 and the inner liner 304 to cool down the outer liner 302 and the inner liner 304, respectively, as the outer liner 302 and the inner liner 304 can be exposed to the hot environment of the combustion chamber 308. In addition to the plurality of outer liner holes 302A, the outer liner 302 also includes one or more outer liner dilution holes 302B. Similarly, in addition to the plurality of inner liner holes 304A, the inner liner 304 also includes one or more inner liner dilution holes 304B. The one or more outer liner dilution holes 302B and the one or more inner liner dilution holes 304B are configured to provide an additional partial air pressure drop across the outer liner 302 and across the inner liner 304 to enhance jet airflow penetration into the combustion chamber 308. In an embodiment, a diameter of the one or more outer liner dilution holes 302B is greater than a diameter of the plurality of outer liner holes 302A. In an embodiment, a diameter of the one or more inner liner dilution holes 304B is greater than a diameter of the plurality of inner liner holes 304A.

[0038] The combustor 300 also includes an outer baffle 312 and an inner baffle 314. The outer baffle 312 has a plurality of outer baffle holes 312A. In an embodiment, the outer baffle 312 or the inner baffle 314, or both, can be metallic baffles or additive baffles. In another embodiment, the outer baffle 312 or the inner baffle 314, or both can be made from a composite material. The inner baffle has a plurality of inner baffle holes 314A. The outer baffle 312 is connected to the outer liner 302 and the inner baffle 314 is connected to the inner liner 304 to define, respectively, an outer cavity 316 between the outer baffle 312 that is connected to the outer liner 302, and an inner cavity 318 between the inner baffle 314 and the inner liner 304.

[0039] In operation, airflow 322 penetrates into the outer cavity 316 through the outer baffle holes 312A. Airflow 324 penetrates into the inner cavity 318 through the inner baffle holes 314A. Air is distributed within the outer cavity 316 and the inner cavity 318. A first portion of airflow 326 from the air within the outer cavity 316 penetrates into the combustion chamber 308 through the plurality of outer liner holes 302A. In addition, a first jet airflow 328 from the air within the outer cavity 316 penetrates into the combustion chamber 308 through the one or more outer liner dilution holes 302B. A second portion of airflow 330 from the air within the inner cavity 318 penetrates into the combustion chamber 308 through the plurality of inner liner holes 304A. In addition, a second jet airflow 332 from the air within the inner cavity 318 penetrates into the combustion chamber 308 through the one or more inner liner dilution holes 304B. The first jet airflow 328 and the second jet airflow 332 further enhance jet airflow penetration into the combustion chamber 308. In an embodiment, the one or more outer liner dilution holes 302B and / or the one or more inner liner dilution holes 304B are provided to introduce airflow jets (e.g., the first jet airflow 328 and / or the second jet airflow 332) into the combustion chamber 308 to stop or to “quench” the reaction to reduce or substantially eliminate undesirable emissions of gases, such as NOx.

[0040] FIG. 4 is a schematic, cross-sectional view of a combustor 400 that can be used in the combustor section 26 of the turbine engine 10, shown in FIG. 1, according to yet another embodiment of the present disclosure. The combustor 400 is similar in many aspects to the combustor 300 shown in FIG. 3. Therefore, only distinguishing features will be discussed further in detail herein. Similar to the combustor 300, the combustor 400 also includes an outer baffle 412 and an inner baffle 414. The outer baffle 412 has a plurality of outer baffle holes 412A. The inner baffle has a plurality of inner baffle holes 414A. The outer baffle 412 is connected to the outer liner 302 to define an outer cavity 416 between the outer baffle 412 and the outer liner 302. The inner baffle 414 is connected to the inner liner 304 to define an inner cavity 418 between the inner baffle 314 and the inner liner 304.

[0041] The outer baffle 412 has a forward outer baffle portion 413 that is connected to a forward portion 303 of the outer liner 302 to define a forward outer cavity 416A and an aft outer baffle portion 415 that is connected to an aft portion 305 of the outer liner 302 to define an aft outer cavity 416B. The inner baffle 414 has a forward inner baffle portion 417 that is connected to a forward portion 307 of the inner liner 304 to define a forward inner cavity 418A and an aft inner baffle portion 419 that is connected to an aft portion 309 of the inner liner 304 to define an aft inner cavity 418B.

[0042] As shown in FIG. 4, the forward outer baffle portion 413 is connected to a first end 303A of the forward portion 303 and to a second end 303B of the forward portion 303. The aft outer baffle portion 415 is connected to a first end 305A of the aft portion 305 and to a second end 305B of the aft portion 305. The second end 303B of the forward portion 303 and the second end 305B of the aft portion 305 are located near the one or more outer liner dilution holes 302B. Similarly, the forward inner baffle portion 417 is connected to a first end 307A of the forward portion 307 and to a second end 307B of the forward portion 307. The aft inner baffle portion 419 is connected to a first end 309A of the aft portion 309 and to a second end 309B of the aft portion 309. The second end 307B of the forward portion 307 and the second end 309B of the aft portion 309 are located near the one or more inner liner dilution holes 304B.

[0043] In this embodiment, the forward outer cavity 416A and the aft outer cavity 416B are not in communication with the one or more outer liner dilution holes 302B in the outer liner 302. Similarly, the forward inner cavity 418A and the aft inner cavity 418B are also not in communication with the one or more inner liner dilution holes 304B in the inner liner 304. The forward outer baffle portion 413 and the aft outer baffle portion 415 are connected to an edge of the one or more outer liner dilution holes 302B. The forward inner baffle portion 417 and the aft inner baffle portion 419 are connected to an edge of the one or more inner liner dilution holes 304B. Therefore, the first jet airflow 328 does not originate from within the forward outer cavity 416A or the aft outer cavity 416B. In addition, the second jet airflow 332 does not originate from within the forward inner cavity 418A or the aft inner cavity 418B. On the other hand, the first portion of airflow 326 through the plurality of outer liner holes 302A originates from the forward outer cavity 416A and the aft outer cavity 416B. The second portion of airflow 330 through the plurality of inner liner holes 304A originates from the forward inner cavity 418A and the aft inner cavity 418B.

[0044] Using the forward outer baffle portion 413 and the aft outer baffle portion 415 maintains the air pressure drop through the plurality of outer liner holes 302A and provides a full air pressure drop through the one or more outer liner dilution holes 302B in the outer liner 302, to enhance the outer liner 302 durability. The outer liner 302 can be a metallic liner or an additive liner. For example, the outer liner 302 can be shaped from a metallic material (e.g., a metallic layer) or additively manufactured.

[0045] Using the forward inner baffle portion 417 and the aft inner baffle portion 419 maintains the air pressure drop through the plurality of inner liner holes 304A and provides a full air pressure drop through the one or more inner liner dilution holes 304B in the inner liner 304, to enhance the inner liner 304 durability. The inner liner 304 can be a metallic liner or an additive liner. For example, the inner liner 304 can be shaped from a metallic material (e.g., a metallic layer) or additively manufactured.

[0046] FIG. 5 is an elevational view of a portion of the outer liner 302 of the combustor 300 with the plurality outer liner holes 302A and the one or more outer liner dilution holes 302B, according to an embodiment of the present disclosure. In the embodiment shown in FIG. 5, a plurality of outer liner dilution holes 302B are provided. The plurality of outer liner dilution holes 302B are spaced apart around a circumference of the combustor 300. In an embodiment, the plurality of outer liner dilution holes 302B are equidistantly spaced apart. In an embodiment, the plurality of outer liner holes 302A are disposed adjacent the plurality of outer liner dilution holes 302B along a circumference of the combustor 300. A diameter of the plurality of outer liner holes 302A is less than a diameter of the plurality of outer liner dilution holes 302B. Although only outer liner 302 is depicted in FIG. 5, similar features can also be provided in the inner liner 304.

[0047] In an embodiment, the outer liner 302 or the inner liner 304, or both, can be a metallic layer or an additive liner.

[0048] For a rich burn combustor, such as the combustor 300 and the combustor 400 as shown in FIGS. 3 and 4, it is desirable to increase the thermal capability of the outer liner 302 and the inner liner 304 by providing an outer baffle 312, 412 connected to the outer liner 302 and an inner baffle 314, 414 connected to the inner liner 304. For example, the use of the outer baffle 412 having a forward outer baffle portion 413 and an aft outer baffle portion 415, which are separated by the one or more outer liner dilution holes 302B, provides a full air pressure drop (i.e., a pressure difference between the outer passage and the combustion chamber 308) through the one or more outer liner dilution holes 302B to enhance jet penetration and, thus, overall gas emission characteristics. For example, the use of the inner baffle 414, having a forward inner baffle portion 417, and an aft inner baffle portion 419, which are separated by the one or more inner liner dilution holes 304B, provides a full air pressure drop through the one or more inner liner dilution holes 304B to enhance jet penetration and, thus, overall gas emission characteristics. For example, by providing a full pressure drop, higher jet airflow penetration into the combustion chamber 308 can be achieved to quench the combustion reaction inside the combustion chamber 308 and to halt the production of undesirable emission gases. The use of the forward outer baffle portion 413, the forward inner baffle portion 417, the aft outer baffle portion 415, and the aft inner baffle portion 419 allows for a full-air pressure drop dilution annularly.

[0049] The outer baffle 312, 412 and the inner baffle 314, 414 can be made from the same material as that of the outer liner 302 and the inner liner 304, such as, a superalloy that is capable of withstanding elevated temperatures, high stresses, and often highly oxidizing environments.

[0050] As described above, in an embodiment, the outer liner 302 and the inner liner 304 are provided with, respectively, the outer baffle 312, 412 and the inner baffle 314, 414. The outer baffle 312, 412 is connected to the outer liner 302 to define the outer cavity 316, 416 therebetween. The inner baffle 314, 414 is connected to the inner liner 304 to define the inner cavity 318, 418 therebetween.

[0051] In another embodiment, the outer liner 302 can be provided with the outer baffle 312, 412 while the inner liner 304 is not provided with the inner baffle 314, 414. In yet another embodiment, the inner liner 304 can be provided with the inner baffle 314, 414 while the outer liner 302 is not provided with the outer baffle 312, 412.

[0052] Further aspects are provided by the subject matter of the following clauses.

[0053] A combustor for a turbine engine includes a dome, an outer liner including a plurality of outer liner holes and one or more outer liner dilution holes, an inner liner including a plurality of inner liner holes and one or more inner liner dilution holes, the dome, the outer liner and the inner liner defining a combustion chamber, an outer baffle connected to the outer liner to define an outer cavity therebetween, the outer baffle including a plurality of outer baffle holes, and an inner baffle connected to the inner liner to define an inner cavity therebetween, the inner baffle including a plurality of inner baffle holes, the plurality of outer liner holes being configured to provide a partial air pressure drop across the outer liner and the plurality of inner liner holes are configured to provide a partial air pressure drop across the inner liner to cool down the outer liner and the inner liner, and the one or more outer liner dilution holes being configured to provide an additional pressure drop across the outer liner and the one or more inner liner dilution holes are configured to provide an additional air pressure drop across the inner liner to enhance penetration of a jet airflow into the combustion chamber.

[0054] The combustor according to the preceding clause, the inner liner, the outer liner, or the inner liner and the outer liner both being metallic liners or additive liners.

[0055] The combustor according to any preceding clause, the inner baffle, the outer baffle, or both being metallic baffles or additive baffles.

[0056] The combustor according to any preceding clause, the one or more outer liner dilution holes being configured to allow a first jet airflow to penetrate into the combustion chamber.

[0057] The combustor according to any preceding clause, the one or more inner liner dilution holes being configured to allow a second jet airflow to penetrate into the combustion chamber.

[0058] The combustor according to any preceding clause, the plurality of outer baffle holes are configured to allow an airflow to enter into the outer cavity between the outer baffle and the outer liner, and the plurality of inner baffle holes are configured to allow an airflow to enter into the inner cavity between the inner baffle and the inner liner.

[0059] The combustor according to any preceding clause, the plurality of outer liner holes being configured to allow a first airflow portion to penetrate into the combustion chamber.

[0060] The combustor according to any preceding clause, the plurality of inner liner holes being configured to allow a second airflow portion to penetrate into the combustion chamber.

[0061] The combustor according to any preceding clause, the outer baffle including a forward outer baffle portion connected to a forward portion of the outer liner, to define a forward outer cavity therebetween, and an aft outer baffle portion connected to an aft portion of the outer liner, to define an aft outer cavity therebetween.

[0062] The combustor according to any preceding clause, the forward outer cavity and the aft outer cavity not being in communication with the one or more outer liner dilution holes.

[0063] The combustor according to any preceding clause, the forward outer baffle portion and the aft outer baffle portion being connected to an edge of the one or more outer liner dilution holes.

[0064] The combustor according to any preceding clause, the jet airflow passing through the one or more outer liner dilution holes not originating from the forward outer cavity or the aft outer cavity.

[0065] The combustor according to any preceding clause, the forward outer baffle portion being connected to a first end of the forward portion of the outer liner and to a second end of the forward portion of the outer liner.

[0066] The combustor according to any preceding clause, the aft outer baffle being connected to a first end of the aft portion of the outer liner and to a second end of the aft portion of the outer liner.

[0067] The combustor according to any preceding clause, the inner baffle having a forward inner baffle portion that is connected to a forward portion of the inner liner to define a forward inner cavity, and an aft inner baffle portion connected to an aft portion of the inner liner to define an aft inner cavity.

[0068] The combustor according to any preceding clause, the forward inner cavity and the aft inner cavity not being in communication with the one or more inner liner dilution holes.

[0069] The combustor according to any preceding clause, the forward inner baffle portion and the aft inner baffle portion being connected to an edge of the one or more inner liner dilution holes.

[0070] The combustor according to any preceding clause, the jet airflow passing through one or more inner liner dilution holes not originating from the forward inner cavity or the aft inner cavity.

[0071] The combustor according to any preceding clause, the forward inner baffle portion being connected to a first end of the forward portion of the inner liner and to a second end of the forward portion of the inner liner.

[0072] The combustor according to any preceding clause, the aft inner baffle being connected to a first end of the aft portion of the inner liner and to a second end of the aft portion of the inner liner.

[0073] The combustor according to any preceding clause, wherein a diameter of the one or more outer liner dilution holes is greater than a diameter of the plurality of outer liner holes.

[0074] The combustor according to any preceding clause, wherein a diameter of the one or more inner liner dilution holes is greater than a diameter of the plurality of inner liner holes.

[0075] A turbine engine includes a combustor. The combustor includes a dome, an outer liner including a plurality of outer liner holes and one or more outer liner dilution holes, an inner liner including a plurality of inner liner holes and one or more inner liner dilution holes, the dome, the outer liner and the inner liner defining a combustion chamber, an outer baffle connected to the outer liner to define an outer cavity therebetween, the outer baffle including a plurality of outer baffle holes, and an inner baffle connected to the inner liner to define an inner cavity therebetween, the inner baffle including a plurality of inner baffle holes, the plurality of outer liner holes being configured to provide a partial air pressure drop across the outer liner and the plurality of inner liner holes are configured to provide a partial air pressure drop across the inner liner to cool down the outer liner and the inner liner, and the one or more outer liner dilution holes being configured to provide an additional pressure drop across the outer liner and the one or more inner liner dilution holes are configured to provide an additional air pressure drop across the inner liner to enhance penetration of a jet airflow into the combustion chamber.

[0076] The turbine engine according to the preceding clause, the inner liner, the outer liner, or the inner liner and the outer liner both being metallic liners or additive liners.

[0077] The turbine engine according to any preceding clause, the inner baffle, the outer baffle, or both being metallic baffles or additive baffles.

[0078] The turbine engine according to any preceding clause, the one or more outer liner dilution holes being configured to allow a first jet airflow to penetrate into the combustion chamber.

[0079] The turbine engine according to any preceding clause, the one or more inner liner dilution holes being configured to allow a second jet airflow to penetrate into the combustion chamber.

[0080] The turbine engine according to any preceding clause, the plurality of outer baffle holes are configured to allow an airflow to enter into the outer cavity between the outer baffle and the outer liner, and the plurality of inner baffle holes are configured to allow an airflow to enter into the inner cavity between the inner baffle and the inner liner.

[0081] The turbine engine according to any preceding clause, the plurality of outer liner holes being configured to allow a first airflow portion to penetrate into the combustion chamber.

[0082] The turbine engine according to any preceding clause, the plurality of inner liner holes being configured to allow a second airflow portion to penetrate into the combustion chamber.

[0083] The turbine engine according to any preceding clause, the outer baffle including a forward outer baffle portion connected to a forward portion of the outer liner, to define a forward outer cavity therebetween, and an aft outer baffle portion connected to an aft portion of the outer liner, to define an aft outer cavity therebetween.

[0084] The turbine engine according to any preceding clause, the forward outer cavity and the aft outer cavity not being in communication with the one or more outer liner dilution holes.

[0085] The turbine engine according to any preceding clause, the forward outer baffle portion and the aft outer baffle portion being connected to an edge of the one or more outer liner dilution holes.

[0086] The turbine engine according to any preceding clause, the jet airflow passing through the one or more outer liner dilution holes not originating from the forward outer cavity or the aft outer cavity.

[0087] The turbine engine according to any preceding clause, the forward outer baffle portion being connected to a first end of the forward portion of the outer liner and to a second end of the forward portion of the outer liner.

[0088] The turbine engine according to any preceding clause, the aft outer baffle being connected to a first end of the aft portion of the outer liner and to a second end of the aft portion of the outer liner.

[0089] The turbine engine according to any preceding clause, the inner baffle having a forward inner baffle portion that is connected to a forward portion of the inner liner to define a forward inner cavity, and an aft inner baffle portion connected to an aft portion of the inner liner to define an aft inner cavity.

[0090] The turbine engine according to any preceding clause, the forward inner cavity and the aft inner cavity not being in communication with the one or more inner liner dilution holes.

[0091] The turbine engine according to any preceding clause, the forward inner baffle portion and the aft inner baffle portion being connected to an edge of the one or more inner liner dilution holes.

[0092] The turbine engine according to any preceding clause, the jet airflow passing through one or more inner liner dilution holes not originating from the forward inner cavity or the aft inner cavity.

[0093] The turbine engine according to any preceding clause, the forward inner baffle portion being connected to a first end of the forward portion of the inner liner and to a second end of the forward portion of the inner liner.

[0094] The turbine engine according to any preceding clause, the aft inner baffle being connected to a first end of the aft portion of the inner liner and to a second end of the aft portion of the inner liner.

[0095] The turbine engine according to any preceding clause, wherein a diameter of the one or more outer liner dilution holes is greater than a diameter of the plurality of outer liner holes.

[0096] The turbine engine according to any preceding clause, wherein a diameter of the one or more inner liner dilution holes is greater than a diameter of the plurality of inner liner holes.

[0097] 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.

Examples

Embodiment Construction

[0009]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, both the foregoing summary of the present disclosure and the following detailed description are exemplary and intended to provide further explanation without limiting the scope of the disclosure as claimed.

[0010]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.

[0011]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.

[0012]The terms “upstream” and “downstream” refer to the relative direction...

TECHNICAL FIELD

[0001] The present disclosure relates generally to a combustor with baffled liners in a turbine engine.BACKGROUND

[0002] Gas turbine engines generally include a fan and a turbo-engine arranged in flow communication with one another, with the turbo-engine disposed downstream of the fan in a direction of flow through the gas turbine engine. The turbo-engine generally includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. With multi-shaft gas turbine engines, the compressor section can include a high pressure (HP) compressor disposed downstream of a low pressure (LP) compressor, and the turbine section can similarly include a low pressure (LP) turbine disposed downstream of a high pressure (HP) turbine. The combustion section includes a combustor that generates combustion gases discharged into the turbine section. With such a configuration, the HP compressor is coupled with the HP turbine via a high pressure (HP) shaft, and the LP compressor is coupled with the LP turbine via a low pressure (LP) shaft.BRIEF DESCRIPTION OF THE DRAWINGS

[0003] Features and advantages will be apparent from the following, more particular, description of various embodiments, as illustrated in the accompanying drawings, wherein like reference numbers generally indicate identical, functionally similar, or structurally similar elements, or both.

[0004] FIG. 1 is a schematic, cross-sectional diagram of a turbine engine, taken along a longitudinal centerline axis of the turbine engine, according to an embodiment of the present disclosure.

[0005] FIG. 2 is a schematic, cross-sectional view of a combustion section of the turbine engine of FIG. 1, according to an embodiment of the present disclosure.

[0006] FIG. 3 is a schematic, cross-sectional view of a combustor that can be used in the combustor section of the turbine engine of FIG. 1, according to another embodiment of the present disclosure.

[0007] FIG. 4 is a schematic, cross-sectional view of a combustor that can be used in the combustor section of the turbine engine of FIG. 1, according to yet another embodiment of the present disclosure.

[0008] FIG. 5 is an elevational view of a portion of an outer liner of the combustor shown in FIG. 3, the outer liner having a plurality of outer liner holes and one or more outer liner dilution holes, according to an embodiment of the present disclosure.DETAILED DESCRIPTION

[0009] 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, both the foregoing summary of the present disclosure and the following detailed description are exemplary and intended to provide further explanation without limiting the scope of the disclosure as claimed.

[0010] 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.

[0011] 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.

[0012] 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.

[0013] The terms “low” and “high,” or their respective comparative degrees (e.g., “lower” and “higher,” where applicable), when used with the compressor, turbine, shaft, or spool components, each refers to relative pressures or relative speeds, or both within an engine unless otherwise specified. For example, a “low-speed” component defines a component configured to operate at a rotational speed, such as a maximum allowable rotational speed, which is lower than that of a “high-speed” component of the engine. Alternatively, unless otherwise specified, the aforementioned terms may be understood in their superlative degree. For example, a “low-pressure turbine” may refer to the lowest maximum pressure within a turbine section, and a “high-pressure turbine” may refer to the highest maximum pressure within the turbine section. The terms “low” or “high” in such aforementioned regards may additionally, or alternatively, be understood as relative to minimum allowable speeds or pressures, or minimum or maximum allowable speeds or pressures relative to normal, desired, steady state, etc., operation of the engine, or any combination thereof.

[0014] 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.

[0015] The singular forms “a,”“an,” and “the” include plural references unless the context clearly dictates otherwise.

[0016] As used herein, the term “axial” refers to directions and orientations that extend substantially parallel to a longitudinal centerline of the turbine engine. Moreover, the terms “radial” and “radially” refer to directions and orientations that extend substantially perpendicular to the longitudinal centerline of the turbine engine. In addition, as used herein, the term “circumferentially” refers to directions and orientations that extend arcuately about the longitudinal centerline of the turbine engine.

[0017] For a rich burn combustor, there is a desire to increase thermal capability of metallic liners or additive liners by adding a baffle, but allowing a full air pressure drop through a dilution hole to maintain the jet penetration and subsequent emissions characteristics. Conventional techniques use individual ‘tower and cap’ dilutions holes. However, this may be cumbersome and expensive to implement on one hundred or more dilution holes. In an embodiment, in order to achieve desired pressure characteristics, a forward baffle and an aft baffle are provided. The use of the forward baffle and the aft baffle allows for full-air pressure drop dilution annularly and, thus, avoid using individual tower / cap assemblies.

[0018] In an embodiment, a combustor includes inner and outer baffles connected to inner and outer liners, respectively. In an embodiment, the combustor includes two inner baffles and two outer baffles connected to the inner and outer liner, respectively. For example, a first baffle (forward baffle) is connected to a forward part of a liner (inner liner or outer liner) and a second baffle (an aft baffle) is connected to an aft part of the liner (inner liner or outer liner).

[0019] The use of the forward baffle and the aft baffle provides the ability to maintain double air pressure drop characteristics to the cooling liners (durability), while allowing for a full air pressure drop to the dilution holes (emissions). This increases the durability of metallic or additively enabled combustor liners (inner liner and outer liner).

[0020] Referring now to the drawings, FIG. 1 is a schematic, cross-sectional diagram 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.

[0021] The turbo-engine 16 depicted in FIG. 1 generally includes an outer casing 18 that is substantially tubular and defines an annular inlet 20. As schematically shown in FIG. 1, the outer casing 18 encases, in serial flow relationship, a compressor section 21 including a booster or a low-pressure (LP) compressor 22 followed downstream by a high-pressure (HP) compressor 24, a combustion section 26, a turbine section 27 including a high-pressure (HP) turbine 28 followed downstream by a low-pressure (LP) turbine 30, and a jet exhaust nozzle section 32. A high-pressure (HP) shaft 34 or a spool drivingly connects the HP turbine 28 to the HP compressor 24, to rotate the HP turbine 28 and the HP compressor 24 in unison. A low-pressure (LP) shaft 36 or a spool drivingly connects the LP turbine 30 to the LP compressor 22, to rotate the LP turbine 30 and the LP compressor 22 in unison. The compressor section 21, the combustion section 26, the turbine section 27, and the jet exhaust nozzle section 32 together define a core air flowpath.

[0022] 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 (airfoils) 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. Each fan blade 40 is 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. The gearbox assembly 46 is shown schematically in FIG. 1. The gearbox assembly 46 includes a plurality of gears for adjusting the rotational speed of the fan shaft 45 and, thus, the fan 38 relative to the LP shaft 36.

[0023] Referring still to the embodiment of FIG. 1, the disk 42 is covered by a rotatable fan hub 48 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 or at least a portion of the turbo-engine 16, or both. The nacelle 50 is supported relative to the turbo-engine 16 by a plurality of circumferentially spaced outlet guide vanes 52 (stator airfoils). Moreover, a downstream section 54 of the nacelle 50 extends over an outer portion of the turbo-engine 16 to define a bypass airflow passage 56 therebetween.

[0024] 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, or both. As the volume of air 58 passes across the fan blades 40 (rotating airfoils), a first portion of air 62 is directed or routed into the bypass airflow passage 56, and a second portion of air 64 is directed or is routed into the upstream section of the core air flowpath, or, more specifically, into the annular inlet 20 of the LP compressor 22. The ratio between the first portion of air 62 and the second portion of air 64 is commonly known as a bypass ratio. The pressure of the second portion of air 64 is then increased, forming compressed air 65, and the compressed air 65 is routed through the HP compressor 24 and into the combustion section 26, where the compressed air 65 is mixed with fuel and burned to generate combustion gases 66. The fuel can be a kerosene-based fuel, such as Jet A-1, or a hydrogen fuel. The fuel can be transported from a fuel reservoir 79 using a fuel pipe 79A to the combustion section 26.

[0025] 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, or both, from the combustion gases 66 is extracted via sequential stages of HP turbine stator vanes 68 (stator airfoils) that are coupled to the outer casing 18 and HP turbine rotor blades 70 (rotating airfoils) that are coupled to the HP shaft 34, thus, causing the HP shaft 34 to rotate, thereby supporting operation of the HP compressor 24. 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 and the kinetic energy is extracted from the combustion gases 66 via sequential stages of LP turbine stator vanes 72 (stator airfoils) that are coupled to the outer casing 18 and LP turbine rotor blades 74 (rotating airfoils) that are coupled to the LP shaft 36, thus, causing the LP shaft 36 to rotate, thereby supporting operation of the LP compressor 22 and rotation of the fan 38 via the gearbox assembly 46.

[0026] 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 pressure of the first portion of air 62 is substantially increased as the first portion of 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.

[0027] The turbine engine 10 depicted in FIG. 1 is by way of example only. In other embodiments, the turbine engine 10 may have any other suitable configuration. For example, in other 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. Moreover, in other embodiments, any other suitable number or configuration of compressors, turbines, shafts, or a combination thereof may be provided. In still other 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, or turboshaft engines.

[0028] The turbine engine 10 discussed herein is suitable for use on aircraft. Suitable aircraft include, for example, airplanes and unmanned aerial vehicles (UAV). In other embodiments, the turbine engine can be any other turbine engine, such as an industrial turbine engine incorporated into a power generation system, or a nautical turbine engine on a ship or other vessel.

[0029] FIG. 2 is a schematic, cross-sectional view of the combustion section 26 of the turbine engine 10, shown in FIG. 1, according to an embodiment of the present disclosure. The combustion section 26 generally includes a combustor 80 that generates the combustion gases discharged into the turbine section, or, more particularly, into the HP turbine 28 (shown in FIG. 1). The combustor 80 includes an outer liner 82, an inner liner 84, and a dome 86. The outer liner 82, the inner liner 84, and the dome 86 together define a combustion chamber 88. In addition, a diffuser 90 is positioned upstream of the combustion chamber 88. The diffuser 90 has an outer diffuser wall 90A and an inner diffuser wall 90B. The inner diffuser wall 90B is closer to the longitudinal centerline axis 12. The diffuser 90 receives an air flow from the compressor section 21 (shown in FIG. 1) and provides a flow of compressed air to the combustor 80. In an embodiment, the diffuser 90 provides the flow of compressed air to a single circumferential row of fuel-air mixers 92. However, a plurality of circumferential rows of fuel-air mixer 92 can also be used. In an embodiment, the dome 86 of the combustor 80 is configured as a single annular dome, and the circumferential row of fuel-air mixers 92 are provided within openings formed in the dome 86 (air feeding dome or combustor dome). However, in other embodiments, a multiple annular dome can also be used. In general, other types of combustors can also be used.

[0030] In an embodiment, the diffuser 90 can be used to slow the high speed, highly compressed air from the HP compressor 24 (shown in FIG. 1) to a velocity optimal for the combustor 80. Furthermore, the diffuser 90 can also be configured to limit flow distortion as much as possible by avoiding undesirable flow effects such as boundary layer separation. Similar to most other turbine engine components, the diffuser 90 is generally designed to be as light as possible to reduce weight of the overall engine.

[0031] A fuel nozzle (not shown) provides fuel transported by fuel pipe 79A from fuel reservoir 79 (shown in FIG. 1) to fuel-air mixers 92 depending upon a desired performance of the combustor 80 at various engine operating states. In the embodiment shown in FIG. 2, an outer cowl 94 (e.g., an annular cowl) and an inner cowl 96 (e.g., an annular cowl) are located upstream of the combustion chamber 88 so as to direct air flow into fuel-air mixers 92. The outer cowl 94 and the inner cowl 96 may also direct a portion of the flow of air from the diffuser 90 to an outer passage 98 defined between the outer liner 82 and an outer casing 100, and an inner passage 102 defined between the inner liner 84 and an inner casing 104. In addition, an inner support cone 106 is further shown as being connected to a nozzle support 108 using a plurality of bolts 110 and nuts 112. Other combustion sections, however, may include any other suitable structural configurations.

[0032] The combustor 80 also has an igniter 114. The igniter 114 is provided to ignite the fuel-air mixture supplied to combustion chamber 88 of the combustor 80. The igniter 114 is attached to the outer casing 100 of the combustor 80 in a substantially fixed manner. Additionally, the igniter 114 extends generally along an axial direction A2, defining a distal end 116 that is positioned proximate to an opening in a combustor member of the combustor 80. The distal end 116 is positioned proximate to an opening 118 within the outer liner 82 of the combustor 80.

[0033] In an embodiment, the dome 86 of the combustor 80, together with the outer liner 82, the inner liner 84, and fuel-air mixers 92, form the combustion chamber 88 and define a swirling flow 130. The air flows through the fuel-air mixers 92 as the air enters the combustion chamber 88. The role of the dome 86 and the fuel-air mixers 92 is to generate turbulence in the air flow to rapidly mix the air with the fuel to create a fuel-air mixture. The fuel-air mixers 92 (also called a swirlers or mixers) establish a local low pressure zone that forces some of the combustion products to recirculate, as illustrated in FIG. 2, creating needed high turbulence.

[0034] The outer liner 82 includes a plurality of outer liner cooling holes 202. The inner liner 84 includes a plurality of inner liner cooling holes 204. The plurality of outer liner cooling holes 202 and the plurality of inner liner cooling holes 204 are provided to cool down the outer liner 82 and the inner liner 84, respectively, as the outer liner 82 and the inner liner 84 can be exposed to the hot environment of the combustion chamber 88.

[0035] FIG. 3 is a schematic, cross-sectional view of a combustor 300 that can be used in the combustor section 26 of the turbine engine 10, shown in FIG. 1, according to another embodiment of the present disclosure. As shown in FIG. 3, the combustor 300 includes an outer liner 302, an inner liner 304, and a dome 306. The outer liner 302, the inner liner 304, and the dome 306 together define a combustion chamber 308.

[0036] In an embodiment, the dome 306 of the combustor 300 is configured as a single annular dome. In other embodiments, multiple annular domes can also be used. In general, other types of combustors can also be used.

[0037] The outer liner 302 includes a plurality of outer liner holes 302A. The inner liner 304 includes a plurality of inner liner holes 304A. The plurality of outer liner holes 302A and the plurality of inner liner holes 304A are configured to provide a partial air pressure drop across the outer liner 302 and the inner liner 304 to cool down the outer liner 302 and the inner liner 304, respectively, as the outer liner 302 and the inner liner 304 can be exposed to the hot environment of the combustion chamber 308. In addition to the plurality of outer liner holes 302A, the outer liner 302 also includes one or more outer liner dilution holes 302B. Similarly, in addition to the plurality of inner liner holes 304A, the inner liner 304 also includes one or more inner liner dilution holes 304B. The one or more outer liner dilution holes 302B and the one or more inner liner dilution holes 304B are configured to provide an additional partial air pressure drop across the outer liner 302 and across the inner liner 304 to enhance jet airflow penetration into the combustion chamber 308. In an embodiment, a diameter of the one or more outer liner dilution holes 302B is greater than a diameter of the plurality of outer liner holes 302A. In an embodiment, a diameter of the one or more inner liner dilution holes 304B is greater than a diameter of the plurality of inner liner holes 304A.

[0038] The combustor 300 also includes an outer baffle 312 and an inner baffle 314. The outer baffle 312 has a plurality of outer baffle holes 312A. In an embodiment, the outer baffle 312 or the inner baffle 314, or both, can be metallic baffles or additive baffles. In another embodiment, the outer baffle 312 or the inner baffle 314, or both can be made from a composite material. The inner baffle has a plurality of inner baffle holes 314A. The outer baffle 312 is connected to the outer liner 302 and the inner baffle 314 is connected to the inner liner 304 to define, respectively, an outer cavity 316 between the outer baffle 312 that is connected to the outer liner 302, and an inner cavity 318 between the inner baffle 314 and the inner liner 304.

[0039] In operation, airflow 322 penetrates into the outer cavity 316 through the outer baffle holes 312A. Airflow 324 penetrates into the inner cavity 318 through the inner baffle holes 314A. Air is distributed within the outer cavity 316 and the inner cavity 318. A first portion of airflow 326 from the air within the outer cavity 316 penetrates into the combustion chamber 308 through the plurality of outer liner holes 302A. In addition, a first jet airflow 328 from the air within the outer cavity 316 penetrates into the combustion chamber 308 through the one or more outer liner dilution holes 302B. A second portion of airflow 330 from the air within the inner cavity 318 penetrates into the combustion chamber 308 through the plurality of inner liner holes 304A. In addition, a second jet airflow 332 from the air within the inner cavity 318 penetrates into the combustion chamber 308 through the one or more inner liner dilution holes 304B. The first jet airflow 328 and the second jet airflow 332 further enhance jet airflow penetration into the combustion chamber 308. In an embodiment, the one or more outer liner dilution holes 302B and / or the one or more inner liner dilution holes 304B are provided to introduce airflow jets (e.g., the first jet airflow 328 and / or the second jet airflow 332) into the combustion chamber 308 to stop or to “quench” the reaction to reduce or substantially eliminate undesirable emissions of gases, such as NOx.

[0040] FIG. 4 is a schematic, cross-sectional view of a combustor 400 that can be used in the combustor section 26 of the turbine engine 10, shown in FIG. 1, according to yet another embodiment of the present disclosure. The combustor 400 is similar in many aspects to the combustor 300 shown in FIG. 3. Therefore, only distinguishing features will be discussed further in detail herein. Similar to the combustor 300, the combustor 400 also includes an outer baffle 412 and an inner baffle 414. The outer baffle 412 has a plurality of outer baffle holes 412A. The inner baffle has a plurality of inner baffle holes 414A. The outer baffle 412 is connected to the outer liner 302 to define an outer cavity 416 between the outer baffle 412 and the outer liner 302. The inner baffle 414 is connected to the inner liner 304 to define an inner cavity 418 between the inner baffle 314 and the inner liner 304.

[0041] The outer baffle 412 has a forward outer baffle portion 413 that is connected to a forward portion 303 of the outer liner 302 to define a forward outer cavity 416A and an aft outer baffle portion 415 that is connected to an aft portion 305 of the outer liner 302 to define an aft outer cavity 416B. The inner baffle 414 has a forward inner baffle portion 417 that is connected to a forward portion 307 of the inner liner 304 to define a forward inner cavity 418A and an aft inner baffle portion 419 that is connected to an aft portion 309 of the inner liner 304 to define an aft inner cavity 418B.

[0042] As shown in FIG. 4, the forward outer baffle portion 413 is connected to a first end 303A of the forward portion 303 and to a second end 303B of the forward portion 303. The aft outer baffle portion 415 is connected to a first end 305A of the aft portion 305 and to a second end 305B of the aft portion 305. The second end 303B of the forward portion 303 and the second end 305B of the aft portion 305 are located near the one or more outer liner dilution holes 302B. Similarly, the forward inner baffle portion 417 is connected to a first end 307A of the forward portion 307 and to a second end 307B of the forward portion 307. The aft inner baffle portion 419 is connected to a first end 309A of the aft portion 309 and to a second end 309B of the aft portion 309. The second end 307B of the forward portion 307 and the second end 309B of the aft portion 309 are located near the one or more inner liner dilution holes 304B.

[0043] In this embodiment, the forward outer cavity 416A and the aft outer cavity 416B are not in communication with the one or more outer liner dilution holes 302B in the outer liner 302. Similarly, the forward inner cavity 418A and the aft inner cavity 418B are also not in communication with the one or more inner liner dilution holes 304B in the inner liner 304. The forward outer baffle portion 413 and the aft outer baffle portion 415 are connected to an edge of the one or more outer liner dilution holes 302B. The forward inner baffle portion 417 and the aft inner baffle portion 419 are connected to an edge of the one or more inner liner dilution holes 304B. Therefore, the first jet airflow 328 does not originate from within the forward outer cavity 416A or the aft outer cavity 416B. In addition, the second jet airflow 332 does not originate from within the forward inner cavity 418A or the aft inner cavity 418B. On the other hand, the first portion of airflow 326 through the plurality of outer liner holes 302A originates from the forward outer cavity 416A and the aft outer cavity 416B. The second portion of airflow 330 through the plurality of inner liner holes 304A originates from the forward inner cavity 418A and the aft inner cavity 418B.

[0044] Using the forward outer baffle portion 413 and the aft outer baffle portion 415 maintains the air pressure drop through the plurality of outer liner holes 302A and provides a full air pressure drop through the one or more outer liner dilution holes 302B in the outer liner 302, to enhance the outer liner 302 durability. The outer liner 302 can be a metallic liner or an additive liner. For example, the outer liner 302 can be shaped from a metallic material (e.g., a metallic layer) or additively manufactured.

[0045] Using the forward inner baffle portion 417 and the aft inner baffle portion 419 maintains the air pressure drop through the plurality of inner liner holes 304A and provides a full air pressure drop through the one or more inner liner dilution holes 304B in the inner liner 304, to enhance the inner liner 304 durability. The inner liner 304 can be a metallic liner or an additive liner. For example, the inner liner 304 can be shaped from a metallic material (e.g., a metallic layer) or additively manufactured.

[0046] FIG. 5 is an elevational view of a portion of the outer liner 302 of the combustor 300 with the plurality outer liner holes 302A and the one or more outer liner dilution holes 302B, according to an embodiment of the present disclosure. In the embodiment shown in FIG. 5, a plurality of outer liner dilution holes 302B are provided. The plurality of outer liner dilution holes 302B are spaced apart around a circumference of the combustor 300. In an embodiment, the plurality of outer liner dilution holes 302B are equidistantly spaced apart. In an embodiment, the plurality of outer liner holes 302A are disposed adjacent the plurality of outer liner dilution holes 302B along a circumference of the combustor 300. A diameter of the plurality of outer liner holes 302A is less than a diameter of the plurality of outer liner dilution holes 302B. Although only outer liner 302 is depicted in FIG. 5, similar features can also be provided in the inner liner 304.

[0047] In an embodiment, the outer liner 302 or the inner liner 304, or both, can be a metallic layer or an additive liner.

[0048] For a rich burn combustor, such as the combustor 300 and the combustor 400 as shown in FIGS. 3 and 4, it is desirable to increase the thermal capability of the outer liner 302 and the inner liner 304 by providing an outer baffle 312, 412 connected to the outer liner 302 and an inner baffle 314, 414 connected to the inner liner 304. For example, the use of the outer baffle 412 having a forward outer baffle portion 413 and an aft outer baffle portion 415, which are separated by the one or more outer liner dilution holes 302B, provides a full air pressure drop (i.e., a pressure difference between the outer passage and the combustion chamber 308) through the one or more outer liner dilution holes 302B to enhance jet penetration and, thus, overall gas emission characteristics. For example, the use of the inner baffle 414, having a forward inner baffle portion 417, and an aft inner baffle portion 419, which are separated by the one or more inner liner dilution holes 304B, provides a full air pressure drop through the one or more inner liner dilution holes 304B to enhance jet penetration and, thus, overall gas emission characteristics. For example, by providing a full pressure drop, higher jet airflow penetration into the combustion chamber 308 can be achieved to quench the combustion reaction inside the combustion chamber 308 and to halt the production of undesirable emission gases. The use of the forward outer baffle portion 413, the forward inner baffle portion 417, the aft outer baffle portion 415, and the aft inner baffle portion 419 allows for a full-air pressure drop dilution annularly.

[0049] The outer baffle 312, 412 and the inner baffle 314, 414 can be made from the same material as that of the outer liner 302 and the inner liner 304, such as, a superalloy that is capable of withstanding elevated temperatures, high stresses, and often highly oxidizing environments.

[0050] As described above, in an embodiment, the outer liner 302 and the inner liner 304 are provided with, respectively, the outer baffle 312, 412 and the inner baffle 314, 414. The outer baffle 312, 412 is connected to the outer liner 302 to define the outer cavity 316, 416 therebetween. The inner baffle 314, 414 is connected to the inner liner 304 to define the inner cavity 318, 418 therebetween.

[0051] In another embodiment, the outer liner 302 can be provided with the outer baffle 312, 412 while the inner liner 304 is not provided with the inner baffle 314, 414. In yet another embodiment, the inner liner 304 can be provided with the inner baffle 314, 414 while the outer liner 302 is not provided with the outer baffle 312, 412.

[0052] Further aspects are provided by the subject matter of the following clauses.

[0053] A combustor for a turbine engine includes a dome, an outer liner including a plurality of outer liner holes and one or more outer liner dilution holes, an inner liner including a plurality of inner liner holes and one or more inner liner dilution holes, the dome, the outer liner and the inner liner defining a combustion chamber, an outer baffle connected to the outer liner to define an outer cavity therebetween, the outer baffle including a plurality of outer baffle holes, and an inner baffle connected to the inner liner to define an inner cavity therebetween, the inner baffle including a plurality of inner baffle holes, the plurality of outer liner holes being configured to provide a partial air pressure drop across the outer liner and the plurality of inner liner holes are configured to provide a partial air pressure drop across the inner liner to cool down the outer liner and the inner liner, and the one or more outer liner dilution holes being configured to provide an additional pressure drop across the outer liner and the one or more inner liner dilution holes are configured to provide an additional air pressure drop across the inner liner to enhance penetration of a jet airflow into the combustion chamber.

[0054] The combustor according to the preceding clause, the inner liner, the outer liner, or the inner liner and the outer liner both being metallic liners or additive liners.

[0055] The combustor according to any preceding clause, the inner baffle, the outer baffle, or both being metallic baffles or additive baffles.

[0056] The combustor according to any preceding clause, the one or more outer liner dilution holes being configured to allow a first jet airflow to penetrate into the combustion chamber.

[0057] The combustor according to any preceding clause, the one or more inner liner dilution holes being configured to allow a second jet airflow to penetrate into the combustion chamber.

[0058] The combustor according to any preceding clause, the plurality of outer baffle holes are configured to allow an airflow to enter into the outer cavity between the outer baffle and the outer liner, and the plurality of inner baffle holes are configured to allow an airflow to enter into the inner cavity between the inner baffle and the inner liner.

[0059] The combustor according to any preceding clause, the plurality of outer liner holes being configured to allow a first airflow portion to penetrate into the combustion chamber.

[0060] The combustor according to any preceding clause, the plurality of inner liner holes being configured to allow a second airflow portion to penetrate into the combustion chamber.

[0061] The combustor according to any preceding clause, the outer baffle including a forward outer baffle portion connected to a forward portion of the outer liner, to define a forward outer cavity therebetween, and an aft outer baffle portion connected to an aft portion of the outer liner, to define an aft outer cavity therebetween.

[0062] The combustor according to any preceding clause, the forward outer cavity and the aft outer cavity not being in communication with the one or more outer liner dilution holes.

[0063] The combustor according to any preceding clause, the forward outer baffle portion and the aft outer baffle portion being connected to an edge of the one or more outer liner dilution holes.

[0064] The combustor according to any preceding clause, the jet airflow passing through the one or more outer liner dilution holes not originating from the forward outer cavity or the aft outer cavity.

[0065] The combustor according to any preceding clause, the forward outer baffle portion being connected to a first end of the forward portion of the outer liner and to a second end of the forward portion of the outer liner.

[0066] The combustor according to any preceding clause, the aft outer baffle being connected to a first end of the aft portion of the outer liner and to a second end of the aft portion of the outer liner.

[0067] The combustor according to any preceding clause, the inner baffle having a forward inner baffle portion that is connected to a forward portion of the inner liner to define a forward inner cavity, and an aft inner baffle portion connected to an aft portion of the inner liner to define an aft inner cavity.

[0068] The combustor according to any preceding clause, the forward inner cavity and the aft inner cavity not being in communication with the one or more inner liner dilution holes.

[0069] The combustor according to any preceding clause, the forward inner baffle portion and the aft inner baffle portion being connected to an edge of the one or more inner liner dilution holes.

[0070] The combustor according to any preceding clause, the jet airflow passing through one or more inner liner dilution holes not originating from the forward inner cavity or the aft inner cavity.

[0071] The combustor according to any preceding clause, the forward inner baffle portion being connected to a first end of the forward portion of the inner liner and to a second end of the forward portion of the inner liner.

[0072] The combustor according to any preceding clause, the aft inner baffle being connected to a first end of the aft portion of the inner liner and to a second end of the aft portion of the inner liner.

[0073] The combustor according to any preceding clause, wherein a diameter of the one or more outer liner dilution holes is greater than a diameter of the plurality of outer liner holes.

[0074] The combustor according to any preceding clause, wherein a diameter of the one or more inner liner dilution holes is greater than a diameter of the plurality of inner liner holes.

[0075] A turbine engine includes a combustor. The combustor includes a dome, an outer liner including a plurality of outer liner holes and one or more outer liner dilution holes, an inner liner including a plurality of inner liner holes and one or more inner liner dilution holes, the dome, the outer liner and the inner liner defining a combustion chamber, an outer baffle connected to the outer liner to define an outer cavity therebetween, the outer baffle including a plurality of outer baffle holes, and an inner baffle connected to the inner liner to define an inner cavity therebetween, the inner baffle including a plurality of inner baffle holes, the plurality of outer liner holes being configured to provide a partial air pressure drop across the outer liner and the plurality of inner liner holes are configured to provide a partial air pressure drop across the inner liner to cool down the outer liner and the inner liner, and the one or more outer liner dilution holes being configured to provide an additional pressure drop across the outer liner and the one or more inner liner dilution holes are configured to provide an additional air pressure drop across the inner liner to enhance penetration of a jet airflow into the combustion chamber.

[0076] The turbine engine according to the preceding clause, the inner liner, the outer liner, or the inner liner and the outer liner both being metallic liners or additive liners.

[0077] The turbine engine according to any preceding clause, the inner baffle, the outer baffle, or both being metallic baffles or additive baffles.

[0078] The turbine engine according to any preceding clause, the one or more outer liner dilution holes being configured to allow a first jet airflow to penetrate into the combustion chamber.

[0079] The turbine engine according to any preceding clause, the one or more inner liner dilution holes being configured to allow a second jet airflow to penetrate into the combustion chamber.

[0080] The turbine engine according to any preceding clause, the plurality of outer baffle holes are configured to allow an airflow to enter into the outer cavity between the outer baffle and the outer liner, and the plurality of inner baffle holes are configured to allow an airflow to enter into the inner cavity between the inner baffle and the inner liner.

[0081] The turbine engine according to any preceding clause, the plurality of outer liner holes being configured to allow a first airflow portion to penetrate into the combustion chamber.

[0082] The turbine engine according to any preceding clause, the plurality of inner liner holes being configured to allow a second airflow portion to penetrate into the combustion chamber.

[0083] The turbine engine according to any preceding clause, the outer baffle including a forward outer baffle portion connected to a forward portion of the outer liner, to define a forward outer cavity therebetween, and an aft outer baffle portion connected to an aft portion of the outer liner, to define an aft outer cavity therebetween.

[0084] The turbine engine according to any preceding clause, the forward outer cavity and the aft outer cavity not being in communication with the one or more outer liner dilution holes.

[0085] The turbine engine according to any preceding clause, the forward outer baffle portion and the aft outer baffle portion being connected to an edge of the one or more outer liner dilution holes.

[0086] The turbine engine according to any preceding clause, the jet airflow passing through the one or more outer liner dilution holes not originating from the forward outer cavity or the aft outer cavity.

[0087] The turbine engine according to any preceding clause, the forward outer baffle portion being connected to a first end of the forward portion of the outer liner and to a second end of the forward portion of the outer liner.

[0088] The turbine engine according to any preceding clause, the aft outer baffle being connected to a first end of the aft portion of the outer liner and to a second end of the aft portion of the outer liner.

[0089] The turbine engine according to any preceding clause, the inner baffle having a forward inner baffle portion that is connected to a forward portion of the inner liner to define a forward inner cavity, and an aft inner baffle portion connected to an aft portion of the inner liner to define an aft inner cavity.

[0090] The turbine engine according to any preceding clause, the forward inner cavity and the aft inner cavity not being in communication with the one or more inner liner dilution holes.

[0091] The turbine engine according to any preceding clause, the forward inner baffle portion and the aft inner baffle portion being connected to an edge of the one or more inner liner dilution holes.

[0092] The turbine engine according to any preceding clause, the jet airflow passing through one or more inner liner dilution holes not originating from the forward inner cavity or the aft inner cavity.

[0093] The turbine engine according to any preceding clause, the forward inner baffle portion being connected to a first end of the forward portion of the inner liner and to a second end of the forward portion of the inner liner.

[0094] The turbine engine according to any preceding clause, the aft inner baffle being connected to a first end of the aft portion of the inner liner and to a second end of the aft portion of the inner liner.

[0095] The turbine engine according to any preceding clause, wherein a diameter of the one or more outer liner dilution holes is greater than a diameter of the plurality of outer liner holes.

[0096] The turbine engine according to any preceding clause, wherein a diameter of the one or more inner liner dilution holes is greater than a diameter of the plurality of inner liner holes.

[0097] 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.

Examples

Embodiment Construction

[0009]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, both the foregoing summary of the present disclosure and the following detailed description are exemplary and intended to provide further explanation without limiting the scope of the disclosure as claimed.

[0010]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.

[0011]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.

[0012]The terms “upstream” and “downstream” refer to the relative direction...

Claims

1. A combustor for a turbine engine, the combustor comprising:a dome;an outer liner comprising a plurality of outer liner holes and one or more outer liner dilution holes;an inner liner comprising a plurality of inner liner holes and one or more inner liner dilution holes, the dome, the outer liner and the inner liner defining a combustion chamber;an outer baffle connected to the outer liner to define an outer cavity therebetween, the outer baffle comprising a plurality of outer baffle holes; andan inner baffle connected to the inner liner to define an inner cavity therebetween, the inner baffle comprising a plurality of inner baffle holes,wherein the plurality of outer liner holes are configured to provide a partial air pressure drop across the outer liner and the plurality of inner liner holes are configured to provide a partial air pressure drop across the inner liner to cool down the outer liner and the inner liner,wherein the one or more outer liner dilution holes are in fluid communication with the outer cavity, and the one or more outer liner dilution holes are configured to provide an additional pressure drop across the outer liner, andwherein the one or more inner liner dilution holes are in fluid communication with the inner cavity, and the one or more inner liner dilution holes are configured to provide an additional air pressure drop across the inner liner to enhance penetration of a jet airflow into the combustion chamber, wherein a diameter of the one or more outer liner dilution holes is greater than a diameter of the plurality of outer liner holes and a diameter of the one or more inner liner dilution holes is greater than a diameter of the plurality of inner liner holes.

2. The combustor according to claim 1, wherein the inner liner, the outer liner, or both are metallic liners or additive liners.

3. The combustor according to claim 1, wherein the inner baffle, the outer baffle, or both are metallic baffles or additive baffles.

4. The combustor according to claim 1, wherein the one or more outer liner dilution holes are configured to allow a first jet airflow to penetrate into the combustion chamber.

5. The combustor according to claim 4, wherein the one or more inner liner dilution holes are configured to allow a second jet airflow to penetrate into the combustion chamber.

6. The combustor according to claim 1, wherein the plurality of outer baffle holes are configured to allow an airflow to enter into the outer cavity between the outer baffle and the outer liner, and the plurality of inner baffle holes are configured to allow an airflow to enter into the inner cavity between the inner baffle and the inner liner.

7. The combustor according to claim 6, wherein the plurality of outer liner holes are configured to allow a first airflow portion to penetrate into the combustion chamber.

8. The combustor according to claim 7, wherein the plurality of inner liner holes are configured to allow a second airflow portion to penetrate into the combustion chamber.

9. A combustor for a turbine engine, the combustor comprising:a dome;an outer liner comprising a plurality of outer liner holes and one or more outer liner dilution holes;an inner liner comprising a plurality of inner liner holes and one or more inner liner dilution holes, the dome, the outer liner and the inner liner defining a combustion chamber;an outer baffle connected to the outer liner to define an outer cavity therebetween, the outer baffle comprising a plurality of outer baffle holes; andan inner baffle connected to the inner liner to define an inner cavity therebetween, the inner baffle comprising a plurality of inner baffle holes,wherein the plurality of outer liner holes are configured to provide a partial air pressure drop across the outer liner and the plurality of inner liner holes are configured to provide a partial air pressure drop across the inner liner to cool down the outer liner and the inner liner,wherein the one or more outer liner dilution holes are configured to provide an additional pressure drop across the outer liner, and the one or more inner liner dilution holes are configured to provide an additional air pressure drop across the inner liner to enhance penetration of a jet airflow into the combustion chamber, andwherein the outer baffle comprises a forward outer baffle portion connected to a first end of a forward portion of the outer liner and to a second end of the forward portion of the outer liner, to define a forward outer cavity therebetween, and an aft outer baffle portion connected to a first end of an aft portion of the outer liner and to a second end of the aft portion of the outer liner, to define an aft outer cavity therebetween, the one or more outer liner dilution holes being provided between the forward outer cavity and the aft outer cavity, the forward outer cavity extending in a longitudinal direction from the first end of the forward portion of the outer liner to the second end of the forward portion of the outer liner, and the aft outer cavity extending from the first end of an aft portion of the outer liner to the second end of the aft portion of the outer liner in the longitudinal direction.

10. The combustor according to claim 9, wherein the forward outer cavity and the aft outer cavity are not in communication with the one or more outer liner dilution holes.

11. The combustor according to claim 9, wherein the forward outer baffle portion and the aft outer baffle portion are connected to an edge of the one or more outer liner dilution holes.

12. The combustor according to claim 9, wherein the jet airflow passing through the one or more outer liner dilution holes does not originate from the forward outer cavity or the aft outer cavity.13-14. (canceled)15. The combustor according to claim 9, wherein the inner baffle has a forward inner baffle portion that is connected to a forward portion of the inner liner to define a forward inner cavity, and an aft inner baffle portion connected to an aft portion of the inner liner to define an aft inner cavity.

16. The combustor according to claim 15, wherein the forward inner cavity and the aft inner cavity are not in communication with the one or more inner liner dilution holes.

17. The combustor according to claim 15, wherein the forward inner baffle portion and the aft inner baffle portion are connected to an edge of the one or more inner liner dilution holes.

18. The combustor according to claim 15, wherein the jet airflow passing through one or more inner liner dilution holes does not originate from the forward inner cavity or the aft inner cavity.

19. The combustor according to claim 15, wherein the forward inner baffle portion is connected to a first end of the forward portion of the inner liner and to a second end of the forward portion of the inner liner.

20. The combustor according to claim 19, wherein the aft inner baffle portion is connected to a first end of the aft portion of the inner liner and to a second end of the aft portion of the inner liner.

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