Combustor with reverse dilution air introduction

Abstract

A combustor for a gas turbine engine includes a forward liner segment having an aft end portion. The forward liner segment at least partially defines a primary combustion chamber. The combustor further includes an aft liner segment having a forward end portion. A passage is defined between the forward liner segment and the aft liner segment. During operation of the combustor, the passage directs a flow of dilution air in a counterflow or reverse direction relative to a flow of combustion gases out of the primary combustion chamber.

Description

Technical Field

[0001] This disclosure relates to a gas turbine engine combustor with a counter-current dilution air introduction. Background Technology

[0002] Gas turbine engines, such as turbofan engines, are used for aircraft propulsion. Gas turbine engines typically include a compressor section, a combustion section, and a turbine section. More specifically, the combustion section includes an annular combustor. In some combustor configurations, such as compact combustors, NOx (nitrogen oxides) formation can be reduced by utilizing a combustion method known as rich-quench-lean or RQL. The inventors of this disclosure have found that improved mixing of dilution air with combustion gases exiting the main combustion zone in an RQL combustor would be beneficial in the art. Attached Figure Description

[0003] The complete and enabling disclosure of this disclosure, including its best mode, is set forth in the specification with reference to the accompanying drawings, for those skilled in the art, wherein:

[0004] Figure 1 This is a schematic cross-sectional view of a gas turbine engine according to an exemplary embodiment of the present disclosure.

[0005] Figure 2 This is a cross-sectional side view of the combustion section of a gas turbine engine according to an exemplary embodiment of the present disclosure.

[0006] Figure 3 This is a cross-sectional side view of the combustion section of a gas turbine engine according to an exemplary embodiment of the present disclosure.

[0007] Figure 4 This is a cross-sectional side view of the combustion section of a gas turbine engine according to an exemplary embodiment of the present disclosure.

[0008] Figure 5 This is an enlarged cross-sectional side view of the combustion section of a gas turbine engine according to an exemplary embodiment of the present disclosure.

[0009] Figure 6 This is an enlarged cross-sectional side view of a portion of the combustion section of a gas turbine engine according to an exemplary embodiment of the present disclosure.

[0010] The repeated use of reference numerals in this specification and drawings is intended to indicate the same or similar features or elements of this disclosure. Detailed Implementation

[0011] Reference will now be made in detail to exemplary embodiments of the subject matter currently disclosed, one or more examples of which are illustrated in the accompanying drawings. Each example is provided by way of explanation and should not be construed as limiting this disclosure. Indeed, it will be apparent to those skilled in the art that various modifications and variations can be made to this disclosure without departing from the scope or spirit of this disclosure. For example, features shown or described as part of one embodiment may be used with another embodiment to produce yet another embodiment. Therefore, this disclosure is intended to cover such modifications and variations that fall within the scope of the appended claims and their equivalents.

[0012] As used herein, the terms “first,” “second,” and “third” are used interchangeably to distinguish one component from another and are not intended to indicate the location or importance of the individual components. Furthermore, the terms “upstream” and “downstream” refer to the relative direction of fluid flow within a fluid path. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction in which the fluid flows.

[0013] The term “exemplary” as used herein means “serving as an example, instance, or illustration.” Any implementation described herein as “exemplary” is not necessarily to be construed as preferred or advantageous over other implementations. Furthermore, all embodiments described herein should be considered exemplary unless otherwise expressly stated. The singular forms “a,” “an,” and “the” include plural references unless the context clearly indicates otherwise. In the context of, for example, “at least one of A, B, and C,” the term “at least one” means only A, only B, only C, or any combination of A, B, and C.

[0014] This disclosure generally relates to a combustor liner design for improving emission reduction. A compact combustor configuration known as a rich-quench-lean (RQL) combustor is used in the gas turbine industry, particularly in the aircraft gas turbine industry, to reduce NOx (nitrogen oxides) emission formation. In an RQL combustor, a fuel-air mixture rich in fuel is supplied to the primary combustion chamber for combustion. Because the fuel-air mixture is fuel-rich, not all fuel is burned in the primary combustion chamber. To burn the remaining fuel in the combustion gases, cooler dilution air is introduced into the combustion gas stream. This dilution air rapidly cools (quenches) the combustion gases, thereby reducing NOx formation, and mixes with these gases to add additional oxygen, thus supplying a lean fuel-air mixture to the secondary combustion chamber to rapidly complete the combustion process, thereby further reducing NOx (nitrogen oxides) and other unwanted emissions. While RQL combustors are useful for emission reduction, further reductions in emissions, particularly NOx, are needed.

[0015] The combustion liner design disclosed herein provides a novel architecture for compact RQL burners, improving operability and NOx control. In at least one embodiment, the burner includes a front liner section defining a primary combustion chamber and a rear liner section defining a secondary combustion chamber downstream of the primary combustion chamber. The primary combustion chamber has a relatively larger volume than the secondary combustion chamber. The larger primary combustion chamber volume provides improved operability, while the smaller secondary combustion chamber accelerates flow for faster mixing / quenching. The channel defined between the front and rear liner sections is oriented to provide a dilution air flow that is opposite or nearly opposite to the combustion gas flow exiting the primary combustion chamber. This channel orientation / counterflow of the dilution flow results in greater turbulence within the combustion gas upstream of the secondary combustion chamber, leading to a more complete / thorough mixing of the dilution air and combustion gas. This effect can reduce NOx more significantly compared to known RQL burners.

[0016] Now refer to the attached diagram, Figure 1 This is a schematic cross-sectional view of one embodiment of the gas turbine engine 10. Figure 1 As shown, the gas turbine engine 10 defines a longitudinal direction L, a radial direction R, and a circumferential direction C. The longitudinal direction L extends parallel to the longitudinal centerline 12 of the gas turbine engine 10, the radial direction R extends orthogonally outward from the longitudinal centerline 12, and the circumferential direction C extends approximately concentrically around the longitudinal centerline 12.

[0017] Typically, the gas turbine engine 10 includes a fan 14, a low-pressure (LP) spool 16, and a high-pressure (HP) spool 18, which are at least partially surrounded by an annular nacelle 20. More specifically, the fan 14 includes a fan rotor 22 and a plurality of fan blades 24 (one shown) coupled to the fan rotor 22. In this respect, the fan blades 24 are spaced apart from each other along the circumferential direction C and extend outward from the fan rotor 22 along the radial direction R. Furthermore, the LP spool 16 and the HP spool 18 are positioned downstream of the fan 14 along a longitudinal centerline 12 (i.e., in the longitudinal direction L). As shown, the LP spool 16 is rotatably coupled to the fan rotor 22, thereby allowing the LP rotor 16 to rotate the fan 14. In addition, a plurality of outlet guide vanes or struts 26, spaced apart from each other in the circumferential direction C, extend along the radial direction R between the outer casing 28 and the nacelle 20 surrounding the LP spool 16 and the HP spool 18. Thus, the strut 26 supports the cabin 20 relative to the outer shell 28, such that the outer shell 28 and the cabin 20 define a bypass airflow passage 30 positioned between them.

[0018] The housing 28 typically surrounds or encloses the compressor section 32, combustion section 34, turbine section 36, and exhaust section 38 in a series flow sequence. The compressor section 32 may include a low-pressure (LP) compressor 40 on an LP spool 16 and a high-pressure (HP) compressor 42 on an HP spool 18, the high-pressure compressor 42 being positioned downstream of the LP compressor 40 along a longitudinal centerline 12. Each compressor 40, 42 may further include one or more rows of stator blades 44 intersecting with one or more rows of compressor blades 46. Furthermore, in some embodiments, the turbine section 36 includes a high-pressure (HP) turbine 48 on an HP spool 18 and a low-pressure (LP) turbine 50 on an LP spool 16, the low-pressure turbine 50 being positioned downstream of the HP turbine 48 along a longitudinal centerline 12. Each turbine 48, 50 may further include one or more rows of stator blades 52 intersecting with one or more rows of turbine rotor blades 54. In a particular embodiment, the turbine section 36 includes a first stator blade or turbine nozzle 52, which is positioned downstream of the combustion section 34 and upstream of the turbine rotor blades 54.

[0019] Furthermore, the LP spool 16 includes a low-pressure (LP) spool 56 and the HP spool 18 includes a high-pressure (HP) spool 58 concentrically positioned around the LP spool 56. In such an embodiment, the HP spool 58 is rotatably coupled to the rotor blades 54 of the HP turbine 48 and the rotor blades 46 of the HP compressor 42, such that rotation of the HP turbine rotor blades 54 rotatably drives the HP compressor rotor blades 46. As shown, the LP spool 56 is directly coupled to the rotor blades 54 of the LP turbine 50 and the rotor blades 46 of the LP compressor 40. Furthermore, the LP spool 56 is coupled to the fan 14 via a gearbox 60. In this respect, rotation of the LP turbine rotor blades 54 rotatably drives the LP compressor rotor blades 46 and the fan blades 24.

[0020] In some embodiments, the gas turbine engine 10 can generate thrust to propel the aircraft. More specifically, during operation, air 62 enters the inlet section 64 of the gas turbine engine 10. As the air 62 flows through the fan 14, it is split into bypass air 66 (indicated by arrow 66) and compressor air 68 (indicated by arrow 68). The bypass air 66 is directed through the bypass airflow passage 30. The compressor air 68 is directed to the inlet 70 of the LP compressor 40, where the rotor blades 46 progressively compress the compressor air 68. The compressor air 68 is then directed to the HP compressor, where the rotor blades 46 continue to progressively compress the compressor air 68. The compressor air 68 is then delivered to the combustion section 34. A portion of the compressor air 68 can be extracted from the HP compressor 42 for cooling and / or other operational purposes.

[0021] In combustion section 34, compressed compressor air 68 is mixed with fuel and burned to produce high-temperature, high-pressure combustion gas 72. The combustion gas 72 then flows through HP turbine 48, where HP turbine rotor blades 54 extract a first portion of its kinetic and / or thermal energy. This energy extraction causes HP shaft 58 to rotate, thereby driving HP compressor 42. The combustion gas 72 then flows through LP turbine 50, where LP turbine rotor blades 54 extract a second portion of its kinetic and / or thermal energy. This energy extraction causes LP shaft 56 to rotate, thereby driving LP compressor 40 and fan 14 via gearbox 60. The combustion gas 72 then exits the gas turbine engine 10 through exhaust section 38.

[0022] Provide the above and Figure 1 The configuration of the gas turbine engine 10 shown is only intended to place the subject matter within an exemplary field of application. Therefore, the subject matter can be readily applied to any configuration of gas turbine engine, including other types of aerospace-based gas turbine engines, sea-based gas turbine engines, and / or land-based / industrial gas turbine engines.

[0023] Figure 2 This is a cross-sectional side view of the combustion section 34 of a gas turbine engine 10 according to an exemplary embodiment of the present disclosure. Figure 2 As shown, combustion section 34 includes an annular burner 100. The annular burner 100 further includes a front liner section 102 having a front inner liner 104 and a front outer liner 106. The front outer liner 106 is radially spaced from the front inner liner 104 in the radial direction R. A first or primary combustion chamber 108 is defined between the front inner liner 104 and the front outer liner 106.

[0024] It should be understood that Figure 1 The exemplary gas turbine engine 10 depicted herein is provided by way of example only, and in other exemplary embodiments, the gas turbine engine 10 may have other configurations. For example, in other exemplary embodiments, aspects of this disclosure may (as appropriate) be incorporated into, for example, a turboprop gas turbine engine, a turboshaft gas turbine engine, or a turbojet gas turbine engine. Various features of the combustor 100 disclosed herein may be incorporated into other combustion systems or configurations, such as a counter-current RQL combustor.

[0025] The burner 100 further includes a rear liner section 110 formed by a rear inner liner 112 and a rear outer liner 114. The rear inner liner 112 and the rear outer liner 114 are radially spaced apart in the radial direction R. A secondary combustion chamber 116 is defined between the rear inner liner 112 and the rear outer liner 114. The rear liner section 110 is positioned downstream of the front liner section 102 relative to the direction in which the combustion gases 72 flow through the burner 100.

[0026] like Figure 2As shown, the burner 100 includes one or more fuel nozzles 118. Although Figure 2 A single fuel nozzle 118 is shown, but the combustion section 34 typically includes a plurality of fuel nozzles 118 spaced circumferentially in a ring array around the longitudinal centerline 12 of the gas turbine engine 10. In a particular embodiment, the combustor 100 includes a swirler or impeller 120 disposed upstream of the primary combustion chamber 108.

[0027] The compressor discharge housing 122 at least partially forms the compressor discharge chamber 124. The compressor discharge housing 122 at least partially surrounds or otherwise encloses the annular burner 100 in the circumferential direction C. The annular burner 100 is in fluid communication with the compressor discharge chamber 124. One or more guide vanes 126 and a diffuser may be used to guide a stream of compressed compressor air 68 from the HP compressor 42 into the compressor discharge chamber 124.

[0028] In various embodiments, the front liner section 102 includes at least one cooling hole or orifice 128 in fluid communication with the compressor exhaust chamber 124. Additionally or alternatively, the rear liner section 110 includes at least one cooling hole or orifice 130 in fluid communication with the compressor exhaust chamber 124. The cooling orifice 128 allows a second portion of the compressor air 68 (referred to herein as cooling air and indicated by arrow 132) to pass through the respective front inner liner 104 or front outer liner 106 and enter the primary combustion chamber 108 during operation. The air may form a cooling boundary layer along the inner surfaces of the respective front inner liner 104 and front outer liner 106. Additionally or alternatively, the cooling orifice 130 may allow cooling air 132 to pass through the respective rear inner liner 112 and / or rear outer liner 114 to form a cooling boundary layer along the inner surfaces of the respective rear inner liner 112 and rear outer liner 114 during operation.

[0029] In various embodiments of this disclosure, the rear end portion 134 of the front liner section 102 overlaps with the front end portion 136 of the rear liner section 110, thereby forming an inner gap or channel 138 and an outer gap or channel 140 therebetween. In some embodiments, the front end portion 136 of the rear liner section 110 is at least partially disposed within the rear end portion 134 of the front liner section 102, thereby forming an inner channel 138 and an outer channel 140 therebetween. In some embodiments, such as Figure 2 As shown, the inner channel 138 and the outer channel 140 extend in the longitudinal direction L or at least almost parallel to the longitudinal direction L.

[0030] Unlike the cooling orifices 128, 130 that guide the cooling air 132 generally parallel or downstream relative to the combustion gas 142 flow, the inner passage 138 and outer passage 140 are oriented relative to the combustion gas 142 flowing from the primary combustion chamber 108 to the secondary combustion chamber 116, guiding, injecting, or causing a portion of the compressor air 68 (referred to herein as dilution air 144) in a generally longitudinal (counter-current / upstream / opposite flow) direction L. This relative orientation of the dilution air 144 flow relative to the combustion gas 142 facilitates a more complete mixing between the combustion gas 142 and the dilution air 144. Furthermore, this configuration provides a more stable combustion process, where the primary combustion chamber has a larger volume VP and the secondary combustion chamber 116 has a smaller volume VS, resulting in a shorter residence time and thus reducing NOx.

[0031] Figure 3 This is a cross-sectional side view of the combustion section 34 of a gas turbine engine 10 according to an alternative embodiment of the present disclosure. In some embodiments, such as Figure 3 As shown, the rear end portion 134 of the front liner section 102 may include one or more recesses or damping chambers 146, 148. The damping chambers 146, 148 may be formed downstream of the fuel nozzle 118 and upstream of the rear liner section 110, respectively, along the front inner liner 104 and the front outer liner 106. The inner wall 150 of the damping chamber 146 may at least partially define an inner passage 138. The inner wall 152 of the damping chamber 148 may at least partially define an outer passage 140. In some embodiments, one or more inlet orifices 154, 156 may provide fluid communication between the compressor exhaust chamber 124 and the damping chambers 146, 148, respectively.

[0032] In some embodiments, such as Figure 3 As shown, one or more exhaust ports 158, 160 may be defined along inner walls 150, 152, respectively. At least one of the exhaust ports 158, 160 may be formed / angled / oriented radially inward or radially outward relative to the longitudinal centerline 12 of the gas turbine engine 10 to guide a portion / flow of cooling air 132 along the inner walls 150, 152, thereby providing thin-film cooling to the front liner section 102 during operation of the combustor 100. In a particular embodiment, damping chambers 146, 148 act as Helmholtz resonators to suppress thermal and / or acoustic oscillations emitted from the combustor 100 during operation.

[0033] Figure 4 Alternative embodiments according to this disclosure Figure 1 A cross-sectional side view of the combustion section 34 of the gas turbine engine 10 is shown. In some embodiments, such as... Figure 4As shown, the front end portion 136 of the rear liner section 110 may include one or more recesses or damping chambers 162, 164. This could be Figure 3 The damping chambers 146, 148 of the front liner section 102 shown are supplementary or alternatives. Damping chambers 162, 164 may be formed downstream of the primary combustion chamber 108 along the rear inner liner 112 and the rear outer liner 114, respectively. The outer wall 166 of the damping chamber 162 may at least partially define the inner passage 138. The outer wall 168 of the damping chamber 164 may at least partially define the outer passage 140. In some embodiments, one or more inlet orifices 170, 172 may provide fluid communication between the compressor exhaust chamber 124 and the damping chambers 162, 164, respectively.

[0034] In some embodiments, one or more exhaust ports 174 may be defined along the inner wall 176 of the rear inner liner 112. Additionally or alternatively, one or more exhaust ports 178 may be defined along the inner wall 180 of the rear outer liner 114. The exhaust ports 174, 178 may be radially inward or radially outward oriented relative to the longitudinal centerline 12 of the gas turbine engine 10 to guide a portion / flow of cooling air 132 along the inner walls 176, 180, thereby providing thin-film cooling to the rear liner section 110 during operation of the combustor 100. In certain embodiments, damping chambers 162, 164 act as Helmholtz resonators to suppress thermal and / or acoustic oscillations emanating from the combustor 100 during operation. In some embodiments, damping chambers 162, 164 may act as mechanical housing reinforcements to reduce thermal distortion / deformation.

[0035] Figure 5 This is based on exemplary embodiments of the present disclosure. Figure 3 An enlarged cross-sectional side view of a portion of the combustor 100 of the gas turbine engine 10 shown. Figure 5 As shown, a preferred radial clearance distance GD can be maintained between the rear end portion 134 of the front liner section 102 and the front end portion 136 of the rear liner section 110 via mechanical spacers or inserts 182, 184 disposed in the respective inner channel 138 and outer channel 140.

[0036] Inserts 182, 184 may be circumferentially spaced along direction C at a specific distance to meter the flow rate of dilution air 144 flowing through the respective inner channel 138 and outer channel 140 and entering the combustion gas 142 stream. Inserts 182, 184 may be configured to allow axial and / or radial relative movement and thermal growth between the rear end portion 134 of the front liner section 102 and the front end portion 136 of the rear liner section 110. Inserts 182, 184 may be rigidly connected to the front liner section 102 and / or the rear liner section 110, or slidably engaged with the front liner section 102 and / or the rear liner section 110. Inserts 182, 184 may include or at least partially define orifices 186, 188. The orifices may be sized to meter the flow rate of dilution air 144 flowing through the respective inner channel 138 and outer channel 140 and entering the combustion gas 142 stream.

[0037] In various embodiments, such as Figure 2 , 3 As shown together with 4 and 5, the inlet section or opening 190 of the rear liner section 110 includes a protrusion or fence 192. More specifically, the fence 192 extends radially into the combustion gas 142 flow relative to the radial direction R, thereby narrowing the opening 190 leading to the secondary combustion chamber 116. Thus, as the combustion gas 142 flows from the primary combustion chamber 108 into the secondary combustion chamber 116, the fence 192 increases the turbulence of the combustion gas 142, thereby promoting faster and more uniform mixing of the dilution air 144.

[0038] Figure 6 Alternative embodiments according to this disclosure Figure 1 An enlarged cross-sectional side view of a portion of the combustor 100 of the gas turbine engine 10 is shown. In one embodiment, at least one of the front inner liner 104 and the front outer liner 106 includes / defines at least one front discrete dilution orifice 194, 196. The front discrete dilution orifice 194, 196 is in fluid communication with the compressor exhaust chamber 124 and the primary combustion chamber 108 and provides annular strips of dilution air 144, which can also be used to cool or quench the combustion gas 142 upstream of the secondary combustion chamber 116.

[0039] In operation, such as Figure 2 , Figure 3 and Figure 4As shown, a portion of the compressor air 68 from the compressor exhaust chamber 124 is directed through and / or across the impeller 120 of the fuel nozzle 118 to impart vortices or turbulence to the compressor air 68. This turbulence or vortex enhances mixing with fuel (as indicated by arrow 198) from the corresponding fuel nozzle 118 upstream of the primary combustion chamber 108. The fuel-rich fuel-air mixture 200 flows downstream from the corresponding fuel nozzle 118 for combustion in the primary combustion chamber 108. Because the fuel-air mixture 200 is fuel-rich, incomplete or partial combustion of the fuel-air mixture 200 occurs in the primary combustion chamber 108. A portion of the compressor air 68 is directed through cooling orifices 128, 130 as cooling air 132 and forms a thermal boundary layer along the inner surfaces of the front inner liner 104 and the front outer liner 106 for thin-film cooling.

[0040] like Figure 5 As shown, dilution air 144 flows through inner channel 138 and outer channel 140 via inserts 182, 184, and then mixes with and quenches / cools combustion gas 142 flowing from primary combustion chamber 108 upstream of secondary combustion chamber 116. The orientation of inner channel 138 and outer channel 140 causes the flow of dilution air 144 to be opposite to or in the opposite direction to the flow of combustion gas 142 within primary combustion chamber 108. This relative orientation of inner channel 138 and outer channel 140, and of dilution air 144, relative to combustion gas 142, is more conducive to a more complete mixing of combustion gas 142 and dilution air 144 compared to known dilution air injection methods (which typically direct the dilution air flow perpendicular to or in the same downstream flow direction as the combustion gas 142 flowing through primary combustion chamber 108 toward secondary combustion chamber 116). Furthermore, compared to the smaller volume VS of the secondary combustion chamber 116, this configuration allows the primary combustion chamber 108 to have a larger volume VP, which results in a more stable combustion process, especially during ground and high-altitude startup, and the smaller volume VS of the secondary combustion chamber 116 or zone results in a shorter residence time, thereby reducing NOx emissions.

[0041] like Figure 5As shown, the inner channel 138 and the outer channel 140 can be positioned at angles 202 and 204, respectively, relative to the axial centerline 206 of the burner 100. The angle 202 can range from 1 degree to 90 degrees. The angle 204 relative to the axial centerline 206 can range from -1 to -90 degrees. In a particular embodiment, the angle 202 can range between 1 degree and 80 degrees. In a particular embodiment, the angle 202 can range between 1 degree and 70 degrees. In a particular embodiment, the angle 202 can range between 1 degree and 60 degrees. In a particular embodiment, the angle 202 can range between 1 degree and 50 degrees. In a particular embodiment, the angle 202 can range between 1 degree and 45 degrees. In a particular embodiment, the angle 204 can range between -1 degree and -80 degrees. In a particular embodiment, the angle 204 can range between -1 degree and -70 degrees. In a particular embodiment, the angle 204 can range between -1 degree and -60 degrees. In a particular embodiment, the angle 204 can range between -1 degree and -50 degrees. In a particular embodiment, the angle 204 can range between -1 degree and 45 degrees.

[0042] The combination of mixing dilution air 144 with combustion gas 142 and rapidly quenching combustion gas 142 results in a significant reduction in NOx formation. In addition to quenching, dilution air 144 also provides additional oxygen to combustion gas 142 to mix with unburned fuel therein. The mixing of dilution air 144 results in a lean fuel-air mixture 208 (e.g., Figure 2 , Figure 3 and Figure 4 (As shown in the diagram) The combustion process is completed by the flow from the primary combustion chamber 108 into the secondary combustion chamber 116.

[0043] In some embodiments, such as Figure 3 As shown, a portion of the compressed compressor air 68 flows from the compressor discharge chamber 124 through inlet orifices 154, 156 and into the corresponding damping chambers 146, 148. This pressurizes the damping chambers 146, 148 and provides cooling to the relevant portions of the front inner liner 104 and the front outer liner 106. The compressed compressor air 68 can then flow through one or more discharge orifices 158, 160 to provide thin-film cooling to the front inner liner 104 and the front outer liner 106, respectively.

[0044] In some embodiments, such as Figure 3As shown, a portion of the compressed compressor air 68 flows from the compressor exhaust chamber 124 through inlet orifices 154, 156 and into the corresponding damping chambers 146, 148 of the front liner section 102. This pressurizes the corresponding damping chambers 146, 148 and provides thin-film cooling to the relevant portions of the front inner liner 104 and the front outer liner 106. The compressed compressor air 68 can then flow through one or more exhaust orifices 158, 160 to provide thin-film cooling to the front inner liner 104 and the front outer liner 106. In some embodiments, the dimensions and / or shapes of the damping chambers 146, 148 and their corresponding exhaust orifices 158, 160 are designed to suppress acoustic and / or thermoacoustic vibrations within the burner 100.

[0045] In some embodiments, such as Figure 4 As shown, a portion of the compressed compressor air 68 flows from the compressor exhaust chamber 124 through inlet orifices 170, 172 and into the corresponding damping chambers 162, 164 of the rear liner section 110. This pressurizes the corresponding damping chambers 162, 164 and provides cooling to the opposing portions of the rear inner liner 112 and the rear outer liner 114. The compressed compressor air 68 can then flow through one or more exhaust orifices 174, 178 to provide thin-film cooling to the rear inner liner 112 and the rear outer liner 114, respectively. In some embodiments, the dimensions and / or shapes of the damping chambers 162, 164 and their corresponding exhaust orifices 174, 178 are designed to suppress acoustic and / or thermoacoustic vibrations within the burner 100.

[0046] Further aspects are provided by the following topics:

[0047] A combustor for a gas turbine engine includes: a front liner section having a rear end portion, wherein the front liner section defines a primary combustion chamber; and a rear liner section having a front end portion, wherein a passage is defined between the front liner section and the rear liner section, wherein during operation of the combustor, the passage guides a dilution air flow in a countercurrent direction relative to combustion gases exiting the primary combustion chamber.

[0048] According to the burner described in the foregoing clause, the front end portion of the rear liner section is at least partially disposed within the rear end portion of the front liner section.

[0049] The burner according to any one of the preceding clauses, wherein the front liner section includes a front inner liner and a front outer liner, wherein the front inner liner at least partially defines a first damping chamber and the front outer liner at least partially defines a second damping chamber, wherein the first damping chamber and the second damping chamber are disposed downstream of the primary combustion chamber and upstream of the rear liner section.

[0050] The burner according to any one of the preceding clauses, wherein the front liner section includes a first damping chamber and a second damping chamber disposed downstream of the primary combustion chamber and upstream of the rear liner section, wherein the front liner section further includes a first plurality of inlet orifices in fluid communication with the first damping chamber and a second plurality of inlet orifices in fluid communication with the second damping chamber.

[0051] The burner according to any one of the preceding clauses, wherein the front liner section includes a first exhaust port in fluid communication with the first damping chamber and the primary combustion chamber, and wherein the front liner section further includes a second exhaust port in fluid communication with the second damping chamber and the primary combustion chamber.

[0052] The burner according to any one of the preceding clauses, wherein during operation of the burner, at least one of the first exhaust port and the second exhaust port directs a flow of cooling air to the front liner section.

[0053] The burner according to any one of the preceding clauses, wherein the rear liner section includes a rear inner liner and a rear outer liner, wherein the rear inner liner at least partially defines a first damping chamber and the rear outer liner at least partially defines a second damping chamber, wherein the first damping chamber and the second damping chamber are disposed downstream of the primary combustion chamber and upstream of the secondary combustion chamber at least partially defined by the rear liner section.

[0054] The burner according to any one of the preceding clauses, wherein the rear liner section includes a first damping chamber and a second damping chamber, the first damping chamber and the second damping chamber being disposed downstream of the primary combustion chamber and upstream of a secondary combustion chamber at least partially defined by the rear liner section, wherein the rear liner section further includes a first plurality of inlet orifices in fluid communication with the first damping chamber and a second plurality of inlet orifices in fluid communication with the second damping chamber.

[0055] The burner according to any one of the preceding clauses, wherein the rear liner section includes a first exhaust port in fluid communication with the first damping chamber, and wherein the rear liner section includes a second exhaust port in fluid communication with the second damping chamber, wherein the first exhaust port and the second exhaust port are disposed upstream of a secondary combustion chamber at least partially defined by the rear liner section.

[0056] The burner according to any one of the preceding clauses, wherein during operation of the burner, at least one of the first exhaust port and the second exhaust port directs a cooling airflow to the rear liner section.

[0057] A gas turbine engine includes: a fan, a compressor section, a combustor section, and a turbine section, the combustor section including a combustor, the combustor including: a front liner section having a rear end portion, wherein the front liner section defines a primary combustion chamber; and a rear liner section having a front end portion, wherein the rear liner section at least partially defines a secondary combustion chamber, wherein a passage is defined between the rear end portion of the front liner section and the front end portion of the rear liner section, wherein the passage is oriented to guide a dilution airflow in a countercurrent direction relative to combustion gases flowing from the primary combustion chamber to the secondary combustion chamber during operation of the combustor.

[0058] The gas turbine engine according to any one of the preceding clauses, wherein the front end portion of the rear liner section is at least partially disposed within the rear end portion of the front liner section.

[0059] The gas turbine engine according to any one of the preceding clauses, wherein at least one of the front liner section and the rear liner section defines a plurality of cooling orifices.

[0060] The gas turbine engine according to any one of the preceding clauses, wherein the front liner section includes a first damping chamber and a second damping chamber, and the rear liner section includes a third damping chamber and a fourth damping chamber, wherein each of the first damping chamber, the second damping chamber, the third damping chamber and the fourth damping chamber is disposed downstream of the primary combustion chamber and upstream of the secondary combustion chamber.

[0061] The gas turbine engine according to any one of the preceding clauses, wherein the front liner section includes a first damping chamber and a second damping chamber disposed downstream of the primary combustion chamber and upstream of the rear liner section, wherein the front liner section further includes a first plurality of inlet orifices in fluid communication with the first damping chamber and a second plurality of inlet orifices in fluid communication with the second damping chamber.

[0062] The gas turbine engine according to any one of the preceding clauses, wherein the front liner section includes a first exhaust port in fluid communication with the first damping chamber and the primary combustion chamber, wherein during operation of the burner, the first exhaust port directs a flow of cooling air to the front liner section.

[0063] The gas turbine engine according to any one of the preceding clauses, wherein the rear liner section includes a first damping chamber and a second damping chamber, wherein the first damping chamber and the second damping chamber are disposed downstream of the primary combustion chamber and upstream of the secondary combustion chamber.

[0064] The gas turbine engine according to any one of the preceding clauses, wherein the rear liner section further includes a first plurality of inlet orifices in fluid communication with the first damping chamber and a second plurality of inlet orifices in fluid communication with the second damping chamber.

[0065] The gas turbine engine according to any one of the preceding clauses, wherein the rear liner section includes a first exhaust port in fluid communication with the first damping chamber and the primary combustion chamber, and wherein the rear liner section further includes a second exhaust port in fluid communication with the second damping chamber and the primary combustion chamber.

[0066] The gas turbine engine according to any one of the preceding clauses, wherein during operation of the burner, at least one of the first exhaust port and the second exhaust port of the rear liner section directs a cooling airflow to the rear liner section.

[0067] This written description uses examples to disclose this disclosure, including best practices, and also enables any person skilled in the art to practice this disclosure, including making and using any apparatus or system and performing any incorporated methods. The patentable scope of this disclosure is defined by the claims, but may include other examples that would occur to a person skilled in the art. Such other examples are intended to fall within the scope of the claims if they include structural elements that are not different from the literal language of the claims, or if they include equivalent structural elements that are not substantially different from the literal language of the claims.

Claims

1. A combustor for a gas turbine engine, characterized in that, The burner includes: A front liner section, the front liner section having a rear end portion, wherein the front liner section defines a primary combustion chamber; and The rear liner section has a front end portion disposed within the rear end portion of the front liner section, wherein a channel is radially defined between the inner surface of the front liner section and the outer surface of the rear liner section, wherein during operation of the burner, the channel guides a dilution airflow in a counter-current direction relative to the combustion gases flowing from the primary combustion chamber.

2. The burner according to claim 1, characterized in that, The front liner section includes a front inner liner and a front outer liner, wherein the front inner liner at least partially defines a first damping chamber and the front outer liner at least partially defines a second damping chamber, wherein the first damping chamber and the second damping chamber are disposed downstream of the primary combustion chamber and upstream of the rear liner section.

3. The burner according to claim 1, characterized in that, The front liner section includes a first damping chamber and a second damping chamber disposed downstream of the primary combustion chamber and upstream of the rear liner section, wherein the front liner section further includes a first plurality of inlet orifices in fluid communication with the first damping chamber and a second plurality of inlet orifices in fluid communication with the second damping chamber.

4. The burner according to claim 3, characterized in that, The front liner section includes a first exhaust port in fluid communication with the first damping chamber and the primary combustion chamber, and the front liner section further includes a second exhaust port in fluid communication with the second damping chamber and the primary combustion chamber.

5. The burner according to claim 4, characterized in that, During operation of the burner, at least one of the first exhaust port and the second exhaust port directs a flow of cooling air to the front liner section.

6. The burner according to claim 1, characterized in that, The rear liner section includes a rear inner liner and a rear outer liner, wherein the rear inner liner at least partially defines a first damping chamber and the rear outer liner at least partially defines a second damping chamber, wherein the first damping chamber and the second damping chamber are disposed downstream of the primary combustion chamber and upstream of the secondary combustion chamber at least partially defined by the rear liner section.

7. The burner according to claim 1, characterized in that, The rear liner section includes a first damping chamber and a second damping chamber, the first damping chamber and the second damping chamber being disposed downstream of the primary combustion chamber and upstream of a secondary combustion chamber at least partially defined by the rear liner section, wherein the rear liner section further includes a first plurality of inlet orifices in fluid communication with the first damping chamber and a second plurality of inlet orifices in fluid communication with the second damping chamber.

8. The burner according to claim 7, characterized in that, The rear liner section includes a first exhaust port in fluid communication with the first damping chamber, and the rear liner section includes a second exhaust port in fluid communication with the second damping chamber, wherein the first exhaust port and the second exhaust port are located upstream of the secondary combustion chamber.

9. The burner according to claim 8, characterized in that, During operation of the burner, at least one of the first exhaust port and the second exhaust port directs a flow of cooling air to the rear liner section.

10. A gas turbine engine, characterized in that, The gas turbine engine includes: The fan, compressor section, burner section, and turbine section, wherein the burner section includes a burner, and the burner includes: A front liner section, the front liner section having a rear end portion, wherein the front liner section defines a primary combustion chamber; and A rear liner section having a front end portion disposed within the rear end portion of the front liner section, wherein the rear liner section at least partially defines a secondary combustion chamber, wherein a channel is radially defined between an inner surface of the rear end portion of the front liner section and an outer surface of the front end portion of the rear liner section, wherein the channel is oriented to guide a dilution airflow in a countercurrent direction relative to combustion gases flowing from the primary combustion chamber to the secondary combustion chamber during operation of the burner.

11. The gas turbine engine according to claim 10, characterized in that, At least one of the front liner section and the rear liner section defines a plurality of cooling orifices.

12. The gas turbine engine according to claim 10, characterized in that, The front liner section includes a first damping chamber and a second damping chamber, and the rear liner section includes a third damping chamber and a fourth damping chamber, wherein each of the first damping chamber, the second damping chamber, the third damping chamber and the fourth damping chamber is disposed downstream of the primary combustion chamber and upstream of the secondary combustion chamber.

13. The gas turbine engine according to claim 10, characterized in that, The front liner section includes a first damping chamber and a second damping chamber disposed downstream of the primary combustion chamber and upstream of the rear liner section, wherein the front liner section further includes a first plurality of inlet orifices in fluid communication with the first damping chamber and a second plurality of inlet orifices in fluid communication with the second damping chamber.

14. The gas turbine engine according to claim 13, characterized in that, The front liner section includes a first exhaust port in fluid communication with the first damping chamber and the primary combustion chamber, wherein during operation of the burner, the first exhaust port directs a flow of cooling air to the front liner section.

15. The gas turbine engine according to claim 10, characterized in that, The rear liner section includes a first damping chamber and a second damping chamber, wherein the first damping chamber and the second damping chamber are disposed downstream of the primary combustion chamber and upstream of the secondary combustion chamber.

16. The gas turbine engine according to claim 15, characterized in that, The rear lining section further includes a first plurality of inlet orifices in fluid communication with the first damping chamber and a second plurality of inlet orifices in fluid communication with the second damping chamber.

17. The gas turbine engine according to claim 16, characterized in that, The rear liner section includes a first exhaust port in fluid communication with the first damping chamber and the primary combustion chamber, and the rear liner section further includes a second exhaust port in fluid communication with the second damping chamber and the primary combustion chamber.

18. The gas turbine engine according to claim 17, characterized in that, During operation of the burner, at least one of the first exhaust port and the second exhaust port of the rear liner section directs a cooling airflow to the rear liner section.

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