Burner dilution holes

Abstract

A turbofan gas turbine engine configured to reduce hot spots within a combustor. The engine includes an axis and a combustor arranged circumferentially about the axis. The combustor includes an annular combustor liner including a front portion and a rear portion. The annular combustor liner is joined to an annular combustor dome via the front portion and defines a chamber, and a nozzle is mounted within the annular combustor dome and configured to inject fuel into a plurality of swirlers. At least one or more dilution openings are distributed circumferentially around the liner such that a region is fluidly connected to the chamber through the annular combustor liner. Each of the plurality of dilution openings includes an opening and a radial support wall, the radial support wall being positioned rearward of the opening such that the radial support wall extends into the chamber.

Description

Technical Field

[0001] The present invention relates generally to gas turbine engines and, more particularly, to combustors therein. Background Art

[0002] In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor to produce hot combustion gases. The hot gases flow downstream through the turbine stages, which extract energy from them. The high-pressure turbine powers the compressor. The low-pressure turbine produces useful work by powering the upstream fan in a typical turbofan gas turbine aircraft engine application, for example.

[0003] Combustor performance is critical to the overall performance of a gas turbine engine. Compressed air is mixed with fuel in the combustor to produce a fuel and air mixture that is ignited to produce combustion gases.

[0004] Additional air is introduced into the burner through the dilution openings. Additional air is introduced through the dilution openings to provide a predetermined fuel to air ratio in a specific area of ​​the burner. Summary of the Invention

[0005] The technology disclosed herein is configured to reduce liner wall gas temperatures by up to 500°F and thereby eliminate hot spots, thereby improving liner durability. Other benefits include good controllability of profile / pattern factors and reduction of NO X The production of.

[0006] According to one aspect of the present invention, a turbofan gas turbine engine is provided that is configured to reduce hot spots within a combustor. The engine includes an axis and a combustor arranged circumferentially around the axis. The combustor includes an annular combustor liner that includes a front portion and a rear portion. The annular combustor liner is joined to an annular combustor dome by the front portion and defines a chamber, and a nozzle is mounted within the annular combustor dome and is configured to inject fuel into a plurality of swirlers. At least one or more dilution openings are distributed circumferentially around the liner so that the region is fluidly connected to the chamber through the annular combustor liner. Each of the one or more dilution openings includes an opening and a radial support wall, the radial support wall being positioned behind the opening so that the radial support wall extends into the chamber.

[0007] According to another aspect of the present invention, a turbofan gas turbine engine is provided, configured to reduce hot spots within a combustor. The engine includes an axis and a combustor circumferentially arranged about the axis. The combustor includes an annular combustor dome. An annular combustor liner includes a front portion and an aft portion, the annular combustor liner being joined to the annular combustor dome via the front portion and defining a chamber, and a nozzle mounted within the annular combustor dome and configured to inject fuel into a plurality of swirlers. One or more dilution openings are distributed circumferentially around the liner, such that a region is fluidly connected to the chamber through the annular combustor liner. Each of the one or more dilution openings includes an opening and a radial support wall positioned rearward of the opening such that the radial support wall extends into the chamber. Each of the one or more dilution openings defines an annular groove. A guide channel fluidly connects the region to the opening. The support wall has a first front surface and a second aft surface. At least one aperture is defined through the support wall, fluidly connecting the first front surface to the second aft surface. A plurality of undulations are formed on at least one of the first front surface and the second aft surface. A plurality of aft liner holes are positioned in the aft portion of the liner wall and are configured to fluidly connect the region with a potential recirculation zone. The support wall defines a substantially perpendicular angle with the axis of the combustor, and a second aft side surface of the support wall is positioned so as to define an obtuse angle with the aft portion.

[0008] According to another aspect of the present invention, a method for reducing hot spots in a combustor turbofan gas turbine engine is provided, the combustor turbofan gas turbine engine comprising: an axis; a combustor, the combustor being circumferentially arranged about the axis; and the combustor comprising an annular combustor dome; an annular combustor liner, the annular combustor liner comprising a front portion and an aft portion, the annular combustor liner being joined to the annular combustor dome by the front portion and defining a chamber, and a nozzle being mounted within the annular combustor dome and configured to inject fuel into a plurality of swirlers; a plurality of dilution openings distributed circumferentially about the liner such that a region is fluidly connected to the chamber through the annular combustor liner; each of the plurality of dilution openings comprising an opening and a radial support wall, the radial support wall being positioned aft of the opening such that the radial support wall extends into the chamber; each of the plurality of dilution openings defining an annular slot; a guide passage fluidly connecting the region with the opening, the method comprising the steps of: flowing combustion gases through the combustion chamber, passing through dilution holes; introducing dilution air into the chamber through the dilution holes; and swirling and mixing the combustion gases due to the introduction of the dilution air in front of the support wall. BRIEF DESCRIPTION OF THE DRAWINGS

[0009] The present invention, according to preferred and exemplary embodiments, together with further objects and advantages thereof, is described in more detail in the following detailed description taken in conjunction with the accompanying drawings, in which:

[0010] Figure 1is an axial partial cross-sectional view of a portion of an exemplary annular combustor of a turbofan gas turbine engine according to the disclosed technology;

[0011] Figure 2 is Figure 1 An enlarged axial cross-sectional view of another combustor according to the disclosed technology, taken at area 2;

[0012] Figure 3 Shows the disclosed technology Figure 2 Alternating wall configuration of dilution openings in the burner;

[0013] Figure 4 shows a detailed view of a dilution well according to the disclosed technology;

[0014] Figure 5 shows another detailed view of a dilution well according to the disclosed technology;

[0015] Figure 6 Another detailed view of a dilution hole according to the disclosed technology is shown, showing a radial wall having a flow constructor formed on its upstream side;

[0016] Figure 7 Another detailed view of a dilution hole according to the disclosed technology is shown, showing a radial wall having a plurality of axial holes formed therethrough;

[0017] Figure 8 Another detailed view of a dilution hole according to the disclosed technology is shown, showing a radial wall having a plurality of outwardly angled holes formed therethrough;

[0018] Figure 9 The wall of the dilution hole with a V-shaped profile is shown;

[0019] Figure 10 The wall of the dilution hole with another profile is shown;

[0020] Figure 11 shows a stepped dilution hole of the present invention having holes formed therethrough behind a radial wall, the holes being configured to weaken the recirculation zone; and

[0021] Figure 12 A stepped dilution hole of the disclosed art is shown having an axial hole formed through a radial wall thereof and an additional hole formed behind the radial wall. DETAILED DESCRIPTION

[0022] Referring to the drawings, wherein like reference numerals refer to like elements throughout the several views, Figure 1The disclosed technology illustrated in FIG is a portion of an exemplary turbofan gas turbine engine including an annular combustor having a liner with stepped dilution openings as described in detail below. Discrete dilution openings are a conventionally known structure for introducing dilution air into a combustor. Discrete dilution openings have an inherent problem of forming hot spots axially aft. In addition, the discrete dilution openings define a recirculation zone aft that serves to draw in hot gases during operation and, therefore, results in increased NOx production. The stepped dilution openings of the present invention address this problem. The present invention will be described in terms of a number of embodiments.

[0023] Now refer to Figure 1 and 2 , annular combustor 10 is a single annular combustor design and is suitably mounted within a casing 11 spaced from a longitudinal or axial centerline axis 12. It will be appreciated that the exemplary turbofan gas turbine engine includes a plurality of annular combustors 10 arranged circumferentially about the centerline axis 12.

[0024] The combustor 10 includes an annular combustor liner 14, each including an integrally joined forward portion 15 and aft portion 16. The forward portion 15 and aft portion 16 are each mechanically joined at a joint 17. The aft portion 16 includes an overlapping liner portion 19 extending over a portion of the forward portion 15. According to one embodiment, the overlapping liner portion 19 is used to create the joint 17 and define the guide passage 58. The annular combustor liner 14 is suitably joined to the annular combustor dome 18 at the upstream end via the forward portion 15. Radially outer and inner shrouds extend axially forward from the dome 18 at the joint with the outer and inner liners to define an annular plenum 24 on the upstream side of the dome 18. The outer surfaces of the dome 18 and the aft liner 14, together with the inner surface of the casing 11, define the region 13.

[0025] like Figure 1 As shown, the engine includes a suitable compressor 26, such as a conventional multi-stage axial flow compressor, which is suitably configured to pressurize the airflow 28 as it flows downstream. The pressurized airflow 28 is directed axially downstream from the compressor 26 through a suitable diffuser and introduced into the plenum 24 through a first annular inlet 34. The combustor 10 and compressor 26, as described above, can have any conventional configuration.

[0026] According to the present invention, Figure 1 The combustor 10 shown in FIG includes a nozzle 41 and a plurality of swirlers 49 suitably mounted in the combustor dome 18. Each nozzle 41 injects fuel into the plurality of swirlers 49, wherein the fuel mixes with the pressurized gas stream air 28 within the throat to produce a fuel and air mixture that is suitably ignited to produce hot combustion gases 36 that together flow downstream through a chamber 55 defined by the combustor liner 14.

[0027] Combustion gases 36 are discharged from the outlet end of the combustor into a high-pressure turbine (not shown), which extracts energy therefrom for powering the compressor 26. A low-pressure turbine (not shown) is disposed downstream of the high-pressure turbine and is suitably configured to produce output power, such as for powering an upstream fan in a typical turbofan gas turbine engine aircraft application.

[0028] The combustor 10 includes one or more dilution holes 50 distributed circumferentially around the liner 14 and configured to fluidly connect the region 13 to the chamber 55. The dilution holes 50 of the disclosed technology are stepped, such as Figure 1 and 2 . Radial dilution support wall 52 is positioned directly behind opening 56 in combustor liner 14 such that radial dilution support wall 52 extends into chamber 55. Radial support wall 52 defines a front surface 53 and an aft surface 54. Opening 56 in combustor liner fluidly connects region 13 to combustor chamber 55 via at least one of the one or more dilution holes 50.

[0029] As a main embodiment, the opening 56 of the stepped dilution hole 50 can be formed as an annular groove or a plurality of adjacent annular grooves. In one embodiment, the opening 56 is completely annular. The shape of the stepped dilution hole can be uniform or non-uniform. In this regard, it can have the following characteristics: Figure 3 A circumferential stepped hole with an undulation 61 is shown. The undulation 61 can be located on both the front side 53 of the wall 52 and the rear side 54 of the wall 52. According to one embodiment, the undulation 61 is located only on the front side 53 of the wall 52. According to another embodiment, the undulation 61 is located only on the rear side 54 of the wall 52. The undulation can have various shapes, such as semicircular, triangular, rectangular, and other geometric shapes. Multiple such steps can be provided as desired.

[0030] like Figure 4 and Figure 5 As shown, the discrete joint overlaps 19 are configured to define guide channels 58. Guide channels 58 are configured to allow dilution fluid to flow axially therethrough from region 13 to aperture 56. The joint overlaps 19 may be vane-shaped such that they are configured to impart swirl to the dilution flow. In this regard, as shown Figure 4 The illustrated vanes 63 are positioned to impart swirl to the dilution flow as it approaches the orifice 56 through the guide passage 58. Figure 5 As shown, the dilution flow of the stepped dilution hole 50 is supplied by holes / slots 57 defined through the joint overlap 19.

[0031] Now refer to Figure 6The biggest advantage of having support walls 52 is improved penetration during mixing. According to some embodiments, support walls 52 provide surfaces that define turbulent features 62 for inducing turbulence. As used herein, the term "turbulent" refers to features that are configured to promote turbulent flow. Figure 6 As shown, the feature 62 is a plurality of dimples, one of which is shown in cross-section. In other embodiments, the feature 62 can be configured as one of the following: a bump, a notch, a step, a groove, other structures, and combinations thereof.

[0032] Now refer to Figure 7 and 8 According to one embodiment, a plurality of holes or perforations 72 are positioned across the support wall 52 to fluidically connect the first front side surface 53 with the second rear side surface 54. In this way, a portion of the dilution flow is introduced from the front of the dilution wall 52 through the dilution wall 52 to the rear region of the dilution wall 52. In this way, wake mitigation is achieved. The plurality of holes or perforations 72 may be circular / rectangular or any shape that may be desired, or they may be block-like structures. The arrangement of the plurality of holes or perforations 72 is determined so that the perforations 72 are configured to mitigate undesirable hydrodynamic forward rafts of the support wall 52. It will be understood that the plurality of perforations 72 may be inclined circumferentially, radially, axially, and combinations thereof. As shown in FIG. Figure 8 The perforations 72 are shown to be angled.

[0033] like Figure 9 As shown, the support wall 52 defines an integral feature 81. The shape of the integral feature 81 can be determined in various embodiments to further enhance the penetration and mixing of the dilution air. The integral feature 81 can be wavy, aerodynamic, or have a stepped, V-shaped, or other characteristic shape to reduce or prevent recirculation zones. Figure 9 As shown, the integral feature 81 defines a V-shape.

[0034] like Figure 10 As shown, it is parallel to the axis 12 looking rearward and substantially perpendicular to the support wall 52, which defines an edge 87. The edge 87 can be equidistant from the axis 12 around the circumferential length of the support wall 42. Alternatively, the distance of the edge 87 from the axis 12 can be varied to define other shapes. Figure 10 In FIG, the support wall 52 is configured such that the edge 87 is wavy. In this regard, the edge 87 defines a peak 91 that is closer to the axis 12 than a valley 93 (shown in dashed lines).

[0035] Now refer to Figure 11, the aft liner holes 74 are positioned downstream or aft of the dilution holes 56 and are configured to direct air into a potential recirculation zone 76 that may be created by the support wall 52. As used herein, the term "recirculation zone" refers to an area in which fluid flow and subsequent mass transfer and heat transfer are reduced relative to the flow of combustor gases 36 through the central region of the combustor chamber 55. One result of the formation of a recirculation zone is the creation of hot spots, and therefore, the function of the aft liner holes 74 is to reduce or eliminate hot spots compared to conventional turbine engines. The aft liner holes 74 may be a single row of slots / round holes, or as Figure 12 As shown, multiple rows of slots / holes can also be used to target specific hot gas zones, thereby enhancing mixing, and can be used as an additional control profile / pattern factor.

[0036] like Figure 11 and 12 As shown, the support wall 52 defines a rear surface 84 and the rear portion 16 defines an inner surface 82. Figure 11 As shown, the rear surface 84 and the inner surface 82 intersect and together define an angle α, which is approximately 90°. Figure 12 As shown, the angle α is greater than 90°. It should be understood that the angle α can be adjusted by changing the thickness of the support wall 52 or changing the thickness of the rear portion 16 (e.g. Figure 12 In the embodiment shown, the support wall 52 is positioned perpendicular to the axis of the burner. It should be understood that the support wall 52 can be tilted forward or backward relative to the axis of the burner 10, and the magnitude of the angle α can be determined by the tilt of the support wall 52 relative to the axis 12 of the burner 10. A plurality of support walls 52 may be provided.

[0037] It will be appreciated that the above structure can be additively manufactured so that the liner is a single part. Thus, there is no need for two mounting features. In some embodiments, the liner is two separate parts that are mounted separately. It will be appreciated that the front portion 15 and the rear portion 16 can also be joined using fasteners such as bolts to allow air to pass axially or radially through the holes.

[0038] The presently disclosed technology may be better understood by describing its operation. During operation of the combustor 10, air 28 is pressurized by the compressor 26. The airflow 28 then flows through the annular inlet 34 into the plenum 24. Combustion gases 36 flow into the combustor chamber 55. Dilution air flows through the holes 50 to mix with the combustion gases 36 and to confine its flow to be more uniform and reduce the formation of hot spots relative to the flow behind conventional dilution holes. In some embodiments, the dilution air is further introduced through the rear liner holes 74 to prevent the formation of recirculation within the potential recirculation zone 76. The technology disclosed herein provides a more predictable profile within the combustor, thereby reducing early turbine maintenance issues associated with conventional technology. When the dilution air flows through the holes 52 to mix with the combustion gases 36, fewer polluting byproducts, such as NOx, are produced relative to conventional dilution air.

[0039] A device, namely a burner including a stepped dilution hole, has been described above. As a primary embodiment, the stepped dilution hole has an opening with an elongated annular groove with two or more joints. In special cases, the hole can be completely annular. The shape of the stepped dilution hole can be uniform or non-uniform. As shown in the figure, it can have a circumferentially stepped hole with undulations. The undulations can have various shapes: semicircular, triangular, rectangular, etc. Multiple such steps can be provided as needed.

[0040] Unless expressly stated otherwise, each feature disclosed in this specification (including any accompanying claims, abstract, and drawings) may be replaced by alternative features serving the same, equivalent, or similar purpose. Thus, unless expressly stated otherwise, each feature disclosed is one example only of a generic series of equivalent or similar features.

[0041] The invention is not limited to the details of the foregoing embodiments. The invention extends to any novel one or any novel combination of features disclosed in this specification (including any accompanying claims, abstract and drawings), or any novel one or any novel combination of steps of any disclosed method or process.

[0042] Further aspects of the invention are provided by the subject matter of the following numbered clauses:

[0043] 1. A turbofan gas turbine engine configured to reduce hot spots within a combustor, the engine comprising: an axis; a combustor arranged circumferentially about the axis; the combustor comprising: an annular combustor dome; an annular combustor liner comprising a front portion and an aft portion, the annular combustor liner being joined to the annular combustor dome by the front portion and defining a chamber, and a nozzle mounted within the annular combustor dome and configured to inject fuel into a plurality of swirlers; at least one or more dilution openings distributed circumferentially about the liner such that a region is fluidly connected to the chamber through the annular combustor liner; and wherein each of the one or more dilution openings comprises an opening and a radial support wall, the radial support wall being positioned rearward of the opening such that the radial support wall extends into the chamber.

[0044] 2. A turbofan gas turbine engine according to any preceding clause, wherein each of the one or more dilution openings defines an annular slot.

[0045] 3. A turbofan gas turbine engine according to any preceding clause, wherein the guide passage fluidly connects the region with the opening.

[0046] 4. A turbofan gas turbine engine according to any preceding clause, wherein blades configured to direct dilution air are positioned in the guide passage.

[0047] 5. The turbofan gas turbine engine of any preceding clause, wherein an aperture is defined through the overlap and fluidly connects the region to one of: a guide passage, an opening, and a combination thereof.

[0048] 6. A turbofan gas turbine engine according to any preceding clause, wherein the support wall has a first front side surface and a second rear side surface, and at least one hole is defined through the support wall so that the first front side surface is fluidly connected to the second rear side surface.

[0049] 7. A turbofan gas turbine engine according to any preceding clause, wherein a plurality of undulations are formed on at least one of the first front side surface and the second rear side surface.

[0050] 8. A turbofan gas turbine engine according to any preceding clause, wherein a plurality of aft liner holes are positioned in the aft portion of the liner wall and are configured to fluidly connect the area with the potential recirculation zone.

[0051] 9. A turbofan gas turbine engine according to any preceding clause, wherein the support wall defines an angle substantially perpendicular to the axis of the combustor, and wherein the second aft side surface of the support wall is positioned such that it defines an obtuse angle with the aft portion.

[0052] 10. The turbofan gas turbine engine of any preceding clause, wherein the guide passage is positioned such that the guide passage is configured to fluidly connect the region with the opening.

[0053] 11. A turbofan gas turbine engine according to any preceding clause, wherein the support wall has a first front side surface and a second rear side surface, and the plurality of undulations are formed on the first front side surface.

[0054] 12. A turbofan gas turbine engine according to any preceding clause, wherein the support wall has a first front side surface and a second rear side surface, the plurality of undulations being formed on the second rear side surface.

[0055] 13. A turbofan gas turbine engine configured to reduce hot spots within a combustor, the engine comprising: an axis; a combustor arranged circumferentially about the axis; the combustor comprising: an annular combustor dome; an annular combustor liner, the annular combustor liner comprising a front portion and an aft portion, the annular combustor liner being joined to the annular combustor dome via the front portion and defining a chamber, and a nozzle mounted within the annular combustor dome and configured to inject fuel into a plurality of swirlers; at least one or more dilution openings distributed circumferentially about the liner such that a region is fluidly connected to the chamber through the annular combustor liner; each of the one or more dilution openings comprising an opening and a radial support wall , a radial support wall is positioned rearward of the opening such that the radial support wall extends into the chamber; each of the one or more dilution openings defines an annular groove; a guide passage fluidly connects the region with the opening; the support wall has a first front side surface and a second rear side surface; at least one hole is defined through the support wall such that the first front side surface is fluidly connected to the second rear side surface; a plurality of undulations are formed on at least one of the first front side surface and the second rear side surface; a plurality of rear liner holes are positioned in the rear portion of the liner wall and are configured to fluidly connect the region with a potential recirculation zone; and wherein the support wall defines a substantially perpendicular angle with an axis of the combustor, and the second rear side surface of the support wall is positioned such that it defines an obtuse angle with the rear portion.

[0056] 14. A turbofan gas turbine engine according to any preceding clause, wherein the guide passage fluidly connects the region with the opening.

[0057] 15. A turbofan gas turbine engine according to any preceding clause, wherein vanes configured to direct dilution air are positioned in the guide passage.

[0058] 16. The turbofan gas turbine engine of any preceding clause, wherein an aperture is defined through the overlap and fluidly connects the region to one of: a guide passage, an opening, and a combination thereof.

[0059] 17. A turbofan gas turbine engine according to any preceding clause, wherein the support wall has a first front side surface and a second rear side surface, and the plurality of undulations are formed on the first front side surface.

[0060] 18. A turbofan gas turbine engine according to any preceding clause, wherein the support wall has a first front side surface and a second rear side surface, the plurality of undulations being formed on the second front side surface.

[0061] 19. A method for reducing hot spots in a combustor turbofan gas turbine engine, the combustor turbofan gas turbine engine comprising: an axis; a combustor, the combustor being circumferentially arranged about the axis; and the combustor comprising: an annular combustor dome; an annular combustor liner, the annular combustor liner comprising a front portion and an aft portion, the annular combustor liner being joined to the annular combustor dome by the front portion and defining a chamber, and a nozzle mounted in the annular combustor dome and configured to inject fuel into a plurality of swirlers; at least one or more dilution openings distributed circumferentially about the liner such that a region is fluidly connected to the chamber through the annular combustor liner; each of the one or more dilution openings comprising an opening and a radial support wall, the radial support wall being positioned aft of the opening such that the radial support wall extends into the chamber; each of the one or more dilution openings defining an annular slot; a guide passage fluidly connecting the region with the opening, the method comprising the steps of: flowing combustion gases through the combustion chamber, past dilution holes; introducing dilution air into the chamber through the dilution holes; and swirling and mixing the combustion gases due to the introduction of the dilution air forward of the support wall.

[0062] 20. A method according to any preceding clause, further comprising the step of introducing dilution air via apertures through the rear lining behind the support wall, thereby reducing potential recirculation zones.

Claims

1. A turbofan gas turbine engine configured to reduce hot spots within a combustor, characterized in that: The turbofan gas turbine engine comprises: axis; a burner, the burner being arranged circumferentially around the axis; The burner comprises: annular burner dome; an annular combustor liner including a front portion and an aft portion, the annular combustor liner being joined to an annular combustor dome by the front portion and defining a chamber, and a nozzle mounted within the annular combustor dome and configured to inject fuel into the plurality of swirlers; at least one or more dilution openings distributed circumferentially about the annular combustor liner such that a region is fluidly connected to the chamber through the annular combustor liner; and wherein each of the one or more dilution openings comprises an opening and a radial support wall, the radial support wall being positioned rearward of the opening such that the radial support wall extends into the chamber; wherein each of the one or more dilution openings defines an annular groove; wherein the radial support wall has a first front side surface and a second rear side surface, and at least one hole is defined through the radial support wall such that the first front side surface is fluidly connected to the second rear side surface.

2. The turbofan gas turbine engine according to claim 1, wherein: in, A guide channel fluidly connects the region with the opening.

3. The turbofan gas turbine engine according to claim 2, characterized in that: in, Vanes configured to guide dilution air are positioned in the guide passage.

4. The turbofan gas turbine engine according to claim 3, characterized in that in, A hole is defined through the overlapping liner portion of the annular combustor liner and fluidly connects the region to one of: the guide passage, the opening, and a combination thereof.

5. The turbofan gas turbine engine according to claim 1, wherein: in, A plurality of undulations are formed on at least one of the first front-side surface and the second rear-side surface.

6. The turbofan gas turbine engine according to claim 5, characterized in that in, A plurality of aft liner holes are positioned in the aft portion of the annular combustor liner and are configured to fluidly connect the region with a potential recirculation zone.

7. The turbofan gas turbine engine according to claim 6, characterized in that: in, The radial support wall defines a substantially perpendicular angle with the axis of the combustor, and the second rear side surface of the radial support wall is positioned so as to define an obtuse angle with the rear portion.

8. The turbofan gas turbine engine according to claim 7, characterized in that in, The guide channel is positioned such that the guide channel is configured to fluidly connect the region with the opening.

9. The turbofan gas turbine engine according to claim 1, wherein: in, The radial support wall has a first front side surface and a second rear side surface, and a plurality of undulations are formed on the first front side surface.

10. The turbofan gas turbine engine according to claim 1, wherein: in, The radial support wall has a first front side surface and a second rear side surface, and a plurality of undulations are formed on the second rear side surface.

11. A turbofan gas turbine engine configured to reduce hot spots within a combustor, characterized in that: The turbofan gas turbine engine comprises: axis; a burner, the burner being arranged circumferentially around the axis; The burner comprises: annular burner dome; an annular combustor liner including a front portion and an aft portion, the annular combustor liner being joined to an annular combustor dome by the front portion and defining a chamber, and a nozzle mounted within the annular combustor dome and configured to inject fuel into the plurality of swirlers; a plurality of dilution openings distributed circumferentially about the annular combustor liner such that a region is fluidly connected to the chamber through the annular combustor liner; each of the plurality of dilution openings includes an opening and a radial support wall positioned rearward of the opening such that the radial support wall extends into the chamber; Each of the plurality of dilution openings defines an annular slot; a guide channel fluidly connecting the region with the opening; The radial support wall has a first front side surface and a second rear side surface; at least one aperture defined through the radial support wall fluidly connecting the first front side surface with the second rear side surface; a plurality of undulations formed on at least one of the first front surface and the second rear surface; a plurality of aft liner holes positioned in the aft portion of the annular combustor liner and configured to fluidly connect the region with a potential recirculation zone; And wherein the radial support wall defines a substantially perpendicular angle with the axis of the combustor, and the second rear side surface of the radial support wall is positioned so as to define an obtuse angle with the rear portion.

12. The turbofan gas turbine engine according to claim 11, characterized in that in, A guide channel fluidly connects the region with the opening.

13. The turbofan gas turbine engine according to claim 12, wherein: in, Vanes configured to guide dilution air are positioned in the guide passage.

14. The turbofan gas turbine engine according to claim 11, wherein: in, A hole is defined through the overlapping liner portion of the annular combustor liner and fluidly connects the region to one of: the guide passage, the opening, and a combination thereof.

15. The turbofan gas turbine engine according to claim 11, wherein: in, The radial support wall has a first front side surface and a second rear side surface, and a plurality of undulations are formed on the first front side surface.

16. The turbofan gas turbine engine according to claim 11, wherein: in, The radial support wall has a first front side surface and a second rear side surface, and a plurality of undulations are formed on the second front side surface.

17. A method for reducing hot spots in a combustor turbofan gas turbine engine, characterized in that The combustor turbofan gas turbine engine comprises: an axis; a combustor, the combustor being circumferentially arranged about the axis; and the combustor comprising: an annular combustor dome; an annular combustor liner, the annular combustor liner comprising a front portion and a rear portion, and the annular combustor liner is joined to the annular combustor dome by the front portion and defines a chamber, and a nozzle is mounted in the annular combustor dome and is configured to inject fuel into a plurality of swirlers; at least one or more dilution openings are distributed circumferentially about the annular combustor liner so that a region is fluidly connected to the chamber through the annular combustor liner; each of the one or more dilution openings comprises an opening and a radial support wall, the radial support wall being positioned rearward of the opening so that the radial support wall extends into the chamber; each of the one or more dilution openings defines an annular groove; a guide channel fluidly connects the region with the opening, and the method comprises the following steps: flowing combustion gases through the chamber, past the one or more dilution openings; introducing dilution air into the chamber through the one or more dilution openings; and Due to the introduction of the dilution air in front of the radial support wall, the combustion gas is swirled and mixed; Dilution air is introduced via holes through the aft portion of the annular combustor liner aft of the radial support wall, thereby reducing potential recirculation zones.

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