A wall structure of an aero-engine combustion chamber liner

Abstract

This application belongs to the technical field of the design of the wall structure of an aero-engine combustion chamber, and specifically relates to a wall structure of an aero-engine combustion chamber. Its designed annular inner wall is formed by electron beam welding between the two side edges of multiple inner wall fan-shaped segments. The annular outer wall is composed of multiple outer wall fan-shaped segments and outer fan-shaped connecting segments distributed at intervals. Through the outward flanges on the two side edges of each outer fan-shaped connecting segment, it is lap brazed or electron beam welded to the outside of the adjacent outer wall fan-shaped segment. The welding connection parts between each outer fan-shaped connecting segment and the outer wall fan-shaped segment are located on the outside, and will not affect the position distribution of the impingement cooling holes and turbulators on it. Corresponding to the circumferential joint surface position between the two side edges of each inner wall fan-shaped segment, multiple impingement cooling holes and their turbulators are distributed on the side wall of each outer fan-shaped segment, which can ensure the cooling effect and avoid ablation.

Description

Technical Field

[0001] This application belongs to the technical field of the structural design of the wall of an aero-engine combustion liner, and particularly relates to a wall structure of an aero-engine combustion liner. Background Art

[0002] In order to pursue higher thrust, it is necessary to increase the temperature before the turbine of an aero-engine, resulting in a significant increase in the temperature inside the combustion liner in the combustion chamber. The wall of the combustion liner is subjected to a high temperature load and is prone to ablation. Therefore, a porous lamellar cooling technology is designed for the wall of the combustion liner to efficiently cool the wall of the combustion liner.

[0003] Currently, for the wall of the combustion liner adopting the porous lamellar cooling technology, the design includes an annular outer wall, and an annular inner wall is arranged inside the outer wall. Among them, there are a plurality of impingement cooling holes on the side wall of the outer wall, and a plurality of turbulators are formed on the inner side. There are a plurality of film holes on the side wall of the inner wall; each turbulator is connected to the outer side of the inner wall by diffusion welding or brazing, as Figure 1 shown. The cooling air can enter between the outer wall and the inner wall through each impingement cooling hole for impingement cooling, and after being turbulized by each turbulator, it is discharged through each film hole, forming a film on the inner side of the inner wall for film cooling. In this way, the wall of the combustion liner can be efficiently cooled to protect the wall of the combustion liner from ablation.

[0004] Limited by the processing technology, the wall of the current combustion liner adopting the porous lamellar cooling technology mainly adopts a circumferential segmented + welding structural form. The outer wall and the inner wall are divided into a plurality of corresponding fan-shaped segments, and the edges of each fan-shaped segment are welded together by electron beam welding. At the circumferential joint surface, it is a solid structure and turbulators cannot be arranged, resulting in poor local cooling effect. In practical applications, ablation is prone to occur.

[0005] In view of the existence of the above technical defects, this application is proposed.

[0006] It should be noted that the disclosure of the above background art content is only used to assist in understanding the inventive concept and technical solution of the present invention, and it does not necessarily belong to the prior art of this patent application. Without clear evidence indicating that the above content was publicly available on the filing date of this application, the above background art should not be used to evaluate the novelty and inventiveness of this application. Summary of the Invention

[0007] The purpose of this application is to provide a wall structure of an aero-engine combustion liner to overcome or mitigate at least one aspect of the known technical defects.

[0008] The technical solution of this application is as follows:

[0009] A wall structure of an aero-engine combustion liner, comprising:

[0010] Multiple outer wall sector segments with multiple impingement cooling holes on their side walls and multiple turbulators formed on the inner side;

[0011] Multiple inner wall sector segments with multiple film holes formed on their side walls, and the outer sides thereof are connected to the turbulators on each outer wall sector segment by diffusion welding or brazing; the two side edges of each inner wall sector segment are welded by electron beam welding to form an annular inner wall; there is a connection gap between adjacent outer wall sector segments;

[0012] Multiple outer sector connecting segments are arranged in each connection gap, with multiple impingement cooling holes on their side walls, multiple turbulators formed on the inner side, and outward flanges on both side edges; there is a small gap between the turbulators on each outer sector connecting segment and the annular inner wall; the outward flanges on both side edges of each outer sector connecting segment overlap the outer sides of the corresponding adjacent outer wall sector segments and are connected to the outer sides of the corresponding adjacent outer wall sector segments by brazing or electron beam welding to form an annular outer wall with each outer wall sector segment.

[0013] According to at least one embodiment of the present application, in the above-mentioned aeroengine flame tube wall structure, the impingement cooling holes, turbulators on the annular outer wall and the film holes on the inner wall sector segments are distributed alternately in the circumferential and axial directions.

[0014] According to at least one embodiment of the present application, in the above-mentioned aeroengine flame tube wall structure, the small gap between the turbulators on each outer sector connecting segment and the annular inner wall is not greater than 0.1 mm.

[0015] According to at least one embodiment of the present application, in the above-mentioned aeroengine flame tube wall structure, each outer sector connecting segment includes at least one row of impingement cooling holes and two rows of turbulators.

[0016] The present application has at least the following beneficial technical effects:

[0017] Provide an aeroengine flame tube wall structure, the annular inner wall of which is formed by electron beam welding between the two side edges of multiple inner wall sector segments, and the annular outer wall is formed by multiple alternately distributed outer wall sector segments and outer sector connecting segments, which are connected by brazing or electron beam welding with the outward flanges on both side edges of each outer sector connecting segment overlapping the outer sides of the adjacent outer wall sector segments. The welding connection part between each outer sector connecting segment and the outer wall sector segment is located on the outside, which will not affect the position distribution of the impingement cooling holes and turbulators thereon. Corresponding to the circumferential joint surface position between the two side edges of each inner wall sector segment, multiple impingement cooling holes and their turbulators are distributed on the side walls of each outer sector segment, which can ensure the cooling effect and avoid ablation. Description of the Drawings

[0018] Figure 1It is a schematic diagram of the wall structure of the existing aero-engine flame tube;

[0019] Figure 2 It is a schematic diagram of the wall structure of the aero-engine flame tube provided by the embodiment of the present application;

[0020] Figure 3 It is a schematic diagram of the circumferential and axial distributions of the impingement cooling holes, turbulators and film holes in the wall structure of the aero-engine flame tube provided by the embodiment of the present application;

[0021] Wherein:

[0022] 1 - Outer wall sector segment; 2 - Inner wall sector segment; 3 - Outer sector connecting segment;

[0023] A - Impingement cooling hole;

[0024] B - Turbulator;

[0025] C - Film hole.

[0026] For better illustration of this embodiment, some components in the drawings are omitted, enlarged or reduced, which do not represent the dimensions of the actual product. In addition, the drawings are only for illustrative purposes and should not be construed as a limitation of this patent. Detailed implementation manners

[0027] To make the technical solutions and their advantages of the present application clearer, the technical solutions of the present application will be further described clearly and completely in conjunction with the drawings. It can be understood that the specific embodiments described herein are only partial embodiments of the present application, which are only used to explain the present application and not to limit the present application. It should be noted that for the convenience of description, only the parts related to the present application are shown in the drawings, and other related parts can refer to the general design. Without conflict, the embodiments in the present application and the technical features in the embodiments can be combined with each other to obtain new embodiments.

[0028] In addition, unless otherwise defined, the technical terms or scientific terms used in the description of this application shall have the ordinary meanings understood by those of ordinary skill in the art to which this application pertains. The words indicating directions such as "upper", "lower", "left", "right", "center", "vertical", "horizontal", "inner", "outer", etc. used in the description of this application are only used to indicate relative directions or positional relationships, rather than implying that the device or component must have a specific orientation, be constructed and operated in a specific orientation. When the absolute position of the object being described changes, its relative positional relationship may also change accordingly. Therefore, it should not be construed as a limitation to this application. The terms "first", "second", "third" and similar terms used in the description of this application are only for descriptive purposes to distinguish different components, and should not be construed as indicating or implying relative importance. The similar words such as "a", "an" or "the" used in the description of this application should not be construed as an absolute limitation on the quantity, but should be understood as having at least one. The similar words such as "including" or "comprising" used in the description of this application are intended to mean that the elements or objects appearing before this word cover the elements or objects listed after this word and their equivalents, without excluding other elements or objects.

[0029] In addition, it should be noted that, unless otherwise clearly specified and limited, the similar words such as "installed", "connected" and "joined" used in the description of this application should be understood in a broad sense. For example, the connection can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be directly connected or indirectly connected through an intermediate medium, and can also be the communication inside two components. Those skilled in the art can understand its specific meaning in this application according to the specific situation.

[0030] The following will further elaborate on this application Figures 1 to 3 in conjunction with the appended

[0031] A wall structure of an aero-engine flame tube, as Figure 2 shown, includes:

[0032] A plurality of outer wall fan-shaped segments 1, which have a plurality of impingement cooling holes A on their side walls and a plurality of turbulator columns B formed on the inner side;

[0033] A plurality of inner wall fan-shaped segments 2, which have a plurality of film holes C formed on their side walls, and the outer sides of which are connected to the turbulator columns B on each of the outer wall fan-shaped segments 1 by diffusion welding or brazing; the two side edges of each inner wall fan-shaped segment 2 are welded by electron beam welding to form an annular inner wall; there is a connection gap between adjacent outer wall fan-shaped segments 1;

[0034] A plurality of outer fan-shaped connecting segments 3 are arranged in each connecting gap. There are a plurality of impingement cooling holes A on its side wall, a plurality of turbulator columns B are formed on the inner side, and outward flanges are provided at both side edges; there is a small gap between the turbulator columns B on each outer fan-shaped connecting segment 3 and the annular inner wall; the outward flanges on both side edges of each outer fan-shaped connecting segment 3 overlap on the outer side of the corresponding adjacent outer wall fan-shaped segment 1 and are connected to the outer side of the corresponding adjacent outer wall fan-shaped segment 1 by brazing or electron beam welding, and are positioned by a rabbet therebetween, forming an annular outer wall with each outer wall fan-shaped segment 1.

[0035] For the aero-engine flame tube wall structure disclosed in the above embodiments, those skilled in the art can understand that its designed annular inner wall is formed by electron beam welding between the two side edges of a plurality of inner wall fan-shaped segments 2, and the annular outer wall is composed of a plurality of outer wall fan-shaped segments 1 and outer fan-shaped connecting segments 3 distributed at intervals. The outward flanges on both side edges of each outer fan-shaped connecting segment 3 overlap on the outer side of the adjacent outer wall fan-shaped segment 1 and are connected by brazing or electron beam welding. The welding connection part between each outer fan-shaped connecting segment 3 and the outer wall fan-shaped segment 1 is located on the outside and will not affect the position distribution of the impingement cooling holes A and turbulator columns B thereon. Corresponding to the circumferential joint surface position between the two side edges of each inner wall fan-shaped segment 2, a plurality of impingement cooling holes A and their turbulator columns B are distributed on the side wall of each outer fan-shaped segment 3, which can ensure the cooling effect and avoid ablation.

[0036] In some alternative embodiments, in the above aero-engine flame tube wall structure, the impingement cooling holes A and turbulator columns B on the annular outer wall and the film holes C on the inner wall fan-shaped segment 2 are distributed alternately in the circumferential and axial directions to ensure the cooling effect, as Figure 3 shown.

[0037] In some alternative embodiments, in the above aero-engine flame tube wall structure, the small gap between the turbulator columns B on each outer fan-shaped connecting segment 3 and the annular inner wall is not greater than 0.1 mm and should not be too large, so as to avoid interference during assembly.

[0038] In some alternative embodiments, in the above aero-engine flame tube wall structure, each outer fan-shaped connecting segment 3 includes at least one row of impingement cooling holes A and two rows of turbulator columns B. The two rows of turbulator columns B can be distributed on both sides of the circumferential joint surface between the corresponding inner wall fan-shaped segments 2 to ensure the cooling effect.

[0039] The embodiments in the specification are described in a progressive manner. The key point of each embodiment is to illustrate the differences from other embodiments. The same or similar parts between the embodiments can be referred to each other.

[0040] So far, the technical solutions of this application have been described in conjunction with the preferred embodiments shown in the accompanying drawings. Those skilled in the art should understand that the protection scope of this application is obviously not limited to these specific embodiments. Without departing from the principle of this application, those skilled in the art can make equivalent changes or substitutions to relevant technical features, and the technical solutions after these changes or substitutions will all fall within the protection scope of this application.

Claims

1. A wall structure of an aero-engine combustion chamber, characterized in that Comprising: A plurality of outer wall sector segments (1), having a plurality of impingement cooling holes (A) on their side walls and a plurality of turbulator posts (B) formed on the inner side; A plurality of inner wall sector segments (2), having a plurality of film holes (C) formed on their side walls, and being connected to the turbulator posts (B) on each outer wall sector segment (1) by diffusion welding or brazing; the two side edges of each inner wall sector segment (2) are welded by electron beam welding to form an annular inner wall; there is a connection gap between adjacent outer wall sector segments (1); A plurality of outer sector connecting segments (3), arranged in each connection gap, having a plurality of impingement cooling holes (A) on their side walls, a plurality of turbulator posts (B) formed on the inner side, and outward flanges at both side edges; there is a small gap between the turbulator posts (B) on each outer sector connecting segment (3) and the annular inner wall; the outward flanges on both side edges of each outer sector connecting segment (3) overlap the outside of the corresponding adjacent outer wall sector segment (1) and are connected to the outside of the corresponding adjacent outer wall sector segment (1) by brazing or electron beam welding to form an annular outer wall with each outer wall sector segment (1); Each outer sector connecting segment (3) includes at least one row of impingement cooling holes (A) and two rows of turbulator posts (B); The two rows of turbulator posts (B) are distributed on both sides of the circumferential joint surface between the corresponding inner wall sector segments (2).

2. The wall structure of an aeroengine combustion chamber according to claim 1, characterized in that The impingement cooling holes (A) and turbulator posts (B) on the annular outer wall and the film holes (C) on the inner wall sector segments (2) are distributed alternately in the circumferential and axial directions.

3. The wall structure of an aeroengine combustion chamber according to claim 1, characterized in that The small gap between the turbulator posts (B) on each outer sector connecting segment (3) and the annular inner wall is not greater than 0.1 mm.

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