An aeroengine combustion chamber

Abstract

An aero-engine combustion chamber comprises: a combustion chamber outer wall having a plurality of mounting holes; a combustion chamber inner wall arranged in the combustion chamber outer wall; a flame tube arranged between the combustion chamber outer wall and the combustion chamber inner wall; a plurality of flame tube heads in the shape of a sector, each having a plurality of longitudinal and transverse intersecting circumferential channels and radial channels, and a plurality of air inlet holes formed thereon; each air inlet hole is distributed in an area surrounded by adjacent circumferential channels and radial channels, and a side wall has two fuel injection holes communicating with two adjacent circumferential channels; each flame tube head is connected to the inlet of the flame tube; a plurality of fuel nozzles are mounted in each mounting hole, and a plurality of groups of nozzle distribution pipelines are connected to the circumferential channels and / or radial channels in each flame tube head; a plurality of annular fuel supply manifolds are sleeved on the outer periphery of the combustion chamber outer wall, and each annular fuel supply manifold is connected to each fuel nozzle through a plurality of supply manifold distribution pipelines, and a corresponding group of nozzle distribution pipelines is communicated.

Description

TECHNICAL FIELD

[0001] The application belongs to the technical field of aero-engine combustion chamber design, and particularly relates to an aero-engine combustion chamber. BACKGROUND

[0002] The aero-engine uses hydrocarbon as fuel, which is easy to produce a large amount of greenhouse gases and causes serious impact on the environment. Therefore, an aero-engine using hydrogen as fuel is designed. However, the high temperature generated after the combustion of hydrogen increases the emission amount of pollutants such as nitrogen oxides. At present, in order to reduce the emission amount of pollutants of the hydrogen fuel aero-engine, a multi-point injection combustion chamber is designed to have a more uniform combustion temperature field, so as to reduce the emission of pollutants.

[0003] In the existing multi-point injection combustion chamber, one flame tube head corresponds to several or even dozens of fuel nozzles, which has a complex structure and a large weight. In addition, each fuel nozzle is designed to be separately supplied with gas, which greatly increases the leakage risk and is difficult to manufacture and assemble.

[0004] The present application is proposed in view of the existence of the above technical defects.

[0005] It should be noted that the disclosure of the above background art is only used to assist in understanding the inventive concept and technical solutions of the present application, and it does not necessarily belong to the prior art of the present patent application. In the absence of explicit evidence that the above content has been disclosed on the filing date of the present application, the above background art should not be used to evaluate the novelty and inventiveness of the present application. SUMMARY

[0006] The purpose of the present application is to provide an aero-engine combustion chamber to overcome or alleviate at least one aspect of the known technical defects.

[0007] The technical solution of the present application is:

[0008] An aero-engine combustion chamber comprises:

[0009] A combustion chamber outer wall having a plurality of mounting holes distributed circumferentially thereon;

[0010] A combustion chamber inner wall arranged in the combustion chamber outer wall;

[0011] A flame tube arranged between the combustion chamber outer wall and the combustion chamber inner wall;

[0012] A plurality of flame tube heads in the shape of a sector, having a plurality of longitudinal and transverse circumferential channels and radial channels therein, and a plurality of air inlet holes formed thereon; each air inlet hole is distributed in a region surrounded by adjacent circumferential channels and radial channels, and the side wall has two fuel injection holes communicating with adjacent two circumferential channels; each flame tube head is connected at the inlet of the flame tube and is spliced to form a whole ring structure;

[0013] Multiple fuel nozzles are installed in various mounting holes and connected to the circumferential and / or radial channels in each flame tube head through multiple sets of nozzle distribution pipelines.

[0014] Multiple annular fuel supply manifolds are fitted around the outer perimeter of the combustion chamber. Each annular fuel supply manifold is connected to various fuel nozzles via multiple supply manifold distribution pipes, and is correspondingly connected to a set of nozzle distribution pipes.

[0015] According to at least one embodiment of this application, in the above-mentioned aero-engine combustion chamber, each radial channel of each flame tube head is symmetrically distributed, the radial channel in the middle is connected to the nozzle distribution pipeline, and the connection part is located at the intersection with the circumferential channel.

[0016] According to at least one embodiment of this application, in the above-described aero-engine combustion chamber, each radial channel is radially cut into multiple flow control sections, and each flow control section is connected to a nozzle distribution pipeline.

[0017] According to at least one embodiment of this application, in the aforementioned aircraft engine combustion chamber, the routing of each nozzle distribution pipe avoids the front of each air intake.

[0018] According to at least one embodiment of this application, in the above-mentioned aero-engine combustion chamber, each flame tube head is integrally formed using an additive manufacturing process.

[0019] According to at least one embodiment of this application, the aforementioned aircraft engine combustion chamber further includes:

[0020] The diffuser is connected at the inlet between the outer wall of the combustion chamber and the inner wall of the combustion chamber. Attached Figure Description

[0021] Figure 1 This is a schematic diagram of an aircraft engine combustion chamber provided in an embodiment of this application;

[0022] Figure 2 This is a perspective view of a portion of the combustion chamber structure of an aero-engine provided in an embodiment of this application;

[0023] Figure 3 yes Figure 2 A three-dimensional view of the middle section structure;

[0024] Figure 4 yes Figure 3 A partial sectional view;

[0025] Figure 5 yes Figure 4 Side view;

[0026] Figure 6 yes Figure 5 A partial sectional view;

[0027] wherein:

[0028] 1 - outer combustion chamber wall; 2 - inner combustion chamber wall; 3 - flame tube; 4 - flame tube head; 5 - fuel nozzle; 6 - nozzle distribution line; 7 - annular fuel supply header; 8 - supply header distribution line; 9 - diffuser;

[0029] A - inlet aperture;

[0030] B - fuel injection aperture.

[0031] In order to better illustrate the present embodiments, some components in the drawings can be omitted, enlarged or reduced, and do not represent the actual size of the product. In addition, the drawings are only used for illustrative description and cannot be understood as a limitation on the present patent. DETAILED DESCRIPTION

[0032] In order to make the technical solutions of the present application and the advantages thereof clearer, the technical solutions of the present application will be further clearly and completely described in detail below in conjunction with the drawings. It should be understood that the specific embodiments described herein are only some embodiments of the present application, which are used to explain the present application, but not to limit the present application. It should be noted that, for the purpose of description, only parts related to the present application are shown in the drawings, and other related parts can be referred to the general design. In the case of no conflict, the embodiments in the present application and the technical features in the embodiments can be combined to obtain new embodiments.

[0033] In addition, unless otherwise defined, the technical terms or scientific terms used in the present application description should be the general meaning understood by the general technical personnel in the field of the present application. The words indicating the relative direction or position relationship, such as "upper", "lower", "left", "right", "center", "vertical", "horizontal", "inner", "outer" and the like used in the present application description, are only used to indicate the relative direction or position relationship, and not to imply that the device or element must have a specific orientation, be constructed and operated in a specific orientation, and the relative position relationship may also change accordingly when the absolute position of the described object changes, therefore, it cannot be understood as a limitation on the present application. The "first", "second", "third" and similar terms used in the present application description are only for the purpose of description, to distinguish different components, and cannot be understood as indicating or implying relative importance. The "one", "an" or "the" and similar terms used in the present application description should not be understood as an absolute limitation on the quantity, but should be understood as the existence of at least one. The "include" or "contain" and similar terms used in the present application description mean that the elements or objects appearing before the word cover the elements or objects listed after the word and their equivalents, and do not exclude other elements or objects.

[0034] In addition, it needs to be explained that, unless otherwise explicitly specified and limited, the "installation", "connection", "connection" and similar words used in the description of the application should be understood broadly, for example, the connection can be fixed connection, or detachable connection, or integrally connected; can be mechanical connection, or electrical connection; can be directly connected, or indirectly connected through intermediate medium, or the internal communication of two elements, and those skilled in the art can understand the specific meaning of the application according to the specific circumstances.

[0035] The following will be combined with the attached Figures 1 to 6 The application is further described in detail.

[0036] An aero-engine combustion chamber, comprising:

[0037] The combustion chamber outer wall 1 has a plurality of circumferentially distributed mounting holes;

[0038] The combustion chamber inner wall 2 is arranged in the combustion chamber outer wall 1;

[0039] The flame tube 3 is arranged between the combustion chamber outer wall 1 and the combustion chamber inner wall 2;

[0040] A plurality of flame tube heads 4 are arranged in the form of a fan, which have a plurality of longitudinal and transverse circumferential channels and radial channels, and a plurality of air inlet holes A are formed on the side wall; each air inlet hole A is distributed in the area surrounded by adjacent circumferential channels and radial channels, and the side wall has two fuel injection holes B which communicate with adjacent two circumferential channels; each flame tube head 4 is connected to the inlet of the flame tube 3 and is spliced to form a whole ring structure;

[0041] A plurality of fuel nozzles 5 are installed in each mounting hole, and the circumferential channels and / or radial channels in each flame tube head 4 are connected through a plurality of nozzle distribution pipelines 6;

[0042] A plurality of annular fuel supply manifolds 7 are arranged outside the outer wall 1 of the combustion chamber, and each annular fuel supply manifold 7 is connected to a plurality of supply manifold distribution pipelines 8 which connect each fuel nozzle 5 and correspondingly connect a group of nozzle distribution pipelines 6;

[0043] The diffuser 9 is connected to the inlet between the combustion chamber outer wall 1 and the combustion chamber inner wall 2.

[0044] The aero-engine combustion chamber disclosed in the above embodiment, when working, the high-pressure air from the compressor can flow into the inlet of the combustion chamber outer wall 1 and the combustion chamber inner wall 2 through the diffuser 8, part of which flows into the space between the combustion chamber outer wall 1 and the outer wall of the flame tube 3, and part of which flows into the space between the combustion chamber inner wall 2 and the inner wall of the flame tube 3. The two parts of the airflow flow circumferentially backward, and can cool the inner and outer walls of the flame tube 3 along the way, and then discharge into the turbine, and the remaining part of the high-pressure air flows into the inside of the flame tube 3 through the air inlet holes A of the flame tube head.

[0045] The aero-engine combustion chamber disclosed in the above embodiment can, in operation, deliver fuel to each annular fuel supply manifold 7 through the fuel delivery pipeline, and the fuel can be distributed by the supply manifold distribution pipeline 8, the fuel nozzle 5, the nozzle distribution pipeline 6 and the circumferential and radial passages thereof, and then enter the corresponding air inlet hole A through the fuel injection hole, mix with the compressed air in the air inlet hole A, and be injected into the flame tube 3 for combustion. The high-temperature gas generated after combustion is discharged from the tail of the flame tube 3 to drive the turbine to work. The fuel mentioned above can be hydrogen.

[0046] For the aero-engine combustion chamber disclosed in the above embodiment, those skilled in the art can understand that the design of the combustion chamber adopts a multi-point injection combustion chamber, which can have a more uniform combustion temperature field, can reduce the emission of pollutants, and can integrate the air inlet and fuel supply functions in the flame tube head 4 through the channel, so that the overall structure is compact, easy to manufacture and assemble, and less likely to leak. In addition, it is convenient to control the fuel supply in different areas, and the heat sink of the fuel can be used to cool the flame tube head 4, reduce the demand for cooling gas for cooling the flame tube head 4, and preheat the fuel to promote more uniform and complete combustion.

[0047] In some optional embodiments, in the aero-engine combustion chamber described above, the radial passages of each flame tube head 4 are symmetrically distributed, the radial passage in the middle is in communication with the nozzle distribution pipeline 6, and the communication part is located at the intersection with the circumferential passage, so as to enhance the symmetry of fuel supply and make the fuel supply to the combustion chamber uniform, so as to have a more uniform combustion temperature field.

[0048] In some optional embodiments, in the aero-engine combustion chamber described above, each radial passage is truncated into a plurality of flow control sections along the radial direction, each flow control section is in communication with one nozzle distribution pipeline 6, nozzle distribution pipelines 6 with the same fuel flow control requirement form a group, and one annular fuel supply manifold 7 supplies fuel, which facilitates the control of fuel supply in different areas, so as to have a more uniform combustion temperature field and reduce the emission of pollutants.

[0049] In some optional embodiments, in the aero-engine combustion chamber described above, the direction of each nozzle distribution pipeline 6 avoids the front of each air inlet hole A, so as to avoid the influence of the nozzle distribution pipeline 6 on the air inlet of the local air inlet hole 4 of the flame tube 3.

[0050] In some optional embodiments, in the aero-engine combustion chamber described above, each flame tube head 4 is integrally formed by additive manufacturing process.

[0051] The various embodiments described in the specification are intended to be exemplary only and are not intended to limit the scope of the application. The scope of the application is defined by the claims.

[0052] The technical solutions of the present application have been described in combination with the preferred embodiments shown in the drawings. It should be understood by those skilled in the art that the protection scope of the present application is obviously not limited to these specific embodiments. Those skilled in the art can make equivalent changes or replacements to the related technical features without departing from the principles of the present application. The technical solutions after the changes or replacements will fall within the protection scope of the present application.

Claims

1. An aircraft engine combustion chamber, characterized in that, include: The outer wall of the combustion chamber (1) has multiple mounting holes distributed circumferentially; The combustion chamber wall (2) is installed inside the combustion chamber wall (1); A flame tube (3) is installed between the outer wall (1) of the combustion chamber and the inner wall (2) of the combustion chamber; Multiple flame tube heads (4) are fan-shaped and have multiple crisscrossing circumferential and radial channels inside. Multiple air inlets (A) are formed on them. Each air inlet (A) is distributed in the area enclosed by adjacent circumferential and radial channels. The side wall has two fuel injection holes (B) that connect to two adjacent circumferential channels. Each flame tube head (4) is connected to the inlet of the flame tube (3) and is spliced ​​together to form a complete ring structure. Multiple fuel nozzles (5) are installed in various mounting holes and connected to the circumferential and / or radial channels in each flame tube head (4) through multiple sets of nozzle distribution pipes (6); Multiple annular fuel supply manifolds (7) are fitted around the outer periphery of the combustion chamber wall (1). Each annular fuel supply manifold (7) is connected to each fuel nozzle (5) through multiple supply manifold distribution pipes (8), and is connected to a set of nozzle distribution pipes (6). Each flame tube head (4) has a radial channel symmetrically distributed. The radial channel in the middle is connected to the nozzle distribution pipeline (6), and the connection point is located at the intersection with the circumferential channel. Each radial channel is cut into multiple flow control segments along the radial direction, and each flow control segment is connected to a nozzle distribution pipeline (6).

2. The aero-engine combustion chamber according to claim 1, characterized in that, The routing of each nozzle distribution pipe (6) avoids the front of each air inlet (A).

3. The aero-engine combustion chamber according to claim 1, characterized in that, Each flame tube head (4) is integrally formed using additive manufacturing process.

4. The aero-engine combustion chamber according to claim 1, characterized in that, Also includes: The diffuser (9) is connected at the inlet between the outer wall (1) of the combustion chamber and the inner wall (2) of the combustion chamber.

Images