An aeroengine combustion chamber component and a connection structure therebetween

Abstract

The application belongs to the technical field of aero-engine combustion chamber component and its connecting structure with a compressor, and particularly relates to an aero-engine combustion chamber component and its connecting structure with a compressor, which comprises a combustion chamber outer wall, a combustion chamber inner wall arranged in the combustion chamber outer wall, a compressor last-stage stator assembly, an annular connecting edge between the outer wall rear end of the compressor last-stage stator assembly and the front end of the combustion chamber outer wall, a connecting edge between the inner wall rear end of the compressor last-stage stator assembly and the front end of the combustion chamber inner wall, and a diffuser made of a light-weight high-temperature-resistant high-specific-strength low-plasticity material, arranged in the combustion chamber outer wall, supported by the combustion chamber outer wall or the combustion chamber inner wall through bolt connection, and connected to the combustion chamber outer wall or the combustion chamber inner wall through the connecting edge.

Description

Technical Field

[0001] This application belongs to the technical field of aero-engine combustion chamber components and their connection structure with the compressor, specifically relating to an aero-engine combustion chamber component and its connection structure with the compressor. Background Technology

[0002] To address the dual requirements of high temperature resistance and lightweight design, diffusers in aero-engine combustion chambers are currently mostly made of lightweight, high-temperature resistant, high-strength, low-plasticity materials, such as γ-TiAl intermetallic compounds. However, these materials have poor weldability and are highly sensitive to notches. In current aero-engines, the diffuser is connected between the inner and outer walls of the combustion chamber and the last-stage stator of the compressor. The connection between the diffuser and the inner and outer walls of the combustion chamber is mostly supported by welding, and it needs to withstand the transmission of axial forces between the inner and outer walls of the combustion chamber. This makes it prone to damage at stress concentration points and surface defects, and the damage is irreparable.

[0003] This application is made in view of the aforementioned technical deficiencies.

[0004] It should be noted that the above background information is only used to assist in understanding the inventive concept and technical solution of this invention, and it does not necessarily belong to the prior art of this application. In the absence of clear evidence that the above information was disclosed on the filing date of this application, the above background information should not be used to evaluate the novelty and inventiveness of this application. Summary of the Invention

[0005] The purpose of this application is to provide an aircraft engine combustion chamber component and its connection structure with the compressor, so as to overcome or mitigate at least one of the known technical defects.

[0006] The technical solution of this application is:

[0007] An aero-engine combustion chamber component and its connection structure with the compressor, comprising:

[0008] The outer wall of the combustion chamber;

[0009] The inner wall of the combustion chamber is installed inside the outer wall of the combustion chamber;

[0010] The rear end of the compressor stator assembly is connected to the front end of the outer wall of the combustion chamber via an annular connecting edge, and the rear end of its inner wall is connected to the front end of the inner wall of the combustion chamber via a connecting edge.

[0011] The diffuser is made of a lightweight, high-temperature resistant, high-strength, low-plasticity material. It is installed inside the outer wall of the combustion chamber and located inside the inlet of the outer wall of the combustion chamber. It is supported by bolts to the outer wall of the combustion chamber or the inner wall of the combustion chamber via connecting edges. The front end of its inner wall is connected to the rear end of the inner wall of the last stage stator assembly of the compressor with a stop fit, and the front end of its outer wall is connected to the rear end of the outer wall of the last stage stator assembly of the compressor with a stop fit.

[0012] According to at least one embodiment of this application, in the above-mentioned aero-engine combustion chamber component and its connection structure with the compressor, the outer rear end of the compressor final stage stator assembly is welded to the inner front end of the combustion chamber outer wall via an annular connecting edge.

[0013] The outer rear end of the inner wall of the compressor's final stage stator assembly is welded to the front end of the combustion chamber wall via a connecting edge.

[0014] The inner wall of the combustion chamber is divided into two sections, front and rear, which are connected by bolts through an outward annular connecting edge.

[0015] The front end of the diffuser inner wall and the rear end of the compressor last stage stator assembly inner wall are fitted with a small clearance in the radial direction. The maximum clearance in the cold state is no more than 0.05 mm, and centering is performed thereon.

[0016] The front end of the diffuser outer wall and the rear end of the compressor last stage stator assembly outer wall are fitted with a large clearance in the radial direction. The maximum clearance in the cold state is not less than 1 mm.

[0017] A W-type elastic sealing ring is installed at the stop position of the front end of the diffuser outer wall and the rear end of the compressor last stage stator assembly. The W-type elastic sealing ring is pre-compressed axially to ensure the seal at the stop position of the front end of the diffuser outer wall and the rear end of the compressor last stage stator assembly. The pre-compression amount of the W-type elastic sealing ring is not greater than 15% of the working stroke. There is an axial gap between the front end of the diffuser inner wall and outer wall and the rear end of the compressor last stage stator assembly. This gap is not greater than 50% of the working stroke of the W-type elastic sealing ring.

[0018] The outer rear end of the diffuser outer wall and the inner front end of the combustion chamber outer wall are connected by bolts through a connecting edge. The nut at this point is a plate nut, which is connected to the connecting edge on the inner front end of the combustion chamber outer wall. The corresponding bolt hole on the connecting edge on the outer rear end of the diffuser outer wall is a runway-shaped hole. The long side of the runway-shaped hole is parallel to the radial direction of the aero-engine, and the length of the long side is at least 2 mm greater than the mean diameter of the corresponding bolt.

[0019] According to at least one embodiment of this application, the aforementioned aero-engine combustion chamber component and its connection structure with the compressor further include:

[0020] The flame tube is installed between the outer wall of the combustion chamber and the inner wall of the combustion chamber. The rear end is axially floatingly connected to the front end of the turbine guide. The inner wall of the turbine guide is supported by bearings on the inner side of the rear end of the inner wall of the combustion chamber through the connecting edge.

[0021] The front cover of the outer wall of the flame tube is connected to the front end of the outer wall of the flame tube;

[0022] The inner wall front cover of the flame tube is connected to the front end of the inner wall of the flame tube and bends towards each other with the outer wall front cover of the flame tube.

[0023] Multiple fuel swirl nozzles are arranged circumferentially at the head of the flame tube, and their fuel injection rods are set through the outer wall of the flame tube and the outer wall of the combustion chamber, and are connected to the outer wall of the combustion chamber by a mounting bracket.

[0024] Multiple ignition nozzles are arranged circumferentially on the flame tube, and their support rods are installed through the outer wall of the combustion chamber and connected to the outer wall of the combustion chamber by mounting brackets.

[0025] According to at least one embodiment of this application, in the above-mentioned aero-engine combustion chamber component and its connection structure with the compressor, there is a radial overlap between the rear end of the outer wall front cover of the flame tube and the front end of the outer wall of the flame tube.

[0026] There is a radial overlap between the rear end of the inner wall of the flame tube and the front end of the inner wall of the flame tube.

[0027] The aero-engine combustion chamber component and its connection structure with the compressor also include:

[0028] Multiple flame tube outer wall positioning rods are installed through the outer wall of the combustion chamber, the rear end of the flame tube outer wall front cover and the front end of the flame tube outer wall in the radial overlapping part, distributed circumferentially, and positioning bushings are installed at the radial overlapping part between the flame tube outer wall front cover and the front end of the flame tube outer wall.

[0029] Multiple positioning rods for the inner wall of the combustion chamber are installed in the radial overlapping area between the inner wall of the combustion chamber, the rear end of the inner wall cover of the combustion chamber and the front end of the inner wall of the combustion chamber, and are distributed circumferentially. Positioning bushings are installed in the radial overlapping area between the inner wall cover of the combustion chamber and the front end of the inner wall of the combustion chamber.

[0030] According to at least one embodiment of this application, in the above-mentioned aero-engine combustion chamber component and its connection structure with the compressor, a support seat is formed on the inner side of the outer wall of the combustion chamber, which surrounds the positioning rods of the outer wall of each flame tube, and the outer contour of the support seat along the rearward axis of the aero-engine is a subsonic airfoil.

[0031] The inner wall of the combustion chamber is formed with a support seat that surrounds the positioning rods of the outer wall of each flame tube. The outer contour of the support seat along the rearward axis of the aero-engine is a subsonic airfoil.

[0032] According to at least one embodiment of this application, in the above-mentioned aero-engine combustion chamber component and its connection structure with the compressor, the diffuser is a multi-channel diffuser, with a flow divider ring provided between its outer wall and inner wall, and supported by multiple support plates along the circumferential direction, wherein each support plate has a notch on its rear edge, which is engaged with the front edge of the flow divider ring.

[0033] According to at least one embodiment of this application, in the above-mentioned aero-engine combustion chamber component and its connection structure with the compressor, the outer rear end of the compressor final stage stator assembly is welded to the inner front end of the combustion chamber outer wall via an annular connecting edge.

[0034] The outer rear end of the inner wall of the compressor's final stage stator assembly is welded to the front end of the combustion chamber wall via a connecting edge.

[0035] The inner wall of the combustion chamber is divided into two sections, front and rear, which are connected by bolts through an outward annular connecting edge.

[0036] The front end of the diffuser inner wall and the rear end stop of the compressor last stage stator assembly inner wall adopt a large clearance fit in the radial direction. The maximum cold-state fit clearance at this location is not less than 1 mm.

[0037] The front end of the diffuser outer wall and the rear end of the compressor last stage stator assembly outer wall are fitted with a large clearance in the radial direction. The maximum clearance in the cold state is not less than 1 mm.

[0038] W-type elastic sealing rings are installed at the stop positions of the inner and outer walls of the diffuser and the inner and outer walls of the last stage stator assembly of the compressor. The W-type elastic sealing rings are pre-compressed axially to ensure the sealing at the stop positions of the inner and outer walls of the diffuser and the last stage stator assembly of the compressor. The pre-compression amount of the W-type elastic sealing rings is not greater than 15% of the working stroke. There is an axial gap between the inner and outer walls of the diffuser and the inner and outer walls of the last stage stator assembly of the compressor. This gap is not greater than 50% of the working stroke of the W-type elastic sealing rings.

[0039] The outer rear end of the diffuser outer wall and the inner front end of the combustion chamber outer wall are connected by bolts through a connecting edge. The connecting edges are fitted with a stop, and the gap in the radial direction is no more than 0.05mm for centering. In addition, the nut at this point is a support nut, which is connected to the connecting edge on the inner front end of the combustion chamber outer wall. The corresponding bolt hole on the connecting edge on the outer rear end of the diffuser outer wall is a runway-shaped hole. The long side of the runway-shaped hole is parallel to the radial direction of the aero-engine, and the length of the long side is at least 2mm greater than the mean diameter of the corresponding bolt.

[0040] According to at least one embodiment of this application, in the above-mentioned aero-engine combustion chamber component and its connection structure with the compressor, the outer rear end of the compressor final stage stator assembly is welded to the inner rear end of the combustion chamber outer wall via an annular connecting edge.

[0041] The rear end of the inner wall of the compressor's last stage stator assembly, the front end of the inner wall of the combustion chamber, and the rear end of the diffuser's inner wall are connected by bolts through an outward annular connecting edge. At this point, the outward annular connecting edge of the diffuser's inner wall rear end and the outward annular connecting edge of the compressor's last stage stator assembly rear end and the combustion chamber's inner wall front end are fitted with a stop joint. The maximum interference in the radial direction of the fit clearance does not exceed 0.1 mm, and the maximum clearance does not exceed 0.025 mm, for centering purposes.

[0042] The front end of the diffuser inner wall and the rear end stop of the compressor last stage stator assembly inner wall adopt a large clearance fit in the radial direction. The maximum cold-state fit clearance at this location is not less than 1 mm.

[0043] The front end of the diffuser outer wall and the rear end of the compressor last stage stator assembly outer wall are fitted with a large clearance in the radial direction. The maximum clearance in the cold state is not less than 1 mm.

[0044] W-shaped elastic sealing rings are installed at the stop positions of the inner and outer walls of the diffuser and the inner and outer walls of the last stage stator assembly of the compressor. The W-shaped elastic sealing rings are pre-compressed axially to ensure the sealing at the stop positions of the inner and outer walls of the diffuser and the last stage stator assembly of the compressor. The pre-compression amount of the W-shaped elastic sealing rings is not greater than 15% of the working stroke. There is an axial gap between the inner and outer walls of the diffuser and the inner and outer walls of the last stage stator assembly of the compressor. This gap is not greater than 50% of the working stroke of the W-shaped elastic sealing rings. Attached Figure Description

[0045] Figures 1-2 This is a schematic diagram of an example 1 of the aero-engine combustion chamber component and its connection structure with the compressor provided in the embodiments of this application;

[0046] Figure 3 This is a schematic diagram of an example two of the aircraft engine combustion chamber components and their connection structure with the compressor provided in the embodiments of this application;

[0047] Figure 4 This is a schematic diagram of an example three of the aircraft engine combustion chamber components and their connection structure with the compressor provided in the embodiments of this application;

[0048] in:

[0049] 1-Outer wall of the combustion chamber; 2-Inner wall of the combustion chamber; 3-Last stage stator assembly of the compressor; 4-Diffuser; 5-W-type elastic sealing ring; 6-Flame tube; 7-Turbine guide; 8-Flame tube outer wall front cover; 9-Flame tube inner wall front cover; 10-Fuel swirl nozzle; 11-Ignition nozzle; 12-Flame tube outer wall positioning rod; 13-Flame tube inner wall positioning rod.

[0050] To better illustrate this embodiment, some parts in the accompanying drawings may be omitted, enlarged, or reduced, and do not represent the actual product dimensions. Furthermore, the drawings are for illustrative purposes only and should not be construed as limiting this application. Detailed Implementation

[0051] To make the technical solution and advantages of this application clearer, the technical solution of this application will be described in a clearer and more complete manner below with reference to the accompanying drawings. It should be understood that the specific embodiments described herein are only some embodiments of this application, and are only used to explain this application, not to limit this application. It should be noted that, for ease of description, only the parts related to this application are shown in the accompanying drawings. Other related parts can be referred to the general design. In the absence of conflict, the embodiments and technical features in the embodiments of this application can be combined with each other to obtain new embodiments.

[0052] Furthermore, unless otherwise defined, the technical or scientific terms used in this application description shall have the ordinary meaning understood by one of ordinary skill in the art to which this application pertains. The terms "upper," "lower," "left," "right," "center," "vertical," "horizontal," "inner," and "outer," etc., used in this application description to indicate relative direction or positional relationship are used only to indicate relative orientation or positional relationship, and do not imply that the device or component must have a specific orientation, or be constructed and operated in a specific orientation. When the absolute position of the described object changes, its relative positional relationship may also change accordingly, and therefore should not be construed as a limitation on this application. The terms "first," "second," "third," and similar terms used in this application description are used only for descriptive purposes to distinguish different components, and should not be construed as indicating or implying relative importance. The terms "a," "one," or "the," etc., used in this application description should not be construed as an absolute limitation on quantity, but should be construed as indicating the existence of at least one. The terms "including," "comprising," etc., used in this application description mean that the element or object preceding the word covers the element or object listed after the word and its equivalents, without excluding other elements or objects.

[0053] Furthermore, it should be noted that, unless otherwise explicitly specified and limited, terms such as “installation,” “connection,” and “linkage” used in the description of this application should be interpreted broadly. For example, a 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 a direct connection or an indirect connection through an intermediate medium; or it can be a connection within two components. Those skilled in the art can understand its specific meaning in this application according to the specific circumstances.

[0054] The following is in conjunction with the appendix Figures 1 to 4 This application will be described in further detail.

[0055] An aero-engine combustion chamber component and its connection structure with the compressor, comprising:

[0056] Combustion chamber exterior wall 1;

[0057] The inner wall 2 of the combustion chamber is installed inside the outer wall 1 of the combustion chamber;

[0058] The rear end of the compressor stator assembly 3 is connected to the front end of the outer wall of the combustion chamber outer wall 1 by an annular connecting edge, and the rear end of its inner wall is connected to the front end of the combustion chamber inner wall 2 by a connecting edge.

[0059] The diffuser 4 is made of a lightweight, high-temperature resistant, high-strength, low-plasticity material. It is installed inside the outer wall 1 of the combustion chamber and located inside the inlet of the outer wall 1. It is supported by bolts to the outer wall 1 or the inner wall 2 of the combustion chamber via connecting edges. The front end of its inner wall is connected to the rear end of the inner wall of the last stage stator assembly 3 of the compressor with a stop fit, and the front end of its outer wall is connected to the rear end of the outer wall of the last stage stator assembly 3 of the compressor with a stop fit.

[0060] Regarding the aero-engine combustion chamber components and their connection structure with the compressor disclosed in the above embodiments, those skilled in the art will understand that, in the case that the diffuser 4 is manufactured from a lightweight, high-temperature resistant, high-strength, low-plasticity material, the outer wall 1 and the inner wall 2 of the combustion chamber are connected through the stator of the compressor's final stage stator assembly 3 to transmit axial forces. The diffuser 4 is supported by bolts to the outer wall 1 or the inner wall 2 of the combustion chamber via connecting edges. The front end of the inner wall of the diffuser 4 is connected to the rear end of the inner wall of the compressor's final stage stator assembly 3 with a stop fit, and the front end of the outer wall is connected to the rear end of the outer wall of the compressor's final stage stator assembly 3 with a stop fit to achieve positioning. This avoids direct welding connections between the diffuser 4 and the outer wall 1 and the inner wall 2 of the combustion chamber, and prevents damage at stress concentration points or surface defect points.

[0061] The aero-engine combustion chamber components and their connection structure with the compressor disclosed in the above embodiments are designed as follows in Example 1: Figure 1 As shown, the outer rear end of the compressor stator assembly 3 is welded to the inner front end of the combustion chamber outer wall 1 via an annular connecting edge.

[0062] The outer rear end of the inner wall of the compressor final stage stator assembly 3 is welded to the front end of the combustion chamber inner wall 2 via a connecting edge.

[0063] The combustion chamber wall 2 is divided into two sections, front and rear, which are connected by bolts through an outward annular connecting edge;

[0064] The front end of the inner wall of the diffuser 4 and the rear end of the inner wall of the last stage stator assembly 3 of the compressor are fitted with a small clearance in the radial direction. The maximum clearance in the cold state is no more than 0.05mm, and centering is performed thereon.

[0065] The front end of the outer wall of the diffuser 4 and the rear end of the outer wall of the last stage stator assembly 3 of the compressor adopt a large clearance fit in the radial direction. The maximum clearance in the cold state is not less than 1 mm, so as to compensate for the radial deformation between the diffuser 4 and the last stage stator assembly 3 of the compressor.

[0066] A W-shaped elastic sealing ring 5 is installed at the stop between the front end of the outer wall of the diffuser 4 and the rear end of the outer wall of the last stage stator assembly 3 of the compressor. The W-shaped elastic sealing ring 5 is pre-compressed axially to ensure the seal at the stop between the front end of the outer wall of the diffuser 4 and the rear end of the outer wall of the last stage stator assembly 3 of the compressor. The pre-compression of the W-shaped elastic sealing ring 5 is not greater than 15% of the working stroke. There is an axial gap between the front end of the inner and outer walls of the diffuser 4 and the rear end of the inner and outer walls of the last stage stator assembly 3 of the compressor. This gap can compensate for the axial deformation between the diffuser 4 and the last stage stator assembly 3 of the compressor. The gap can be designed to be no greater than 50% of the working stroke of the W-shaped elastic sealing ring 5 to ensure that the compression of the W-shaped elastic sealing ring 5 does not exceed 50% of the working stroke and to avoid damage to the W-shaped elastic sealing ring 5.

[0067] The outer rear end of the diffuser 4 is connected to the inner front end of the outer wall of the combustion chamber 1 by bolts through a connecting edge. The nut at this point is a plate nut, which is connected to the connecting edge on the inner front end of the outer wall of the combustion chamber 1. The corresponding bolt hole on the connecting edge of the outer rear end of the diffuser 4 is a runway-shaped hole. The long side of the runway-shaped hole is parallel to the radial direction of the aero-engine to reduce the stress concentration at the bolt hole on the connecting edge of the outer rear end of the diffuser 4. The length of the long side is at least 2 mm greater than the mean diameter of the corresponding bolt to allow for radial deformation of the diffuser 4.

[0068] In Example 1, the disclosed aero-engine combustion chamber components and their connection structure with the compressor, such as... Figure 1 As shown, it also includes:

[0069] The flame tube 6 is set between the outer wall 1 of the combustion chamber and the inner wall 2 of the combustion chamber. Its rear end is axially floatingly connected to the front end of the turbine guide 7. The inner wall of the turbine guide 7 is supported by a pre-rotating nozzle on the inner side of the rear end of the inner wall 2 of the combustion chamber through the connecting edge.

[0070] The front cover 8 of the outer wall of the flame tube is connected to the front end of the outer wall of the flame tube 6;

[0071] The inner wall front cover 9 of the flame tube is connected to the front end of the inner wall of the flame tube 6 and is bent towards each other with the outer wall front cover 8 of the flame tube.

[0072] Multiple fuel swirl nozzles 10 are arranged circumferentially at the head of the flame tube 6, and their fuel injection rods are arranged through the front cover 8 of the outer wall of the flame tube and the outer wall 1 of the combustion chamber, and are connected to the outer wall 1 of the combustion chamber by means of a mounting bracket.

[0073] Multiple ignition nozzles 11 are arranged circumferentially on the flame tube 6, and their support rods are set through the outer wall 1 of the combustion chamber and connected to the outer wall 1 of the combustion chamber by means of a mounting base.

[0074] In Example 1, the disclosed aero-engine combustion chamber components and their connection structure with the compressor, such as... Figure 2 As shown, there is a radial overlap between the rear end of the outer wall front cover 8 of the flame tube and the front end of the outer wall of the flame tube 6.

[0075] There is a radial overlap between the rear end of the inner wall front cover 9 of the flame tube and the front end of the inner wall of the flame tube 6.

[0076] The aero-engine combustion chamber component and its connection structure with the compressor also include:

[0077] Multiple flame tube outer wall positioning rods 12 are provided through the radial overlap between the outer wall 1 of the combustion chamber, the rear end of the flame tube outer wall front cover 8 and the front end of the flame tube 6 outer wall, and are distributed circumferentially. Positioning bushings are provided at the radial overlap between the rear end of the flame tube outer wall front cover 8 and the front end of the flame tube 6 outer wall.

[0078] Multiple flame tube inner wall positioning rods 13 are installed through the radial overlap between the inner wall 2 of the combustion chamber, the rear end of the inner wall front cover 9 of the flame tube, and the front end of the inner wall of the flame tube 6. They are distributed circumferentially and positioning bushings are installed at the radial overlap between the inner wall front cover 9 of the flame tube and the front end of the inner wall of the flame tube 6. Thus, multiple flame tube outer wall positioning rods 12 and flame tube inner wall positioning rods 13 can be used to position the flame tube 6 at multiple points on the outer wall 1 of the combustion chamber and the inner wall 2 of the combustion chamber, while allowing a certain degree of radial deformation.

[0079] In Example 1, the disclosed aero-engine combustion chamber components and their connection structure with the compressor, such as... Figure 2 As shown, the inner side of the outer wall 1 of the combustion chamber is formed with a support seat that surrounds the positioning rod 12 of each flame tube outer wall. The outer contour of the support seat along the rearward axis of the aero-engine is a subsonic airfoil to reduce the pressure loss of the airflow.

[0080] The inner wall 2 of the combustion chamber is formed with a support seat that surrounds the positioning rod 13 of each flame tube. The outer contour of the support seat along the rearward axis of the aero-engine is a subsonic airfoil to reduce the pressure loss of the airflow.

[0081] The aero-engine combustion chamber components and their connection structure with the compressor disclosed in the above embodiments are designed as follows in Example 2: Figure 3As shown, the outer rear end of the compressor stator assembly 3 is welded to the inner front end of the combustion chamber outer wall 1 via an annular connecting edge.

[0082] The outer rear end of the inner wall of the compressor final stage stator assembly 3 is welded to the front end of the combustion chamber inner wall 2 via a connecting edge.

[0083] The combustion chamber wall 2 is divided into two sections, front and rear, which are connected by bolts through an outward annular connecting edge;

[0084] The front end of the inner wall of the diffuser 4 and the rear end of the inner wall of the last stage stator assembly 3 of the compressor adopt a large clearance fit in the radial direction. The maximum clearance in the cold state is not less than 1 mm, so as to compensate for the radial deformation between the diffuser 4 and the last stage stator assembly 3 of the compressor.

[0085] The front end of the outer wall of the diffuser 4 and the rear end of the outer wall of the last stage stator assembly 3 of the compressor adopt a large clearance fit in the radial direction. The maximum cold-state fit clearance is not less than 1mm to compensate for the radial deformation between the diffuser 4 and the last stage stator assembly 3 of the compressor.

[0086] A W-shaped elastic sealing ring 5 is installed at the stop positions of the inner and outer walls of the diffuser 4 and the inner and outer walls of the compressor's last stage stator assembly 3. The W-shaped elastic sealing ring 5 is pre-compressed axially to ensure the sealing at the stop positions of the inner and outer walls of the diffuser 4 and the compressor's last stage stator assembly 3. The pre-compression amount of the W-shaped elastic sealing ring 5 is not greater than 15% of the working stroke. There is also an axial gap between the inner and outer walls of the diffuser 4 and the inner and outer walls of the compressor's last stage stator assembly 3 to compensate for axial deformation between the diffuser 4 and the compressor's last stage stator assembly 3. This gap can be designed to be no greater than 50% of the working stroke of the W-shaped elastic sealing ring 5 to ensure that the compression amount of the W-shaped elastic sealing ring 5 does not exceed 50% of the working stroke and to avoid damage to the W-shaped elastic sealing ring 5.

[0087] The outer rear end of the diffuser 4 is bolted to the inner front end of the outer wall of the combustion chamber 1 via a connecting edge. The connecting edges have a stop fit with a radial clearance not exceeding 0.05mm for centering. Furthermore, the nut at this point is a support nut, attached to the connecting edge on the inner front end of the outer wall of the combustion chamber 1. The corresponding bolt hole on the outer rear end of the diffuser 4 is a runway-shaped hole, with its long side parallel to the radial direction of the aero-engine. This reduces stress concentration at the bolt hole on the outer rear end of the diffuser 4, and the long side is at least 2mm longer than the corresponding bolt's mean diameter to allow for radial deformation of the diffuser 4.

[0088] The aero-engine combustion chamber components and their connection structure with the compressor disclosed in the above embodiments are designed as follows in Example 3: Figure 4 As shown, the outer rear end of the compressor stator assembly 3 is welded to the inner rear end of the combustion chamber outer wall 1 via an annular connecting edge.

[0089] The rear end of the inner wall of the compressor final stage stator assembly 3, the front end of the combustion chamber inner wall 2, and the rear end of the inner wall of the diffuser 4 are connected by bolts through an outward annular connecting edge. At this point, the outward annular connecting edge of the rear end of the inner wall of the diffuser 4 and the outward annular connecting edge of the rear end of the inner wall of the compressor final stage stator assembly 3 and the front end of the combustion chamber inner wall 2 are fitted with a stop. The maximum interference in the radial direction of the fit clearance does not exceed 0.1 mm, and the maximum clearance does not exceed 0.025 mm, for centering purposes.

[0090] The front end of the inner wall of the diffuser 4 and the rear end of the inner wall of the last stage stator assembly 3 of the compressor adopt a large clearance fit in the radial direction. The maximum clearance in the cold state is not less than 1 mm, so as to compensate for the radial deformation between the diffuser 4 and the last stage stator assembly 3 of the compressor.

[0091] The front end of the outer wall of the diffuser 4 and the rear end of the outer wall of the last stage stator assembly 3 of the compressor adopt a large clearance fit in the radial direction. The maximum clearance in the cold state is not less than 1 mm, so as to compensate for the radial deformation between the diffuser 4 and the last stage stator assembly 3 of the compressor.

[0092] A W-shaped elastic sealing ring 5 is installed at the stop positions of the inner and outer walls of the diffuser 4 and the inner and outer walls of the compressor's last-stage stator assembly 3. The W-shaped elastic sealing ring 5 is pre-compressed axially to ensure sealing at the stop positions of the inner and outer walls of the diffuser 4 and the compressor's last-stage stator assembly 3. The pre-compression of the W-shaped elastic sealing ring 5 is no more than 15% of the working stroke. There is also an axial gap between the inner and outer walls of the diffuser 4 and the inner and outer walls of the compressor's last-stage stator assembly 3 to compensate for axial deformation between the diffuser 4 and the compressor's last-stage stator assembly 3. This gap can be designed to be no more than 50% of the working stroke of the W-shaped elastic sealing ring 5 to ensure that the compression of the W-shaped elastic sealing ring 5 does not exceed 50% of the working stroke, thus avoiding damage to the W-shaped elastic sealing ring 5.

[0093] In some optional embodiments, in the above-mentioned aero-engine combustion chamber components and their connection structure with the compressor, the diffuser 4 is a multi-channel diffuser, with a flow-dividing ring provided between its outer and inner walls, and supported by multiple support plates along the circumference, wherein each support plate has a notch on its rear edge, which is engaged with the front edge of the flow-dividing ring.

[0094] The various embodiments in the specification are described in a progressive manner, with each embodiment focusing on the differences from other embodiments. The same or similar parts between the various embodiments can be referred to each other.

[0095] The technical solution of this application has been described in conjunction with the preferred embodiments shown in the accompanying drawings. Those skilled in the art should understand that the scope of protection of this application is obviously not limited to these specific embodiments. Without departing from the principles of this application, those skilled in the art can make equivalent changes or substitutions to the relevant technical features, and the technical solutions after these changes or substitutions will all fall within the scope of protection of this application.

Claims

1. An aero-engine combustion chamber component and its connection structure with the compressor, characterized in that, include: outer wall of the combustion chamber (1); The combustion chamber wall (2) is installed inside the combustion chamber wall (1); The diffuser (4) is made of a lightweight, high-temperature resistant, high-strength, low-plasticity material and is installed inside the outer wall (1) of the combustion chamber, located inside the inlet of the outer wall (1) of the combustion chamber. The outer rear end of the compressor stator assembly (3) is welded to the inner front end of the combustion chamber outer wall (1) via an annular connecting edge. The outer rear end of the inner wall of the compressor final stage stator assembly (3) is welded to the front end of the combustion chamber inner wall (2) via a connecting edge; The combustion chamber wall (2) is divided into two sections, front and rear, which are connected by bolts through an outward annular connecting edge; The front end of the inner wall of the diffuser (4) and the rear end of the inner wall of the last stage stator assembly (3) of the compressor are fitted with a small clearance in the radial direction. The maximum clearance in the cold state is no more than 0.05 mm, and centering is performed. The front end of the outer wall of the diffuser (4) and the rear end of the outer wall of the last stage stator assembly (3) of the compressor adopt a large clearance fit in the radial direction. The maximum clearance in the cold state is not less than 1 mm. A W-type elastic sealing ring (5) is installed at the stop between the front end of the outer wall of the diffuser (4) and the rear end of the outer wall of the last stage stator assembly (3) of the compressor. The W-type elastic sealing ring (5) is pre-compressed axially to ensure the seal at the stop between the front end of the outer wall of the diffuser (4) and the rear end of the outer wall of the last stage stator assembly (3) of the compressor. The pre-compression amount of the W-type elastic sealing ring (5) is not greater than 15% of the working stroke. There is an axial gap between the front end of the inner and outer walls of the diffuser (4) and the rear end of the inner and outer walls of the last stage stator assembly (3) of the compressor. The gap is not greater than 50% of the working stroke of the W-type elastic sealing ring (5). The outer rear end of the diffuser (4) and the inner front end of the outer wall of the combustion chamber (1) are connected by bolts through the connecting edge. The nut is a plate nut, which is connected to the connecting edge of the inner front end of the outer wall of the combustion chamber (1). The corresponding bolt hole on the connecting edge of the outer rear end of the diffuser (4) is a runway-shaped hole. The long side of the runway-shaped hole is parallel to the radial direction of the aero-engine, and the length of the long side is at least 2 mm greater than the corresponding bolt's mean diameter.

2. The aero-engine combustion chamber component and its connection structure with the compressor according to claim 1, characterized in that, Also includes: The flame tube (6) is set between the outer wall (1) of the combustion chamber and the inner wall (2) of the combustion chamber. The rear end is axially floatingly connected to the front end of the turbine guide (7). The inner wall of the turbine guide (7) is supported by a bearing on the inner side of the rear end of the inner wall (2) of the combustion chamber through the connecting edge. The outer wall front cover (8) of the flame tube is connected to the front end of the outer wall of the flame tube (6); The inner wall front cover (9) of the flame tube is connected to the front end of the inner wall of the flame tube (6) and bends towards each other with the outer wall front cover (8) of the flame tube; Multiple fuel swirl nozzles (10) are arranged circumferentially at the head of the flame tube (6), and their fuel injection rods are arranged through the front cover (8) of the outer wall of the flame tube and the outer wall (1) of the combustion chamber, and are connected to the outer wall (1) of the combustion chamber by means of a mounting base; Multiple ignition nozzles (11) are arranged circumferentially on the flame tube (6), and their support rods are arranged through the outer wall (1) of the combustion chamber and connected to the outer wall (1) of the combustion chamber by means of a mounting base.

3. The aero-engine combustion chamber component and its connection structure with the compressor according to claim 2, characterized in that, There is a radial overlap between the rear end of the outer wall front cover (8) of the flame tube and the front end of the outer wall of the flame tube (6); There is a radial overlap between the rear end of the inner wall front cover (9) of the flame tube and the front end of the inner wall of the flame tube (6); The aero-engine combustion chamber component and its connection structure with the compressor also include: Multiple flame tube outer wall positioning rods (12) are provided in the radial overlapping part between the outer wall of the combustion chamber (1), the rear end of the flame tube outer wall front cover (8) and the front end of the flame tube (6), and are distributed in the circumferential direction. Positioning bushings are provided in the radial overlapping part between the flame tube outer wall front cover (8) and the front end of the flame tube (6). Multiple flame tube inner wall positioning rods (13) are provided in the radial overlapping part between the inner wall of the combustion chamber (2), the rear end of the flame tube inner wall front cover (9) and the front end of the flame tube (6), distributed in the circumferential direction, and positioning bushings are provided in the radial overlapping part between the flame tube inner wall front cover (9) and the front end of the flame tube (6).

4. The aero-engine combustion chamber component and its connection structure with the compressor according to claim 3, characterized in that, The inner side of the outer wall of the combustion chamber (1) is formed with a support seat that surrounds the positioning rod (12) of each flame tube outer wall. The outer contour of the support seat along the rearward axis of the aero-engine is a subsonic airfoil. The inner side of the combustion chamber wall (2) is formed with a support seat that surrounds the positioning rod (13) of each flame tube inner wall. The outer contour of the support seat along the rearward axis of the aero-engine is a subsonic airfoil.

5. The aero-engine combustion chamber component and its connection structure with the compressor according to claim 1, characterized in that, The diffuser (4) is a multi-channel diffuser with a flow divider ring between its outer and inner walls and multiple support plates along the circumference. Each support plate has a notch on its rear edge, which is stuck on the front edge of the flow divider ring.

6. An aero-engine combustion chamber component and its connection structure with the compressor, characterized in that, include: outer wall of the combustion chamber (1); The combustion chamber wall (2) is installed inside the combustion chamber wall (1); The diffuser (4) is made of a lightweight, high-temperature resistant, high-strength, low-plasticity material and is installed inside the outer wall of the combustion chamber (1) and located inside the inlet of the outer wall of the combustion chamber (1); the compressor final stage stator assembly (3) is welded between the outer rear end of the outer wall and the inner front end of the outer wall of the combustion chamber (1) through an annular connecting edge. The outer rear end of the inner wall of the compressor final stage stator assembly (3) is welded to the front end of the combustion chamber inner wall (2) via a connecting edge; The combustion chamber wall (2) is divided into two sections, front and rear, which are connected by bolts through an outward annular connecting edge; The front end of the inner wall of the diffuser (4) and the rear end of the inner wall of the last stage stator assembly (3) of the compressor adopt a large clearance fit in the radial direction. The maximum clearance in the cold state is not less than 1 mm. The front end of the outer wall of the diffuser (4) and the rear end of the outer wall of the last stage stator assembly (3) of the compressor adopt a large clearance fit in the radial direction. The maximum clearance in the cold state is not less than 1 mm. A W-type elastic sealing ring (5) is provided at the front end of the inner and outer walls of the diffuser (4) and the rear end of the inner and outer walls of the compressor final stage stator assembly (3). The W-type elastic sealing ring (5) is pre-compressed axially to ensure the sealing at the front end of the inner and outer walls of the diffuser (4) and the rear end of the inner and outer walls of the compressor final stage stator assembly (3). The pre-compression amount of the W-type elastic sealing ring (5) is not greater than 15% of the working stroke. There is an axial gap between the front end of the inner and outer walls of the diffuser (4) and the rear end of the inner and outer walls of the compressor final stage stator assembly (3). This gap is not greater than 50% of the working stroke of the W-type elastic sealing ring (5). The outer rear end of the diffuser (4) and the inner front end of the outer wall of the combustion chamber (1) are connected by bolts through the connecting edge. The connecting edge is fitted with a stop, and the gap is no more than 0.05mm in the radial direction for centering. The nut is a plate nut, which is connected to the connecting edge of the inner front end of the outer wall of the combustion chamber (1). The corresponding bolt hole on the connecting edge of the outer rear end of the diffuser (4) is a runway-shaped hole. The long side of the runway-shaped hole is parallel to the radial direction of the aero-engine, and the length of the long side is at least 2mm greater than the corresponding bolt's mean diameter.

7. An aero-engine combustion chamber component and its connection structure with the compressor, characterized in that, include: outer wall of the combustion chamber (1); The combustion chamber wall (2) is installed inside the combustion chamber wall (1); The diffuser (4) is made of a lightweight, high-temperature resistant, high-strength, low-plasticity material and is installed inside the outer wall (1) of the combustion chamber, located inside the inlet of the outer wall (1) of the combustion chamber. The outer rear end of the compressor stator assembly (3) is welded to the inner front end of the combustion chamber outer wall (1) via an annular connecting edge. The rear end of the inner wall of the compressor final stage stator assembly (3), the front end of the combustion chamber inner wall (2), and the rear end of the inner wall of the diffuser (4) are connected by bolts through an outward annular connecting edge. The outward annular connecting edge of the rear end of the inner wall of the diffuser (4) and the outward annular connecting edge of the rear end of the inner wall of the compressor final stage stator assembly (3) and the front end of the combustion chamber inner wall (2) are fitted with a stop. The maximum interference in the radial direction of the fit clearance does not exceed 0.1 mm, and the maximum clearance does not exceed 0.025 mm. Centering is performed. The front end of the inner wall of the diffuser (4) and the rear end of the inner wall of the last stage stator assembly (3) of the compressor adopt a large clearance fit in the radial direction. The maximum clearance in the cold state is not less than 1 mm. The front end of the outer wall of the diffuser (4) and the rear end of the outer wall of the last stage stator assembly (3) of the compressor adopt a large clearance fit in the radial direction. The maximum clearance in the cold state is not less than 1 mm. A W-type elastic sealing ring (5) is provided in the inner wall and outer wall front end of the diffuser (4) and the inner wall and outer wall rear end stop of the compressor final stage stator assembly (3). The W-type elastic sealing ring (5) is pre-compressed along the axial direction to ensure the sealing of the inner wall and outer wall front end of the diffuser (4) and the inner wall and outer wall rear end stop of the compressor final stage stator assembly (3). The pre-compression amount of the W-type elastic sealing ring (5) is not greater than 15% of the working stroke. There is a gap in the axial direction between the inner wall and outer wall front end of the diffuser (4) and the inner wall and outer wall rear end of the compressor final stage stator assembly (3). The gap is not greater than 50% of the working stroke of the W-type elastic sealing ring (5).

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