An aero-engine core engine air bleed cooling structure

Abstract

The application relates to an aero-engine core engine bleed air cooling structure. A cooling path is constructed by utilizing a combustion chamber outer casing, a combustion chamber inner casing, a combustion chamber flame tube, a combustion chamber inlet diffuser, a bleed air return gas collecting cover and a transmission shaft connecting a turbine first-stage rotor disc and a compressor last-stage rotor disc, and the return gas is utilized to cool turbine front-end rotor blades. The overall structure is simple and compact, an additional bleed air pipe does not need to be arranged on the outside to cool the turbine through compressor inter-stage bleed air, a deformation compensation structure does not need to be additionally designed, the sealing is reliable, gas leakage and other problems are not prone to occur, in addition, the bleed air return cooling channel directly performing bleed air can be designed in a ring shape, the ring-shaped bleed air return cooling channel does not affect the circumferential uniformity of the bleed air position flow field, and the return gas can be expanded and reduced in speed in the bleed air return gas collecting cavity and the bleed air return gas diffuser cavity during the bleed air process, so that the return gas is uniformly mixed and the stability of the return gas cooling can be ensured.

Description

Technical Field

[0001] This application belongs to the field of bleed air cooling design technology for aero-engine core engines, specifically relating to a bleed air cooling structure for aero-engine core engines. Background Technology

[0002] The core of an aero-engine consists of a compressor, a combustion chamber, and a turbine connected in sequence. As the thrust requirements of aero-engines increase, the temperature at the turbine inlet rises significantly. To protect the turbine from erosion, an bleed air cooling structure is designed, which uses bleed air from the compressor to cool the turbine.

[0003] Currently, the bleed air cooling structure of aero-engine cores utilizes multiple circumferentially distributed bleed air pipes to cool the turbine by drawing bleed air from the compressor interstage to the turbine inlet. This technical solution has the following drawbacks:

[0004] 1) Multiple air intake pipes distributed in the circumferential direction will occupy a large radial space, which will increase the radial dimension of the aero engine as a whole, resulting in an increase in the overall mass of the aero engine, which does not meet the current requirements for weight reduction of aero engines.

[0005] 2) Multiple air intake pipes are independent of each other in the circumferential direction, which seriously affects the uniformity of the flow field in the circumferential direction. Therefore, a larger air collection chamber needs to be designed to uniform pressure, which will further increase the radial dimension of the aero engine and increase the overall mass of the aero engine.

[0006] 3) Using multiple circumferentially distributed bleed pipes to bleed air from the outside of the compressor stage to the turbine inlet can only effectively cool the turbine front guide vanes, but it is difficult to cool the turbine front rotor blades.

[0007] 4) In order to compensate for the uncoordinated deformation between the compressor and the turbine, the middle part of each air intake pipeline is designed to use a floating connection and use a sealing ring for sealing. However, the sealing effect is poor under high temperature, and the sealing part is easily damaged and fails when the uncoordinated deformation between the compressor and the turbine is large.

[0008] This application is made in view of the aforementioned technical deficiencies.

[0009] 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

[0010] The purpose of this application is to provide a core air bleed cooling structure for an aero-engine to overcome or mitigate at least one of the known technical defects.

[0011] The technical solution of this application is:

[0012] A core bleed air cooling structure for an aero-engine includes:

[0013] Air compressor;

[0014] The combustion chamber includes an outer combustion chamber casing, an inner combustion chamber casing, a combustion chamber flame tube, and a combustion chamber inlet diffuser. The front end of the outer combustion chamber casing is connected to the rear end of the outer compressor casing via a connecting edge. The inner combustion chamber casing is disposed inside the outer combustion chamber casing. The combustion chamber flame tube is disposed between the outer combustion chamber casing and the inner combustion chamber casing. The front end of the outer wall of the combustion chamber inlet diffuser is connected to the rear end of the outer compressor casing, and the rear end of the outer wall is connected to the front end of the outer combustion chamber casing via a connecting edge. The front end of the inner wall is connected to the rear end of the inner compressor casing, and the rear end of the inner wall is connected to the front end of the inner combustion chamber casing via a connecting edge.

[0015] The turbine's outer casing is connected to the rear end of the outer casing of the combustion chamber via a connecting edge, and its inner casing is connected to the rear end of the inner casing of the combustion chamber via a connecting edge. Its first-stage rotor disc is connected to the last-stage rotor disc of the compressor via a drive shaft.

[0016] in,

[0017] The combustion chamber has an exhaust gas recirculation cooling channel inside, and the inlet of the exhaust gas recirculation cooling channel extends to the inner rear end of the combustion chamber.

[0018] The outer wall of the front end of the combustion chamber has multiple circumferentially distributed bleed air return cooling holes, and each bleed air return cooling hole is connected to the bleed air return cooling channel.

[0019] The bleed air cooling structure for the aero-engine core also includes:

[0020] Multiple bleed air return and collection hoods are circumferentially connected to the outer wall of the front end of the outer casing of the combustion chamber, forming multiple bleed air return and collection chambers between them and the outer casing of the combustion chamber; each bleed air return and collection chamber is connected to each bleed air return cooling hole.

[0021] The front side wall of the combustion chamber has multiple circumferentially distributed bleed air return flow holes, and each bleed air return flow hole is connected to each bleed air return collection chamber.

[0022] A bleed air return diffuser cavity is formed between the inner side of the front end of the outer casing of the combustion chamber and the outer wall of the diffuser at the combustion chamber inlet. The bleed air return diffuser cavity is connected to each bleed air return flow hole.

[0023] The outer and inner walls of the combustion chamber inlet diffuser are supported by multiple circumferentially arranged support plates, each of which has an air bleed return flow channel. The inlet of each air bleed return flow channel is formed on the outer wall of the combustion chamber inlet diffuser and communicates with the air bleed return diffuser chamber. The outlet of each air bleed return flow channel is formed on the inner wall of the combustion chamber inlet diffuser.

[0024] A bleed air return channel is formed between the combustion chamber casing and the drive shaft, and the bleed air return channel connects to the outlet of each bleed air return flow channel.

[0025] According to at least one embodiment of this application, the above-described aero-engine core bleed air cooling structure further includes:

[0026] Multiple drag reducers are installed in the bleed air return diffuser chamber and connected to the inner side of the front end of the outer casing of the combustion chamber, communicating with each bleed air return flow hole.

[0027] According to at least one embodiment of this application, in the above-described aero-engine core bleed air cooling structure, multiple reinforcing ribs distributed circumferentially are formed inside the inlet end of the bleed air return cooling channel.

[0028] According to at least one embodiment of this application, in the above-mentioned aero-engine core engine bleed air cooling structure, the outer casing side wall of the combustion chamber has a plurality of fuel nozzle mounting holes along the circumferential direction, and a fuel nozzle mounting seat is formed around each fuel nozzle mounting hole.

[0029] The core air cooling structure of the aero-engine also includes:

[0030] Multiple fuel nozzles are connected to each fuel nozzle mounting base via connecting edges. The fuel injection end extends into the outer casing of the combustion chamber through the fuel nozzle mounting hole and is installed on the head of the combustion chamber flame tube.

[0031] This application has at least the following beneficial technical effects:

[0032] This invention provides a core bleed air cooling structure for an aero-engine. The design utilizes the outer combustor casing, inner combustor casing, combustor flame tube, combustor inlet diffuser, bleed air return shroud, and drive shaft connecting the turbine first-stage rotor disk and the compressor last-stage rotor disk to construct a cooling path. The return air cools the turbine front-end rotor blades. The overall structure is simple and compact, eliminating the need for additional bleed air pipes on the outside for compressor-stage bleed air cooling of the turbine. It does not occupy a large radial space, reducing the overall radial dimension and mass of the aero-engine. Furthermore, it eliminates the need for additional deformation compensation structures, ensuring reliable sealing and minimizing the risk of gas leakage. In addition, the bleed air return cooling channel, which directly receives the return air, can be designed in a ring shape, avoiding any impact on the circumferential uniformity of the flow field at the bleed air location. During the bleed air process, the return air is repeatedly diffused and decelerated using the bleed air return shroud and bleed air return diffuser, ensuring uniform mixing and stable cooling of the return air. Attached Figure Description

[0033] Figure 1 This is a schematic diagram of the bleed air cooling structure of the aero-engine core provided in the embodiments of this application;

[0034] Figure 2 This is a schematic diagram of the combustion chamber, its bleed air return hood, and drag reducer provided in the embodiments of this application;

[0035] Figure 3 yes Figure 2 A sectional view along the axial direction;

[0036] Figure 4 yes Figure 3 Partial sectional view from direction II;

[0037] in:

[0038] 1-Compressor; 2-Combustion chamber; 3-Turbine; 4-Drive shaft; 5-Bleed air return shroud; 6-Drag reducer; 7-Reinforcing rib; 8-Fuel nozzle;

[0039] 21-Outer combustion chamber casing; 22-Inner combustion chamber casing; 23-Combustion chamber flame tube; 24-Combustion chamber inlet diffuser;

[0040] A-Air evacuation and collection chamber;

[0041] B-Air evacuation and diffuser chamber;

[0042] C - Air reflux circulation channel;

[0043] D-Air reflux channel;

[0044] E-Fuel injector mounting hole;

[0045] F - Ignition nozzle mounting hole;

[0046] G-Test Hole;

[0047] H-lead hole.

[0048] 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

[0049] 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.

[0050] 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.

[0051] 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.

[0052] The following is in conjunction with the appendix Figures 1 to 4 This application will be described in further detail.

[0053] A core bleed air cooling structure for an aero-engine includes:

[0054] Compressor 1;

[0055] Combustion chamber 2 includes an outer combustion chamber casing 21, an inner combustion chamber casing 22, a combustion chamber flame tube 23, and a combustion chamber inlet diffuser 24. The front end of the outer combustion chamber casing 21 is connected to the rear end of the outer casing of compressor 1 via a connecting edge. The inner combustion chamber casing 22 is disposed inside the outer combustion chamber casing 21. The combustion chamber flame tube 23 is disposed between the outer combustion chamber casing 21 and the inner combustion chamber casing 22. The front end of the outer wall of the combustion chamber inlet diffuser 24 is connected to the rear end of the outer casing of compressor 1, and the rear end of the outer wall is connected to the front end of the outer combustion chamber casing 21 via a connecting edge. The front end of the inner wall is connected to the rear end of the inner casing of compressor 1, and the rear end of the inner wall is connected to the front end of the inner combustion chamber casing 22 via a connecting edge.

[0056] The turbine 3 has its outer casing front end connected to the rear end of the combustion chamber outer casing 21 via a connecting edge, its inner casing front end connected to the rear end of the combustion chamber inner casing 22 via a connecting edge, and its first-stage rotor disc connected to the last-stage rotor disc of the compressor 1 via a drive shaft 4.

[0057] in,

[0058] The combustion chamber 21 has an exhaust gas recirculation cooling channel inside, the inlet end of which extends to the inner side of the rear end of the combustion chamber 21, and the height of the exhaust gas recirculation cooling channel is not less than 1.5 mm.

[0059] The outer wall of the front end of the combustion chamber 21 has multiple circumferentially distributed bleed air return cooling holes, and each bleed air return cooling hole is connected to the bleed air return cooling channel.

[0060] The bleed air cooling structure for the aero-engine core also includes:

[0061] Multiple bleed air return and collection hoods 5 are circumferentially connected to the outer wall of the front end of the combustion chamber 21, forming multiple bleed air return and collection chambers A between them and the combustion chamber 21; each bleed air return and collection chamber A is connected to each bleed air return cooling hole.

[0062] The front side wall of the combustion chamber 21 has multiple circumferentially distributed bleed air return flow holes, and each bleed air return flow hole is connected to each bleed air return collection chamber A;

[0063] A bleed air return diffuser chamber B is formed between the inner side of the front end of the outer casing 21 of the combustion chamber and the outer wall of the diffuser 24 at the combustion chamber inlet. The bleed air return diffuser chamber B is connected to each bleed air return flow hole.

[0064] The outer and inner walls of the combustion chamber inlet diffuser 24 are supported by multiple circumferentially arranged support plates, each of which has an air bleed return flow channel C. The inlet of each air bleed return flow channel C is formed on the outer wall of the combustion chamber inlet diffuser 24 and communicates with the air bleed return diffuser chamber B. The outlet of each air bleed return flow channel C is formed on the inner wall of the combustion chamber inlet diffuser 24.

[0065] A bleed air return guide channel D is formed between the casing 22 and the drive shaft 4 in the combustion chamber, and the bleed air return guide channel D is connected to the outlet of each bleed air return flow channel C.

[0066] Regarding the bleed air cooling structure for the aero-engine core engine disclosed in the above embodiments, those skilled in the art will understand that during operation, the outlet airflow of compressor 1 enters combustion chamber 2 via combustion chamber inlet diffuser 24. This airflow can be divided into two paths. One path enters the space between the inner walls of the combustion chamber casing 22 and the combustion chamber flame tube 23. This airflow can enter the combustion chamber flame tube 23 through the air inlet on the inner wall of the combustion chamber flame tube 23 for combustion and cooling within the combustion chamber flame tube 23. The other path enters the space between the outer walls of the combustion chamber casing 21 and the combustion chamber flame tube 23. This airflow can partially enter the space between the outer walls of the combustion chamber flame tube 23 through the air inlet on the outer wall of the combustion chamber flame tube 23. The airflow enters the combustion chamber flame tube 23 for combustion and cooling. A portion of the remaining airflow can be diverted into the turbine 3 to cool the turbine 3 front guide vane, while another portion enters the induced draft air return cooling channel. It then enters the induced draft air return collecting chamber A through various induced draft air return flow holes, flows into the induced draft air return diffuser chamber B through various induced draft air return flow channels C, and then flows into the induced draft air return guide channel D as return air to the turbine 3 first-stage rotor disk. This allows for further cooling flow path design for the turbine 3 first-stage rotor disk and its rotor blades, achieving efficient cooling of the turbine front rotor blades.

[0067] Regarding the bleed air cooling structure for the aero-engine core engine disclosed in the above embodiments, those skilled in the art will understand that its design utilizes the outer combustion chamber casing 21, the inner combustion chamber casing 22, the combustion chamber flame tube 23, the combustion chamber inlet diffuser 24, the bleed air return shroud 5, and the drive shaft 4 connecting the first-stage rotor disk of the turbine 3 and the last-stage rotor disk of the compressor 1 to construct a cooling path. The return air is used to cool the turbine's front-end rotor blades. The overall structure is simple and compact, eliminating the need for additional bleed air pipes on the outside for compressor interstage bleed air cooling of the turbine. It does not occupy a large radial space, which can reduce the radial size of the aero-engine and reduce the overall weight of the aero-engine. It does not require the design of additional deformation compensation structure, and the sealing is reliable, making it less prone to gas leakage and other problems. In addition, the bleed air return cooling channel that directly bleeds the return air can be designed in a ring shape, which will not affect the circumferential uniformity of the flow field at the bleed air position. During the bleed air process, the bleed air return collecting chamber A and the bleed air return diffuser chamber B can be used to diffuse and decelerate the return air multiple times, so as to mix it evenly and ensure the stability of the return air cooling.

[0068] In some optional embodiments, the above-described aero-engine core bleed air cooling structure further includes:

[0069] Multiple drag reducers 6 are installed in the bleed air return diffuser chamber B and connected to the inner front end of the outer casing 21 of the combustion chamber. They are connected to each bleed air return flow hole and can decelerate the airflow through a specific aerodynamic shape when the return air flows into the bleed air return diffuser chamber B, thereby reducing the sudden diffusion loss.

[0070] In some optional embodiments, in the above-mentioned aero-engine core bleed air cooling structure, the drag reducer 6 can be designed to reduce the airflow velocity by no less than 50%. The drag reducer 6 can be designed with equal pressure gradient or equal velocity gradient along the airflow direction. A guide vane can be set inside to meet the requirement of a greater velocity reduction ratio and ensure that the airflow does not undergo significant flow separation.

[0071] In some optional embodiments, in the above-mentioned bleed air cooling structure of the aero-engine core, multiple reinforcing ribs 7 are formed in the inlet end of the bleed air return cooling channel, which are distributed circumferentially. The number of reinforcing ribs 7 is not less than 10, the thickness is not less than 2 mm, the length is not less than 10 mm, and they extend along the aero-engine axis. This can divide the inlet of the bleed air return collecting chamber A into multiple independent areas, rectify the return air entering the bleed air return collecting chamber A, and enable the return air to flow parallel to the aero-engine axis to reduce flow loss.

[0072] In some optional embodiments, in the above-described aero-engine core bleed air cooling structure, the side wall of the combustion chamber 21 has a plurality of fuel nozzle mounting holes E along the circumferential direction, and a fuel nozzle mounting seat is formed around each fuel nozzle mounting hole E.

[0073] The core air cooling structure of the aero-engine also includes:

[0074] Multiple fuel nozzles 8 are connected to each fuel nozzle mounting base via connecting edges. The fuel injection end extends into the outer casing 21 of the combustion chamber through the fuel nozzle mounting hole E and is installed at the head of the combustion chamber flame tube 23 so as to supply fuel into the combustion chamber flame tube 23 for combustion.

[0075] In some optional embodiments, in the above-mentioned bleed air cooling structure of the aero-engine core engine, the side wall of the combustion chamber outer casing 21 may also be provided with functional holes such as ignition nozzle mounting hole F, test hole G, and lead wire hole H, and surrounded by a mounting base for mounting the ignition nozzle, test sensor, and lead wire, etc. The mounting base can be locally reinforced and thickened to improve the overall rigidity of the combustion chamber outer casing 21 and prevent harmful vibration.

[0076] In some optional embodiments, in the above-mentioned bleed air cooling structure of the aero-engine core, components such as the combustion chamber outer casing 21, the bleed air return shroud 5, and the drag reducer 6 can be integrally formed by metal additive manufacturing process. Machining allowances are reserved at positions with high dimensional accuracy requirements, such as connecting edges and mounting seats. After integral forming and hot isostatic pressing treatment, the final finishing is carried out.

[0077] 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.

[0078] 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. A core bleed air cooling structure for an aero-engine, characterized in that, include: Compressor (1); Combustion chamber (2) includes an outer combustion chamber casing (21), an inner combustion chamber casing (22), a combustion chamber flame tube (23), and a combustion chamber inlet diffuser (24). The front end of the outer combustion chamber casing (21) is connected to the rear end of the outer casing of the compressor (1) via a connecting edge. The inner combustion chamber casing (22) is installed inside the outer combustion chamber casing (21). The combustion chamber flame tube (23) is installed between the outer combustion chamber casing (21) and the inner combustion chamber casing (22). The front end of the outer wall of the combustion chamber inlet diffuser (24) is connected to the rear end of the outer casing of the compressor (1), and the rear end of the outer wall is connected to the front end of the outer combustion chamber casing (21) via a connecting edge. The front end of the inner wall is connected to the rear end of the inner casing of the compressor (1), and the rear end of the inner wall is connected to the front end of the inner combustion chamber casing (22) via a connecting edge. The turbine (3) has its outer casing front end connected to the combustion chamber outer casing (21) rear end via a connecting edge, its inner casing front end connected to the combustion chamber inner casing (22) rear end via a connecting edge, and its first-stage rotor disk connected to the compressor (1) last-stage rotor disk via a drive shaft (4). in, The combustion chamber (21) has an exhaust air recirculation cooling channel, the inlet of which extends to the inner rear end of the combustion chamber (21); The outer wall of the combustion chamber (21) at the front end has multiple circumferentially distributed bleed air return cooling holes, and each bleed air return cooling hole is connected to the bleed air return cooling channel. The bleed air cooling structure for the aero-engine core also includes: Multiple bleed air return and collection hoods (5) are connected circumferentially to the outer wall of the front end of the combustion chamber (21), forming multiple bleed air return and collection chambers (A) between them and the combustion chamber (21); each bleed air return and collection chamber (A) is connected to each bleed air return cooling hole; The front side wall of the combustion chamber (21) has multiple circumferentially distributed bleed air return flow holes, and each bleed air return flow hole is connected to each bleed air return collection chamber (A). A bleed air return diffuser chamber (B) is formed between the inner front end of the outer casing (21) of the combustion chamber and the outer wall of the diffuser (24) at the combustion chamber inlet. The bleed air return diffuser chamber (B) is connected to each bleed air return passage. The outer and inner walls of the combustion chamber inlet diffuser (24) are supported by multiple circumferentially arranged support plates, each of which has an air bleed return flow channel (C); the inlet of each air bleed return flow channel (C) is formed on the outer wall of the combustion chamber inlet diffuser (24) and communicates with the air bleed return diffuser chamber (B); the outlet of each air bleed return flow channel (C) is formed on the inner wall of the combustion chamber inlet diffuser (24); A bleed air return guide channel (D) is formed between the combustion chamber casing (22) and the drive shaft (4), and the bleed air return guide channel (D) is connected to the outlet of each bleed air return flow channel (C).

2. The bleed air cooling structure for the core engine of an aero-engine according to claim 1, characterized in that, Also includes: Multiple drag reducers (6) are installed in the bleed air return diffuser chamber (B), connected to the inner side of the front end of the outer casing (21) of the combustion chamber, and connected to each bleed air return flow hole.

3. The bleed air cooling structure for the core engine of an aero-engine according to claim 1, characterized in that, Multiple reinforcing ribs (7) are formed inside the inlet end of the air duct return cooling channel.

4. The bleed air cooling structure for the core engine of an aero-engine according to claim 1, characterized in that, The combustion chamber (21) has multiple fuel nozzle mounting holes (E) along the circumferential direction on its side wall, and a fuel nozzle mounting seat is formed around each fuel nozzle mounting hole (E); The core air cooling structure of the aero-engine also includes: Multiple fuel nozzles (8) are connected to each fuel nozzle mounting base via connecting edges. The fuel injection end extends into the outer casing (21) of the combustion chamber through the fuel nozzle mounting hole (E) and is installed at the head of the combustion chamber flame tube (23).

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