Aero-engine compressor disc cavity air bleed composite vortex reducer

Abstract

The application belongs to the technical field of aero-engine compressor disc cavity bleed air structure design, and particularly relates to aero-engine compressor disc cavity bleed air composite vortex reducer, wherein an outer edge part mounting ring and a disc core part mounting ring are arranged between two-stage rotor wheel discs to construct inner and outer two-layer cavities, and the airflow in the compressor flow channel can be introduced into the space between the two-stage rotor wheel discs through the bleed air holes on the drum, the airflow entering the space between the two-stage rotor wheel discs is affected by each counter-rotation nozzle, the flow direction is changed, the airflow enters the inner layer cavity, flows to the disc core through each vortex reduction pipe, and disc cavity bleed air is performed, and the counter-rotation nozzle and the vortex reduction pipe are respectively arranged at the outer edge part close to the rotor wheel disc and the disc core part to segmentally reduce the vortex, the advantages of the vortex reduction pipe and the counter-rotation nozzle are coupled in the inner and outer layer cavities, the generation of free vortex between the two-stage rotor wheel discs can be effectively reduced, and the pressure loss of the airflow can be reliably reduced.

Description

Technical Field

[0001] This application belongs to the technical field of air bleed structure design for compressor disks of aero-engines, and specifically relates to a composite vortex reducer for air bleed in compressor disks of aero-engines. Background Technology

[0002] The operation of an aero engine requires bleed air from the compressor flow channel to cool high-temperature components and seal the chambers.

[0003] In aero engines, air is drawn from the compressor flow channel, usually by using a disc cavity bleed method. An air bleed hole is opened on the drum between the two-stage rotor discs to draw the airflow from the compressor flow channel to the space between the two-stage rotor discs and then draw it out through the disc center.

[0004] In aero engines, air is drawn from the compressor flow channel using a disc cavity. The airflow in the compressor flow channel enters between the two-stage rotor discs through the air bleed holes on the drum. Due to the high-speed rotation of the rotor discs, complex vortices are generated. At the same time, it is affected by centrifugal force and Coriolis force, making it difficult to flow radially towards the disc center. If it is drawn out through the disc center, the pressure loss is large, which makes it difficult to meet the requirements for cooling high-temperature components and sealing the chamber.

[0005] Currently, the following two methods are commonly used to reduce pressure loss when using disc-cavity bleed air in the compressor flow path of aero engines:

[0006] 1) The pressure loss is reduced by using the anti-rotation nozzle vortex reducer. It is easy to install. The anti-rotation nozzle changes the rotation direction of the airflow between the two-stage rotor discs, reducing the generation of vortices and reducing the pressure loss of the disc cavity induced air. However, the anti-rotation nozzle is close to the drum and far from the disc center. After the airflow passes through the anti-rotation nozzle, it flows towards the disc center, which still produces a relatively complex flow and the drag reduction effect is unstable.

[0007] 2) Using tubular vortex reducers to reduce pressure loss involves installing vortex reducers at the center of the two-stage rotor disk. These vortex reducers direct airflow towards the center, reducing pressure loss in the disk cavity. However, the vortex reducers are located far from the drum. Airflow entering the space between the two-stage rotor disks through the bleed holes on the drum generates free vortices before reaching the vortex reducers. Furthermore, there is a significant angle between the flow direction entering the vortex reducers and the inlet cross-section of the vortex reducers, resulting in substantial pressure loss. Even with longer vortex reducers, achieving ideal drag reduction is difficult, and the overall mass of the aero-engine increases, potentially causing vibrations, threatening normal operation, and even posing a danger.

[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 composite vortex reducer for the compressor disk cavity of an aero-engine, in order to overcome or mitigate at least one of the known technical defects.

[0011] The technical solution of this application is:

[0012] A compressor disk cavity bleed air composite vortex suppressor for an aero-engine includes:

[0013] The two-stage rotor disk has an annular stop edge at the center of the front stage rotor disk on its rear side; the rear stage rotor disk has an annular stop edge at the center of the rear stage rotor disk and an annular stop edge at the outer edge of the rear stage rotor disk on its front side; the annular stop edge at the center of the rear stage rotor disk has an outward annular folded edge at the center of the rear stage rotor disk.

[0014] The drum has multiple circumferentially distributed air vents. Its rear end is formed on the front side of the rear stage rotor disk near the outer edge, and its front end points to the rear side of the front stage rotor disk. It has an inward annular fold at the front end of the drum.

[0015] An outer edge mounting ring is set between two stage rotor discs. Its rear outer side abuts against the inner side of the annular stop edge of the outer edge of the rear stage, and its front end has an outward annular mounting edge of the outer edge. The outward annular mounting edge of the outer edge extends into the space between the inward annular folded edge of the front stage rotor disc and the front end of the drum, and is connected to the inward annular folded edge of the front stage rotor disc and the front end of the drum by bolts.

[0016] Multiple counter-rotating nozzles are circumferentially mounted on an outer edge mounting ring; the outer edge mounting ring has localized recesses around each counter-rotating nozzle.

[0017] The disk center mounting ring has multiple circumferentially distributed disk center mounting holes and is set between the two-stage rotor disks. Its front outer side abuts against the inner side of the annular stop edge of the front stage disk center, and its rear outer side abuts against the inner side of the annular stop edge of the rear stage disk center. The rear outer side has an outward annular mounting edge of the disk center. The outward annular mounting edge of the disk center and the outward annular folded edge of the rear stage disk center are connected by bolts.

[0018] Multiple anti-vortex tubes are installed in the mounting holes of each disk center, with the inlet end pointing to each anti-rotation nozzle, and the outer wall of the outlet end having an annular stop edge at the outlet end of the anti-vortex tube; the annular stop edge at the outlet end of each anti-vortex tube abuts against the inner side of the mounting ring at the disk center.

[0019] According to at least one embodiment of this application, in the above-described aero-engine compressor disk cavity bleed air composite vortex reducer, each anti-rotation nozzle is directly opposite each bleed air hole.

[0020] According to at least one embodiment of this application, in the above-mentioned aero-engine compressor disk cavity bleed air composite vortex reducer, the local recesses surrounding each anti-rotation nozzle on the outer edge of the mounting ring are conical or spherical.

[0021] According to at least one embodiment of this application, in the above-mentioned aero-engine compressor disk cavity bleed air composite vortex reducer, the outer edge mounting ring is integrally formed with each anti-rotation nozzle.

[0022] According to at least one embodiment of this application, in the above-mentioned aero-engine compressor disk cavity bleed air composite vortex reducer, each vortex reducer tube is a conduit of equal diameter.

[0023] According to at least one embodiment of this application, in the above-mentioned aero-engine compressor disk cavity bleed air composite vortex reducer, the outer annular mounting edge of the outer edge, the front stage rotor disk, and the inner annular folded edge of the drum front end are centered and positioned using a stop.

[0024] According to at least one embodiment of this application, in the above-mentioned aero-engine compressor disk cavity bleed air composite vortex reducer, the outer annular mounting edge of the disk center has an outward flange, which presses against the outer annular folded edge of the rear stage disk center.

[0025] This application has at least the following beneficial technical effects:

[0026] This invention provides a composite vortex suppressor for aero-engine compressor disks. The design incorporates an outer edge mounting ring and a center mounting ring to create two chambers between two-stage rotor disks. Airflow from the compressor channel is introduced into the space between the two-stage rotor disks through bleed holes on the drum. In the outer chamber, the airflow is redirected by anti-rotation nozzles and flows into the inner chamber, where it is guided through vortex suppressor tubes to the center for vortex suppression. Near the outer edge and center of the rotor disks, anti-rotation nozzles and vortex suppressor tubes respectively provide segmented vortex suppression. The vortex suppressor tubes and anti-rotation nozzles in the inner and outer chambers are coupled to effectively reduce the generation of free vortices between the two-stage rotor disks. This reliably reduces pressure loss and meets the requirements for cooling high-temperature components and sealing the chambers during aero-engine operation. Furthermore, shorter vortex suppressor tubes, not exceeding one-third the length of the rotor disk, can be used, avoiding an increase in the overall mass of the aero-engine and preventing vibration that could threaten its normal operation. Attached Figure Description

[0027] Figure 1This is a schematic diagram of the air-induced vortex reducer for the compressor disk cavity of an aero-engine provided in the embodiments of this application;

[0028] Figure 2 This is a schematic diagram of the outer edge mounting ring and its reverse-rotation nozzle provided in an embodiment of this application;

[0029] Figure 3 This is a schematic diagram of the vortex reducer provided in an embodiment of this application;

[0030] Figure 4 This is a schematic diagram of the mounting ring and its anti-vortex tube at the center of the disk provided in the embodiments of this application;

[0031] in:

[0032] 1-Rotor disc; 2-Drum; 3-Outer edge mounting ring; 4-Reverse swirl nozzle; 5-Disc center mounting ring; 6-Vortex reducer;

[0033] A- Circular stop edge at the center of the front stage disc;

[0034] B-Ring stop edge at the rear center of the plate;

[0035] C-Rear stage outer edge ring stop edge;

[0036] D-The outward circular fold at the rear center of the plate;

[0037] F - Inward annular fold at the front end of the drum.

[0038] 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

[0039] 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.

[0040] 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.

[0041] 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.

[0042] The following is in conjunction with the appendix Figures 1 to 4 This application will be described in further detail.

[0043] A compressor disk cavity bleed air composite vortex suppressor for an aero-engine includes:

[0044] A two-stage rotor disk 1, wherein the rear side of the front stage rotor disk has an annular stop edge A at the center of the front stage disk; the front side of the rear stage rotor disk has an annular stop edge B at the center of the rear stage disk and an annular stop edge C at the outer edge of the rear stage disk; the annular stop edge B at the center of the rear stage disk has an outward annular folded edge D at the center of the rear stage disk.

[0045] The drum 2 has multiple circumferentially distributed air vents, and its rear end is formed on the front side of the rear stage rotor disk near the outer edge, while its front end points to the rear side of the front stage rotor disk. It has an inward annular fold F at the front end of the drum.

[0046] An outer edge mounting ring 3 is installed between the two-stage rotor discs 1. Its rear outer side abuts against the inner side of the annular stop edge C of the outer edge of the rear stage, and its front end has an outward annular mounting edge of the outer edge. The outward annular mounting edge of the outer edge extends into the space between the front stage rotor disc and the inward annular folded edge F of the front end of the drum, and is connected to the front stage rotor disc and the inward annular folded edge F of the front end of the drum by bolts. A stop can be used for centering and positioning.

[0047] Multiple anti-rotating nozzles 4 are circumferentially mounted on the outer edge mounting ring 3, and can be integrally formed with the outer edge mounting ring 3; the outer side of the outer edge mounting ring 3 is partially recessed around each anti-rotating nozzle 4, in the shape of a cone or a spherical socket.

[0048] The disk center mounting ring 5 has multiple disk center mounting holes distributed circumferentially. It is set between the two-stage rotor disks 1. Its front outer side abuts against the inner side of the annular stop edge A of the front stage disk center, and its rear outer side abuts against the inner side of the annular stop edge B of the rear stage disk center. The rear outer side has an outward annular mounting edge of the disk center. The outward annular mounting edge of the disk center and the outward annular folded edge D of the rear stage disk center are connected by bolts.

[0049] Multiple anti-vortex tubes 6 are equal diameter conduits, installed in the mounting holes of each disk center, with their inlet ends pointing to each anti-vortex nozzle 4, and each anti-vortex nozzle 4 can be directly facing each air intake hole.

[0050] Each vortex reducer tube 6 has an annular stop edge at its outlet end on its outer wall; the annular stop edge at the outlet end of each vortex reducer tube is attached to the inner side of the mounting ring 5 at the center of the disk.

[0051] Regarding the aero-engine compressor disk cavity bleed air composite vortex suppressor disclosed in the above embodiments, those skilled in the art will understand that, with the outer edge mounting ring 3 and the disk center mounting ring 5 constructing inner and outer chambers between the two-stage rotor disks 1, the airflow in the compressor flow channel can be introduced into the space between the two-stage rotor disks 1 through the bleed air holes on the drum 2. The airflow entering the space between the two-stage rotor disks 1 is affected by the various anti-rotation nozzles 4 in the outer chamber, changing its flow direction and entering the inner chamber. It then flows through the various vortex suppressor tubes 6 towards the disk center, performing disk cavity bleed air. Near the outer edge of the rotor disk 1, In the core area, vortex reduction is achieved in segments using anti-vortex nozzles 4 and anti-vortex tubes 6. The advantages of the anti-vortex tubes 6 and anti-vortex nozzles 4 are coupled in the inner and outer chambers, which can effectively reduce the generation of free vortices between the two-stage rotor disks 1. This can reliably reduce the pressure loss of the airflow, meet the requirements of cooling high-temperature components and sealing the chambers during the operation of the aero-engine, and allow for the use of shorter anti-vortex tubes 6, with a length not exceeding one-third of the rotor disk 1, thus avoiding an increase in the overall mass of the aero-engine and preventing vibration that could threaten the normal operation of the aero-engine.

[0052] For the aero-engine compressor cavity bleed air composite vortex reducer disclosed in the above embodiments, those skilled in the art will understand that its two-stage rotor disk 1, drum 2, outer edge mounting ring 3, disk center mounting ring 5, and vortex reducer tube 6 are assembled using annular stop edges, folded edges, mounting edges, etc., with matching bolts, making the overall structure simple and compact, easy to disassemble and assemble, and under the action of centrifugal force, the assembly between components is reliable, which can avoid the separation between components.

[0053] For the aero-engine compressor disk cavity air intake composite vortex reducer disclosed in the above embodiments, those skilled in the art will understand that its design of each anti-rotation nozzle 4 to each air intake hole, vortex reducer tube 6, and the design of the outer edge mounting ring 3 surrounding each anti-rotation nozzle 4 with local concavity can efficiently guide the airflow entering between the two-stage rotor disk 1 and reduce pressure loss.

[0054] In some optional embodiments, in the above-mentioned aero-engine compressor disk cavity bleed air composite vortex reducer, the outward annular mounting edge of the disk center has an outward flange, which presses on the outward annular folded edge D of the disk center of the subsequent stage. The mounting ring 5 of the disk center and the outward flange can lock the annular stop edge B of the disk center of the subsequent stage and the outward annular folded edge D of the disk center of the subsequent stage from the inside and outside sides, so that the overall structure is stable and reliable.

[0055] 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.

[0056] 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 composite vortex suppressor for compressor disk cavity of an aero-engine, characterized in that, include: Two-stage rotor disk (1), wherein the front stage rotor disk has an annular stop edge (A) at the center of the front stage disk on the rear side. The front side of the rear stage rotor disc has an annular stop edge (B) at the center of the rear stage disc and an annular stop edge (C) at the outer edge of the rear stage disc; the annular stop edge (B) at the center of the rear stage disc has an outward annular folded edge (D) at the center of the rear stage disc. The drum (2) has multiple circumferentially distributed air vents. Its rear end is formed on the front side of the rear stage rotor disk near the outer edge, and its front end points to the rear side of the front stage rotor disk. It has an inward annular fold (F) at the front end of the drum. An outer edge mounting ring (3) is set between two stage rotor discs (1). Its rear outer side is attached to the inner side of the annular stop edge (C) of the outer edge of the rear stage, and its front end has an outward annular mounting edge of the outer edge. The outward annular mounting edge of the outer edge extends into the space between the front stage rotor disc and the inward annular folded edge (F) of the front end of the drum, and is connected to the front stage rotor disc and the inward annular folded edge (F) of the front end of the drum by bolts. Multiple counter-rotating nozzles (4) are mounted circumferentially on the outer edge mounting ring (3); the outer edge mounting ring (3) has a partial recess around each counter-rotating nozzle (4); The disk center mounting ring (5) has multiple disk center mounting holes distributed circumferentially. It is set between the two-stage rotor disks (1). Its front outer side abuts against the inner side of the annular stop edge (A) of the front stage disk center, and its rear outer side abuts against the inner side of the annular stop edge (B) of the rear stage disk center. The rear outer side has an outward annular mounting edge of the disk center. The outward annular mounting edge of the disk center and the outward annular folded edge (D) of the rear stage disk center are connected by bolts. Multiple anti-vortex tubes (6) are installed in the mounting holes of each disk center, with the inlet end pointing to each anti-rotation nozzle (4) and the outer wall of the outlet end having an annular stop edge at the outlet end of the anti-vortex tube; the annular stop edge at the outlet end of each anti-vortex tube is attached to the inner side of the mounting ring (5) at the disk center. The length of the vortex reducer (6) does not exceed one-third of the rotor disk 1.

2. The aero-engine compressor disk cavity bleed air composite vortex suppressor according to claim 1, characterized in that, Each counter-rotating nozzle (4) is directly opposite each air intake hole.

3. The aero-engine compressor disk cavity bleed air composite vortex suppressor according to claim 1, characterized in that, The outer edge of the mounting ring (3) has a local concave shape or a ball-shaped depression surrounding each anti-rotation nozzle (4).

4. The aero-engine compressor disk cavity bleed air composite vortex suppressor according to claim 1, characterized in that, The outer edge mounting ring (3) is integrally formed with each anti-rotation nozzle (4).

5. The aero-engine compressor disk cavity bleed air composite vortex suppressor according to claim 1, characterized in that, Each vortex reducer (6) is a conduit of equal diameter.

6. The aero-engine compressor disk cavity bleed air composite vortex suppressor according to claim 1, characterized in that, The outer annular mounting edge, the front stage rotor disc, and the inner annular folded edge (F) at the front end of the drum are centered and positioned using a stop.

7. The aero-engine compressor disk cavity bleed air composite vortex suppressor according to claim 1, characterized in that, The outer annular mounting edge of the disc core has an outward flange, which presses against the outer annular fold (D) of the rear disc core.

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