A thin-walled annular disc cavity air bleed structure for an aero-engine compressor

Abstract

This application specifically relates to a thin-walled annular disk cavity bleed air structure for an aero-engine compressor, including a front-stage rotor disk, a rear-stage rotor disk, a thin-walled bleed air ring, and a corrugated damping ring. The front-stage rotor disk and the rear-stage rotor disk are arranged opposite to each other and connected by a drum. The drum has multiple circumferentially distributed outer edge bleed air holes. The front side of the rear-stage rotor disk has a support ring located inside the drum, which has multiple circumferentially distributed disk center bleed air holes. The thin-walled bleed air ring is disposed between the front-stage rotor disk and the rear-stage rotor disk, located inside the drum. Its inner edge is connected to the support ring, and its outer edge has a forward connecting ring connected to the front-stage rotor disk. A bleed cavity is formed between the thin-walled bleed air ring and the rear-stage rotor disk, and the bleed cavity connects to each outer edge bleed air hole and the disk center bleed air hole. The rear side of the thin-walled bleed air ring has multiple circumferentially distributed radial guide fins. The corrugated damping ring is disposed between the drum and the forward connecting ring.

Description

Technical Field

[0001] This application belongs to the field of air bleed design technology for aero-engine compressor disc cavity, specifically relating to a thin-walled annular disc cavity air bleed structure for aero-engine compressor. Background Technology

[0002] Increasing the turbine inlet temperature of an aero-engine can effectively improve its efficiency and thrust-to-weight ratio, but it also leads to a deterioration of the working environment of the hot-end components of the aero-engine, requiring further improvement in the cooling effect of the hot-end components.

[0003] Currently, most aero-engine turbines use film cooling. The required cooling air is usually drawn from a suitable location in the compressor, led out from the compressor disk cavity, and guided from the disk center to the turbine section to cool the turbine disk and turbine blades.

[0004] Aircraft engine compressors operate at high speeds. Air is drawn into the compressor disk cavity, and this air enters a typical rotating disk cavity, where it forms strong vortices during radial inward flow. Figure 1 As shown, if the gas forms a free vortex, the swirling flow will become increasingly intense during the radial inward flow, resulting in a very large pressure loss.

[0005] Designing a vortex suppressor in the compressor disc bleed air system can effectively reduce the pressure loss of the bleed air. Currently, common vortex suppressors include tubular vortex suppressors, anti-swirl nozzle vortex suppressors, and guide vane vortex suppressors, such as... Figure 2 As shown, radial guidance promotes radial flow of air, disrupts free vortices, reduces pressure loss, and makes it easier for air to flow from the compressor hub to the disk center.

[0006] Tubular vortex suppressors lack vibration damping measures in their bleed pipes. As rotor components, they rotate with the compressor rotor, generating significant vibrations that can damage themselves and result in low reliability. Baffle-type vortex suppressors rely on the location of the bleed port and the span between the two disks to function, and their large size and weight of the bleed disk and guide vanes hinder weight reduction. Furthermore, current compressor disk bleed structures have large radii and rotation radii, generating significant stress during operation, and are structurally complex with poor assemblability, requiring some parts to be assembled blindly.

[0007] In view of the aforementioned technical deficiencies, this application is hereby filed. Summary of the Invention

[0008] The purpose of this application is to provide a thin-walled annular disc cavity air bleed structure for an aero-engine compressor, in order to overcome or mitigate at least one of the known technical defects.

[0009] The technical solution of this application is:

[0010] A thin-walled annular disc cavity air bleed structure for an aero-engine compressor, characterized in that it includes a front-stage rotor disc, a rear-stage rotor disc, a thin-walled air bleed ring, and a corrugated damping ring;

[0011] The front-stage rotor disk and the rear-stage rotor disk are arranged opposite to each other and connected by a drum. The drum has multiple circumferentially distributed outer edge air vents.

[0012] The front side of the rear rotor disc has a support ring, which is located inside the drum and has multiple circumferentially distributed disc center air vents.

[0013] The thin-walled bleed air ring is set between the front-stage rotor disc and the rear-stage rotor disc, located inside the drum. Its inner edge is connected to the support ring, and its outer edge has a forward connecting ring, which is connected to the front-stage rotor disc.

[0014] A flow chamber is formed between the thin-walled air duct ring and the rear rotor disk. The flow chamber is connected to the outer edge air duct and the disk center air duct. The rear side of the thin-walled air duct ring has multiple radially distributed guide fins.

[0015] The corrugated damping ring is positioned between the drum and the forward connecting ring.

[0016] According to at least one embodiment of this application, in the above-described thin-walled annular disc cavity air bleed structure of the aero-engine compressor, the rear end of the drum is formed on the rear stage rotor disk, and the front end has a first inward folded edge, which is connected to the front stage rotor disk by bolt fasteners.

[0017] The air vents on the outer edge of the drum are distributed near the rear rotor disc;

[0018] The space between the drum and the inner ring of the stator connected to its outer side is sealed with honeycomb teeth.

[0019] According to at least one embodiment of this application, in the above-described thin-walled annular disc cavity air bleed structure of the aero-engine compressor, the support ring is formed on the front side of the rear stage rotor disk.

[0020] According to at least one embodiment of this application, in the above-described thin-walled annular disc cavity bleed air structure of the aero-engine compressor, the thin-walled bleed air ring is disposed close to the rear stage rotor disk.

[0021] The front end of the support ring has a second inward folded edge, and the inner edge of the thin-walled air duct ring is connected to the second inward folded edge by bolt fasteners, and the inner edge is raised forward.

[0022] The front end of the forward connecting ring has a third inward folded edge, which is connected to the front rotor disc by bolt fasteners;

[0023] The thin-walled air intake ring is composed of multiple sector-shaped segments spliced ​​together circumferentially.

[0024] Each radial guide fin is welded to the rear side of the thin-walled air intake ring.

[0025] According to at least one embodiment of this application, in the above-described thin-walled annular disc cavity air bleed structure of the aero-engine compressor, the front end of the corrugated damping ring has a fourth inward folded edge, and the fourth inward folded edge is connected to the front stage rotor disk by bolt fasteners.

[0026] The first inward fold, the third inward fold, and the fourth inward fold can share bolt fasteners for connection on the front rotor disc;

[0027] The corrugated damping ring is made up of multiple arc-shaped segments spliced ​​together circumferentially.

[0028] According to at least one embodiment of this application, the above-described thin-walled annular disc cavity air bleed structure for an aero-engine compressor further includes an air bleed pipe.

[0029] The air intake pipe is installed inside the disk center of the rear rotor, forming an air intake channel with the disk center of the rear rotor. The disk center of the front rotor has a rearward connecting ring, which is inclined inward and is sleeved on the air intake pipe through a rearward annular segment, making sealed contact with the air intake pipe, and a sealing ring is provided therebetween. Attached Figure Description

[0030] Figure 1 This is a schematic diagram of bleed air from the compressor disk cavity of a current aero-engine;

[0031] Figure 2 This is a schematic diagram of a common type of vortex reducer;

[0032] Figure 3 This is a schematic diagram of the thin-walled annular disc cavity air bleed structure of the aero-engine compressor provided in the embodiments of this application;

[0033] Figure 4 This is a schematic diagram of the front-stage rotor disk provided in an embodiment of this application;

[0034] Figure 5 yes Figure 4 View from direction A;

[0035] Figure 6 This is a schematic diagram of the downstream rotor disk provided in an embodiment of this application;

[0036] Figure 7 yes Figure 6 View from direction B;

[0037] Figure 8 This is a schematic diagram of the thin-walled air-guiding ring provided in an embodiment of this application;

[0038] Figure 9 yes Figure 8 The C-direction view;

[0039] Figure 10 This is a schematic diagram of the corrugated damping ring provided in an embodiment of this application;

[0040] Figure 11 yes Figure 8 The D-direction view;

[0041] in:

[0042] 1-Pre-stage rotor disc; 2-Rear-stage rotor disc; 3-Thin-walled bleed air ring; 4-Corrugated damping ring; 5-Drum; 6-Stator inner ring; 7-Support ring; 8-Radial guide fins; 9-Bleed air pipe.

[0043] To better illustrate this embodiment, some content in the accompanying drawings may be omitted, enlarged, or reduced. They are for illustrative purposes only and should not be construed as limiting the scope of this application. Detailed Implementation

[0044] 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, and other related parts can be referred to the general design.

[0045] 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 word "comprising" as used in this application description indicates that the concept preceding the word encompasses the concepts listed following the word and their equivalents, without excluding other related concepts.

[0046] Furthermore, the terms indicating location used in the description of this application are only used to indicate relative directions or positional relationships. When the absolute position of the described object changes, its relative positional relationship may also change accordingly. It should also be noted that, unless otherwise explicitly specified and limited, terms such as "installation" and "connection" used in the description of this application should be interpreted broadly. For example, a connection can be a fixed connection or a detachable 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. Those skilled in the art can understand its specific meaning in this application according to the specific circumstances.

[0047] To address the challenges of high compressor speeds, complex operating conditions, and limited disk space in aero-engine compressors, where the bleed air mechanism must rotate with the rotor, resulting in severe vibration issues, this application provides a thin-walled annular disk-type bleed air structure for aero-engine compressors. Figure 3As shown, it includes a front-stage rotor disc 1, a rear-stage rotor disc 2, a thin-walled bleed air ring 3, a corrugated damping ring 4, and a bleed air pipe 9.

[0048] Fore-stage rotor disk 1 Figure 4-5 As shown, the rear rotor disk 2 is as follows Figure 6-7 As shown.

[0049] The front-stage rotor disk 1 and the rear-stage rotor disk 2 are arranged opposite to each other and connected by a drum 5. The drum 5 has multiple circumferentially distributed outer edge air vents.

[0050] The rear end of the drum 5 is formed on the rear rotor disk 2, and the front end has a first inward folded edge, which is connected to the front rotor disk 1 by bolts. The air vents on the outer edge of the drum 5 are distributed close to the rear rotor disk 2. The drum 5 and the stator inner ring 6 that is sleeved on its outer side are sealed with honeycomb teeth.

[0051] The front side of the rear rotor disk 2 has a support ring 7, which is located inside the drum 5 and has multiple circumferentially distributed disc center air vents.

[0052] The support ring 7 is formed on the front side of the rear rotor disk 2.

[0053] Thin-walled air duct ring 3 Figure 8-9 As shown, it is disposed between the front rotor disk 1 and the rear rotor disk 2, located inside the drum 5, with its inner edge connected to the support ring 7 and its outer edge having a forward connecting ring, which is connected to the front rotor disk 1.

[0054] A flow chamber is formed between the thin-walled air duct ring 3 and the rear rotor disk 2. The flow chamber is connected to the outer edge air duct and the disk center air duct. The rear side of the thin-walled air duct ring 3 has multiple radially distributed guide fins 8.

[0055] The thin-walled bleed air ring 3 is positioned close to the rear rotor disk 2. The support ring 7 has a second inward folded edge at its front end. The inner edge of the thin-walled bleed air ring 3 is connected to the second inward folded edge by bolts, and the inner edge curves forward. The forward connecting ring has a third inward folded edge at its front end, which is connected to the front rotor disk 1 by bolts. The thin-walled bleed air ring 3 is composed of multiple fan-shaped segments spliced ​​circumferentially. Each radial guide fin 8 is welded to the rear side of the thin-walled bleed air ring 3.

[0056] Corrugated damping ring 4 Figure 10-11 As shown, it is positioned between the drum 5 and the forward connecting ring.

[0057] The corrugated damping ring 4 has a fourth inward fold at its front end, which is connected to the front rotor disk 1 by bolts. The first, third, and fourth inward folds can share the same bolts for connection on the front rotor disk 1. The corrugated damping ring 4 is composed of multiple arc-shaped segments spliced ​​circumferentially, corresponding to the various sector-shaped segments of the thin-walled bleed air ring 3.

[0058] The air intake pipe 9 is installed inside the center of the rear rotor disk 2, forming an air intake channel between it and the center of the rear rotor disk 2. The center of the front rotor disk 1 has a rearward connecting ring, which is inclined inward and is sleeved on the air intake pipe 9 through a rearward annular segment, making sealed contact with the air intake pipe 9, and a sealing ring is provided therebetween.

[0059] The thin-walled annular disc cavity air bleed structure of the aero-engine compressor disclosed in the above embodiments allows high-pressure gas from the compressor flow channel to enter the air bleed cavity through the outer edge air bleed hole, flow radially within the air bleed cavity, and then exit through the disc center air bleed hole and flow out through the air bleed channel to supply cooling for hot-end components. During the radial flow of high-pressure gas within the air bleed cavity, the radial guide fins 8 can guide the airflow, disrupt airflow vortices, and reduce pressure loss.

[0060] The thin-walled annular disc cavity air bleed structure of the aero-engine compressor disclosed in the above embodiments is fixed to the front-stage rotor disk 1 and the rear-stage rotor disk 2 by bolts and fasteners. Radial guide fins 8 are welded on the thin-walled air bleed ring 3, and the forward connecting ring on the thin-walled air bleed ring 3 is tightly attached to the inner wall of the drum 5. It is easy to assemble and disassemble, and can be assembled by referring to the following process:

[0061] The thin-walled air bleed ring 3 and the corrugated damping ring 4 are installed into the drum 5. They can be inserted through the lower front end of the support ring 7. The inner edge of the thin-walled air bleed ring 3 is connected to the second inward folded edge of the support ring 7 by bolt fasteners.

[0062] The first inward folded edge at the front end of the drum 5, the third inward folded edge at the front end of the forward connecting ring on the thin-walled air bleed ring 3, and the fourth inward folded edge at the front end of the corrugated damping ring 4 are connected to the front rotor disk 1 by bolts and fasteners.

[0063] Push the air intake pipe 9 into the center of the rear rotor disk 2, and make sealing contact with the rear annular section of the rear connecting ring on the front rotor disk 1 to complete the assembly.

[0064] In the thin-walled annular disc cavity air bleed structure of the aero-engine compressor disclosed in the above embodiments, the thin-walled air bleed ring 3 is fixed at both ends, and the corrugated damping ring 4 can absorb the vibration energy of the structure during operation, thereby reducing vibration and improving the reliability and life of the structure.

[0065] In the thin-walled annular disc cavity air bleed structure of the aero-engine compressor disclosed in the above embodiments, the use of the thin-walled air bleed ring 3 can greatly reduce the structural weight and volume, shorten the size of the air bleed cavity, significantly reduce the size of the radial guide fins 8, and allow weight-reducing grooves to be opened on the support structure, which can further reduce the structural weight.

[0066] The closer the bleed tube of a tubular vortex reducer is to the bleed hole, the higher the bleed efficiency. However, due to the excessive length of the bleed tube, vibration problems prevent it from reaching the vicinity of the bleed hole. In contrast, the more guide vanes a vortex reducer has, the higher the bleed efficiency. However, due to the large size of the guide vanes and limited assembly space, too many cannot be arranged. In the thin-walled annular disc cavity bleed structure of the aero-engine compressor disclosed in the above embodiments, the thin-walled structure combined with fins makes it easier to approach the bleed hole and allows for a reasonable increase in the number of fins. The size of the fins can be adjusted according to the location of the bleed hole, and the size of the bleed cavity can be reduced by moving closer to the bleed hole, which can effectively improve bleed efficiency, reduce pressure loss, and reduce structural stress.

[0067] Current vortex reducer structures require the rotor to be connected first before assembly. Some bolts and fasteners require special tooling and cannot be operated visually, making assembly difficult. However, the thin-walled ring-type disc cavity air bleed structure of the aero-engine compressor disclosed in the above embodiment can first fix one end of the thin-walled air bleed ring 3 to the rotor disc, and then fix the other end by connecting bolts and fasteners after positioning through the disc stop, thus completing the assembly. The assembly process is simple and clear, the assemblability is greatly improved, and it is not limited by the rotor structure space.

[0068] 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 thin-walled annular disc cavity air bleed structure for an aero-engine compressor, characterized in that, It includes a front-stage rotor disc (1), a rear-stage rotor disc (2), a thin-walled air bleed ring (3), and a corrugated damping ring (4). The front rotor disk (1) and the rear rotor disk (2) are arranged opposite to each other and connected by a drum (5). The drum (5) has multiple circumferentially distributed outer edge air vents. The front side of the rear rotor disk (2) has a support ring (7), which is located inside the drum (5) and has multiple circumferentially distributed disc center air vents. The thin-walled air intake ring (3) is set between the front rotor disk (1) and the rear rotor disk (2), located inside the drum (5), with its inner edge connected to the support ring (7) and its outer edge having a forward connecting ring connected to the front rotor disk (1). A flow chamber is formed between the thin-walled air duct ring (3) and the rear rotor disk (2). The flow chamber is connected to the outer edge air duct and the disk center air duct. The rear side of the thin-walled air duct ring (3) has multiple radially distributed guide fins (8). The corrugated damping ring (4) is placed between the drum (5) and the forward connecting ring.

2. The thin-walled annular disc cavity air bleed structure for an aero-engine compressor according to claim 1, characterized in that, The rear end of the drum (5) is formed on the rear rotor disk (2), and the front end has a first inward folded edge, which is connected to the front rotor disk (1) by bolt fasteners. The air vents on the outer edge of the drum (5) are distributed near the rear rotor disk (2); The drum (5) and the inner ring (6) of the stator connected to its outer side are sealed with honeycomb teeth.

3. The thin-walled annular disc cavity air bleed structure for an aero-engine compressor according to claim 2, characterized in that, The support ring (7) is formed on the front side of the rear rotor disk (2).

4. The thin-walled annular disc cavity air bleed structure for an aero-engine compressor according to claim 3, characterized in that, The thin-walled bleed air ring (3) is positioned close to the rear rotor disc (2); The front end of the support ring (7) has a second inward folded edge, and the inner edge of the thin-walled air duct ring (3) is connected to the second inward folded edge by bolt fasteners, and the inner edge is raised forward; The front end of the forward connecting ring has a third inward folded edge, which is connected to the front rotor disc (1) by bolt fasteners; The thin-walled air-drawing ring (3) is composed of multiple fan-shaped segments spliced ​​together circumferentially; Each radial guide fin (8) is welded to the rear side of the thin-walled air intake ring (3).

5. The thin-walled annular disc cavity air bleed structure for an aero-engine compressor according to claim 4, characterized in that, The front end of the corrugated damping ring (4) has a fourth inward folded edge, which is connected to the front rotor disk (1) by bolt fasteners; The first inward fold, the third inward fold, and the fourth inward fold can be connected by bolt fasteners on the front rotor disc (1); The corrugated damping ring (4) is made up of multiple arc segments spliced ​​together circumferentially.

6. The thin-walled annular disc cavity air bleed structure for an aero-engine compressor according to claim 5, characterized in that, It also includes the airway (9); The air intake pipe (9) is installed inside the disk center of the rear rotor disk (2) to form an air intake channel with the disk center of the rear rotor disk (2). The disk center of the front rotor disk (1) has a rearward connecting ring. The rearward connecting ring is inclined inward and is sleeved on the air intake pipe (9) through a rearward annular segment, and is in sealed contact with the air intake pipe (9), and a sealing ring is provided therebetween.

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