Honeycomb bionic structure turbine disc of aero-engine

Abstract

The invention discloses a honeycomb bionic structure turbine disc of an aero-engine, and belongs to the technical field of aero-engine turbine disc design. The turbine disc comprises the disc edge, the radial plates and the hub, the radial plates are arranged to be of the structure of the left radial plate and the right radial plate, compared with a traditional single-radial-plate turbine disc, the weight can be reduced by 10%-15%, meanwhile, a plurality of honeycomb annular cavities which are communicated with one another are formed in the left radial plate, the right radial plate and the hub, the weight of the turbine disc is further reduced, and the service life of the turbine disc is prolonged. Compared with a traditional solid turbine disc, the weight of the structure is reduced by 20%-40%, and in addition, due to the fact that the specific strength of the honeycomb annular cavities is far higher than that of a traditional solid structure, the honeycomb annular cavities can bear larger loads under the same weight; according to the turbine disc, the cooling air inlet, the disc edge air collecting cavity and the cooling air outlet holes are formed, the disc center of the turbine disc communicates with the disc edge mortise area, cooling air can cool the disc edge, the radial plate, the hub and blades in the disc edge mortise area, and the cooling efficiency of the turbine disc is improved.

Description

Technical Field

[0001] The present invention belongs to the technical field of aero-engine turbine disk design, and particularly relates to an aero-engine honeycomb bionic structure turbine disk. Background Art

[0002] The turbine disk is a key component of an aero-engine. Most of them adopt a single web and solid structure, resulting in a large weight, which affects the thrust-to-weight ratio and efficiency improvement of the engine. During the operation of the turbine disk, the environmental temperature is high, the rotation speed is high, and the pressure is high. The harsh operating conditions of high temperature, high pressure, and high rotation speed require higher strength and better rigidity in the design of aero-engine turbine disks. Therefore, higher requirements are put forward for the specific strength and specific stiffness characteristics of the structure. At the same time, it is required that the turbine disk structure can achieve lightweight design to significantly reduce the weight of the engine.

[0003] In the prior art, in order to prevent bursting along the meridian plane, when designing the aero-engine turbine disk, the disk body of the aero-engine turbine disk is designed to be thick and heavy, and the disk body is thickened and weighted to increase the rupture speed of the disk. In addition, there are no other measures. Such improvements make the turbine disk too heavy, increasing the overall self-weight load of the aero-engine and reducing the overall performance of the aircraft. Summary of the Invention

[0004] Based on the problems existing in the above background art, the present invention aims to provide an aero-engine honeycomb bionic structure turbine disk, which solves the problems that the aero-engine turbine disk in the prior art is too heavy, increasing the overall self-weight load of the aero-engine and reducing the overall performance of the aircraft.

[0005] The embodiments of the present invention are implemented as follows:

[0006] The embodiment of the present invention provides an aero-engine honeycomb bionic structure turbine disk, which includes a rim, a web, and a hub; the hub is located inside the rim, the web is located between the rim and the hub, the outer side surface of the web is fixedly connected to the inner ring surface of the rim, and the inner side surface of the web is fixedly connected to the outer ring surface of the hub; a plurality of honeycomb annular cavities are arranged inside the web and the hub, and each honeycomb annular cavity is circumferentially arranged with the symmetry center line of the turbine disk as the axis; adjacent two honeycomb annular cavities are arranged at intervals and communicated through honeycomb inter-cooling holes; a plurality of cooling air inlets are arranged on the inner ring surface of the hub, and the plurality of cooling air inlets are communicated with a plurality of honeycomb annular cavities close to the inner ring surface of the hub.

[0007] A circumferential rim air collecting cavity is arranged on the top of the rim with the symmetry center line of the turbine disk as the axis, and the rim air collecting cavity is communicated with a plurality of honeycomb annular cavities located at the outer edge of the web; a rim dovetail groove area is circumferentially arranged on the outer ring surface of the rim with the symmetry center line of the turbine disk as the axis; the rim dovetail groove area and the rim air collecting cavity are communicated through cooling air outlet holes.

[0008] The basic principle of the honeycomb bionic structure turbine disk in the present invention is as follows: Cooling air enters the interior of the honeycomb annular cavity from the cooling air inlet located at the center of the turbine disk. It enters the hub and the interior of the web through the inter-honeycomb cooling holes in the honeycomb annular cavity to cool the disk body, and finally enters the disk rim of the turbine disk through the cooling air outlet holes. The cooling air enters the tenon groove area of the disk rim and further enters the interior of the turbine rotor blades to cool the blades. Multiple interconnected honeycomb annular cavities are arranged inside the web and the hub. This can not only reduce the weight of the turbine disk. Compared with the traditional solid turbine disk, the structural weight reduction reaches 20% - 40%. Moreover, since the specific strength of the honeycomb annular cavity is much higher than that of the traditional solid structure, it can bear a greater load under the same weight. The specific stiffness of the honeycomb annular cavity is also much higher than that of the traditional solid structure, and it has better anti-deformation ability under the same weight. The honeycomb annular cavity can absorb a large amount of energy through its own deformation when being impacted, and has good anti-impact performance. The performance of the honeycomb annular cavity can be flexibly adjusted by changing parameters such as the shape, size, and wall thickness of the honeycomb annular cavity to meet the requirements of different application scenarios, and it has strong designability.

[0009] In summary, multiple honeycomb annular cavities are arranged inside the web and the hub, enabling the entire turbine disk to better ensure the stiffness, damage tolerance characteristics, and lightweight design of the turbine disk on the premise of meeting the strength requirements, and solving the problem in the prior art that the turbine disk of the aeroengine is too heavy, increasing the overall self-weight load of the aeroengine and reducing the overall performance of the aircraft.

[0010] As an alternative solution to the above embodiment, the web includes a left web and a right web arranged symmetrically. The left web and the right web are inclined, there is an included angle between the left web and the right web, the outer sides of the left web and the right web are joined, and the inner sides of the left web and the right web are forked and separated; in the present application, by setting the web into the structure of the left web and the right web, compared with the traditional single-web turbine disk, a weight reduction of 10% - 15% can be achieved, further reducing the weight of the turbine disk and realizing lightweight design.

[0011] As an alternative solution to the above embodiment, the cross-sectional shape of each honeycomb annular cavity is a regular pentagon, a regular hexagon, a circle, or an ellipse.

[0012] In the alternative solution of the above embodiment, the cross-sectional shape of the honeycomb annular cavity can be preferably a regular hexagon, so that multiple honeycomb annular cavities form a honeycomb structure, which can utilize space more efficiently than other polygons, avoid waste, can evenly disperse external forces in all directions, avoid stress concentration, and while realizing the lightweight design of the entire turbine disk, the turbine disk also has high strength and stability.

[0013] As an alternative to the above embodiments, the space between two adjacent honeycomb annular cavities is provided with a honeycomb partition.

[0014] Furthermore, the thickness of the honeycomb partition is greater than or equal to 2 mm. The thickness of the honeycomb partition is adjusted according to the strength design requirements of the turbine disk. In areas with high stress, the thickness of the honeycomb partition is increased, the side length is reduced, and the distribution density of the plurality of honeycomb annular cavities is increased; in areas with low stress, the thickness dimension of the honeycomb partition is reduced, and the distribution density of the plurality of honeycomb annular cavities is decreased.

[0015] As an alternative to the above embodiments, a metal transition layer is provided between the honeycomb annular cavity and the surfaces of the left web, right web, and hub, and the thickness of the metal transition layer is greater than or equal to 3 mm. The above setting ensures the surface integrity and impact resistance of the turbine disk.

[0016] As an alternative to the above embodiments, a plurality of inter-honeycomb cooling holes are evenly arranged circumferentially in each honeycomb annular cavity, and the diameter of each inter-honeycomb cooling hole is greater than 1 mm, ensuring good manufacturability of the inter-honeycomb cooling holes.

[0017] As an alternative to the above embodiments, the disk rim, web, and hub are integrally formed by additive manufacturing.

[0018] The beneficial effects of the present invention are as follows:

[0019] 1. For a honeycomb bionic structure turbine disk in the present invention, by setting the web into the structure of a left web and a right web, compared with a traditional single-web turbine disk, a weight reduction of 10% - 15% can be achieved. At the same time, a plurality of interconnected honeycomb annular cavities are arranged inside the left web, right web, and hub, further reducing the weight of the turbine disk. Compared with a traditional solid turbine disk, the structural weight reduction reaches 20% - 40%. Moreover, since the specific strength of the honeycomb annular cavity is much higher than that of the traditional solid structure, a greater load can be borne under the same weight, realizing the lightweight design of the turbine disk and solving the problem in the prior art that the turbine disk of an aeroengine is too heavy, increasing the overall self-weight load of the aeroengine and reducing the overall performance of the aircraft.

[0020] 2. For a honeycomb bionic structure turbine disk in the present invention, by providing a cooling air inlet, a disk rim air collecting cavity, and a cooling air outlet hole, the disk center of the turbine disk is connected to the disk rim tenon groove area, enabling the cooling air to cool the disk rim, web, hub, and the blades in the disk rim tenon groove area, improving the cooling efficiency of the turbine disk, and reducing the working temperature of the turbine disk by 50°C - 100°C.

[0021] 3. A honeycomb bionic structure turbine disk in the present invention, the cross-section of the honeycomb annular cavity is preferably a regular hexagon. The honeycomb annular cavity with a regular hexagon structure can absorb a large amount of energy through its own deformation when being impacted, and has good impact resistance. This enables the entire turbine disk to effectively reduce its weight and improve the cooling efficiency while meeting the strength requirements, thereby enhancing the performance of the aero-engine. BRIEF DESCRIPTION OF THE DRAWINGS

[0022] In order to more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the drawings required to be used in the embodiments. Obviously, the drawings described below are only some embodiments of the present invention. For those of ordinary skill in the art, without creative efforts, other drawings can also be obtained based on these drawings. As shown by the drawings, the above-mentioned and other objects, features, and advantages of the present invention will become clearer. The same reference numerals indicate the same parts in all the drawings. The drawings are not deliberately drawn to scale in actual size, and the emphasis is on showing the gist of the present invention.

[0023] Figure 1 It is a schematic structural diagram of the meridian plane of the honeycomb bionic structure turbine disk of the aero-engine.

[0024] Figure 2 It is Figure 1 the schematic structural diagram in the A - A direction in

[0025] Figure 3 It is Figure 1 the enlarged structural diagram at B in

[0026] Figure 4 It is a three-dimensional sectional structural diagram of the honeycomb bionic structure turbine disk of the aero-engine.

[0027] Among them, 1. Disk rim; 2. Left spoke plate; 3. Right spoke plate; 4. Hub; 5. Cooling air inlet; 6. Disk rim air collecting cavity; 7. Turbine disk symmetry center line; 8. Cooling air outlet hole; 9. Disk rim tenon groove area; 10. Honeycomb annular cavity; 11. Inter-honeycomb cooling holes; 12. Honeycomb partition. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0028] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the following will clearly and completely describe the technical solutions in the embodiments of the present invention with reference to the drawings in the embodiments of the present invention. Obviously, the described embodiments are some, but not all, of the embodiments of the present invention. Usually, the components of the embodiments of the present invention described and shown in the drawings here can be arranged and designed in various different configurations.

[0029] Accordingly, the following detailed description of the embodiments of the present invention provided in the accompanying drawings is not intended to limit the scope of the claimed invention, but merely represents selected embodiments of the present invention. All other embodiments obtained by those of ordinary skill in the art based on the embodiments of the present invention without creative efforts shall fall within the scope of protection of the present invention.

[0030] It should be noted that like reference numerals and letters denote like items in the following drawings. Therefore, once an item is defined in one drawing, it does not require further definition and explanation in subsequent drawings.

[0031] As Figures 1 to 4 shown, the present invention provides a honeycomb bionic structure turbine disk for an aeroengine, which comprises a rim 1, a web and a hub 4; the hub 4 is located inside the rim 1, the web is located between the rim 1 and the hub 4, the outer side surface of the web is fixedly connected to the inner ring surface of the rim 1, and the inner side surface of the web is fixedly connected to the outer ring surface of the hub 4; a plurality of honeycomb annular cavities 10 are arranged inside the web and the hub 4, and each of the honeycomb annular cavities 10 is circumferentially arranged with the symmetry center line 7 of the turbine disk as the axis; adjacent two honeycomb annular cavities 10 are arranged at intervals and communicated through honeycomb inter-cooling holes 11. A plurality of cooling air inlets 5 are arranged on the inner ring surface of the hub 4, and the plurality of cooling air inlets 5 are communicated with the plurality of honeycomb annular cavities 10 close to the inner ring surface of the hub 4.

[0032] A circumferential rim air collecting cavity 6 is arranged on the top of the rim 1 with the symmetry center line 7 of the turbine disk as the axis, and the rim air collecting cavity 6 is communicated with the plurality of honeycomb annular cavities 10 located at the outer edge of the web; a rim dovetail groove area 9 is circumferentially arranged on the outer ring surface of the rim 1 with the symmetry center line 7 of the turbine disk as the axis; a cooling air outlet hole 8 is arranged between the rim dovetail groove area 9 and the rim air collecting cavity 6.

[0033] The basic principle of the honeycomb bionic structure turbine disk for aero-engines is as follows: Cooling air enters the interior of the honeycomb annular cavity 10 from the cooling air inlet 5 located at the center of the turbine disk. It passes through the inter-honeycomb cooling holes 11 of the honeycomb annular cavity 10 and enters the hub 4 and the internal cooling disk body of the web plate. Finally, it enters the disk rim 1 of the turbine disk through the cooling air outlet holes 8. The cooling air enters the tenon groove area 9 of the disk rim and further enters the interior of the turbine rotor blades to cool the blades, improving the cooling efficiency of the turbine disk. This can reduce the operating temperature of the turbine disk by 50°C to 100°C. Multiple interconnected honeycomb annular cavities 10 are provided inside the web plate and the hub 4. This can not only reduce the weight of the turbine disk. Compared with traditional solid turbine disks, the structural weight reduction reaches 20% - 40%. Moreover, since the specific strength of the honeycomb annular cavity 10 is much higher than that of traditional solid structures, it can withstand greater loads under the same weight. The specific stiffness of the honeycomb annular cavity 10 is also much higher than that of traditional solid structures, and it has better anti-deformation ability under the same weight. When the honeycomb annular cavity 10 is impacted, it can absorb a large amount of energy through its own deformation and has good anti-impact performance. The performance of the honeycomb annular cavity 10 can be flexibly adjusted by changing parameters such as the shape, size, and wall thickness of the honeycomb annular cavity 10 to meet the requirements of different application scenarios, and it has strong designability.

[0034] In summary, multiple honeycomb annular cavities 10 are provided inside the web plate and the hub 4, enabling the entire turbine disk to better ensure the stiffness, anti-damage tolerance characteristics, and lightweight design of the turbine disk on the premise of meeting the strength requirements, and solving the problem in the prior art that the aero-engine turbine disk is too heavy, increasing the overall self-weight load of the aero-engine and reducing the overall performance of the aircraft.

[0035] Preferably but not limited to, the disk rim 1, the web plate, and the hub 4 are integrally formed by additive manufacturing. Due to the setting of multiple honeycomb annular cavities 10, the structure and manufacturing process of the turbine disk will become more complex. Additive manufacturing technology constructs a three-dimensional structure by layer-by-layer material stacking and can manufacture complex structural parts that are difficult to achieve by traditional processing methods, such as flow channels and thin walls, and is particularly suitable for the manufacturing of aero-engine turbine disks.

[0036] Preferably, the web plate includes a left web plate 2 and a right web plate 3 that are symmetrically arranged. The left web plate 2 and the right web plate 3 are inclined, and there is an included angle between the left web plate 2 and the right web plate 3. The outer sides of the left web plate 2 and the right web plate 3 are connected, and the inner sides of the left web plate 2 and the right web plate 3 are forked and separated; in this application, by setting the web plate into the structure of the left web plate 2 and the right web plate 3, compared with traditional single-web turbine disks, a weight reduction of 10% - 15% can be achieved, further reducing the weight of the turbine disk and realizing lightweight design.

[0037] Specifically, the cross-sectional shape of each honeycomb annular cavity 10 is a regular pentagon, a regular hexagon, a circle, or an ellipse.

[0038] In an alternative of the above embodiment, the cross-sectional shape of the honeycomb annular cavity 10 can be preferably a regular hexagon, so that a plurality of honeycomb annular cavities 10 form a honeycomb structure. The regular hexagon can make more efficient use of space than other polygons, avoid waste, and can evenly disperse external forces in all directions to avoid stress concentration. While achieving the lightweight design of the entire turbine disk, the turbine disk also has high strength and stability.

[0039] The two adjacent honeycomb annular cavities 10 are spaced by a honeycomb partition 12. The thickness of the honeycomb partition 12 is greater than or equal to 2 mm, and the thickness of the honeycomb partition 12 is adjusted according to the strength design requirements of the turbine disk. In areas with high stress, the thickness of the honeycomb partition 12 is increased, the side length is reduced, and the distribution density of the plurality of honeycomb annular cavities 10 is increased; in areas with low stress, the thickness dimension of the honeycomb partition 12 is reduced, and the distribution density of the plurality of honeycomb annular cavities 10 is reduced.

[0040] As an alternative of the above embodiment, a metal transition layer is provided between the honeycomb annular cavity 10 and the surfaces of the left spoke plate 2, the right spoke plate 3, and the hub 4, and the thickness of the metal transition layer is greater than or equal to 3 mm. The above setting ensures the surface integrity and impact resistance of the turbine disk.

[0041] Specifically, a plurality of the honeycomb inter-cooling holes 11 are evenly arranged circumferentially in each honeycomb annular cavity 10, and the diameter of each honeycomb inter-cooling hole 11 is greater than 1 mm to ensure good manufacturability of the honeycomb inter-cooling holes 11.

[0042] In summary, compared with the traditional solid turbine disk, the weight of the honeycomb bionic structure turbine disk in the present invention is reduced by 20% - 40%, and it can bear a greater load under the same weight; a cooling air flow channel is provided in the turbine disk, and the cooling air flow channel refers to the cooling air inlet 5, the rim air collecting cavity 6, the cooling air outlet hole 8, the rim dovetail groove area 9, the honeycomb annular cavity 10, and the honeycomb inter-cooling holes 11, which can direct the cooling air to the blades in the rim dovetail groove area 9, so that the turbine disk has a high cooling efficiency and can reduce the working temperature of the turbine disk by 50°C - 100°C.

[0043] The above is only the preferred embodiment of the present invention and is not used to limit the present invention. For those skilled in the art, the present invention can have various changes and modifications. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention shall be included in the protection scope of the present invention.

Claims

1. A honeycomb bionic structure turbine disk for an aeroengine, characterized in that It includes a disk rim, a web and a hub; The hub is located inside the disk rim, the web is located between the disk rim and the hub, the outer side surface of the web is fixedly connected to the inner ring surface of the disk rim, and the inner side surface of the web is fixedly connected to the outer ring surface of the hub; Multiple honeycomb annular cavities are provided inside the web and the hub, and each of the honeycomb annular cavities is circumferentially arranged with the symmetry center line of the turbine disk as the axis; Adjacent two honeycomb annular cavities are arranged at intervals and communicated through honeycomb inter-cooling holes; Multiple cooling air inlets are provided on the inner ring surface of the hub, and the multiple cooling air inlets are communicated with multiple honeycomb annular cavities close to the inner ring surface of the hub; A ring of disk rim gas collecting cavity is circumferentially arranged at the top of the disk rim with the symmetry center line of the turbine disk as the axis, and the disk rim gas collecting cavity is communicated with multiple honeycomb annular cavities located at the outer edge of the web; A disk rim dovetail groove area is circumferentially arranged on the outer ring surface of the disk rim with the symmetry center line of the turbine disk as the axis; The disk rim dovetail groove area and the disk rim gas collecting cavity are communicated through cooling air outlet holes.

2. The honeycomb bionic structure turbine disk of the aeroengine according to claim 1, characterized in that, The web includes a left web and a right web which are symmetrically arranged, the left web and the right web are inclined, there is an included angle between the left web and the right web, the outer side surfaces of the left web and the right web are joined, and the inner side surfaces of the left web and the right web are forked and separated.

3. The honeycomb bionic structure turbine disk of the aeroengine according to claim 1, characterized in that, The cross-sectional shape of each honeycomb annular cavity is a regular pentagon, a regular hexagon, a circle or an ellipse.

4. The honeycomb bionic structure turbine disk of the aeroengine according to claim 1, characterized in that Adjacent two honeycomb annular cavities are separated by a honeycomb partition.

5. The honeycomb bionic structure turbine disk of the aeroengine according to claim 4, characterized in that, The thickness of the honeycomb partition is greater than or equal to 2 mm, and the thickness of the honeycomb partition is adjusted according to the strength design requirements of the turbine disk. In the area with large stress, the thickness of the honeycomb partition is increased, the side length is reduced, and the distribution density of the multiple honeycomb annular cavities is increased; In the area with small stress, the thickness dimension of the honeycomb partition is reduced, and the distribution density of the multiple honeycomb annular cavities is reduced.

6. The honeycomb bionic structure turbine disk of the aeroengine according to claim 2, characterized in that A metal transition layer is provided between the honeycomb annular cavity and the surfaces of the left web, the right web and the hub, and the thickness of the metal transition layer is greater than or equal to 3 mm.

7. The honeycomb bionic structure turbine disk of the aeroengine according to claim 1, characterized in that, Each honeycomb annular cavity is circumferentially and uniformly provided with multiple honeycomb inter-cooling holes, and the diameter of each honeycomb inter-cooling hole is greater than 1 mm.

8. The honeycomb bionic structure turbine disk of an aeroengine according to any one of claims 1 to 7, characterized in that The disk rim, the web and the hub are integrally formed by additive manufacturing.

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