A honeycomb bionic structure turbine disk for an aircraft engine

Abstract

The present invention discloses a turbine disk with a honeycomb biomimetic structure for an aero-engine, belonging to the technical field of aero-engine turbine disk design. The turbine disk comprises a rim, spokes, and a hub. By arranging the spokes into a left and right structure, the disk can achieve a weight reduction of 10% to 15% compared to a conventional single-spoke turbine disk. Multiple interconnected honeycomb annular cavities are provided within the left and right spokes and the hub, further reducing the weight of the turbine disk. Compared to a conventional solid turbine disk, the weight reduction is 20% to 40%. Furthermore, because the specific strength of the honeycomb annular cavities is much higher than that of a conventional solid structure, the disk can withstand greater loads at the same weight. A cooling air inlet, a rim air collecting cavity, and a cooling air outlet are provided to connect the center of the turbine disk with the tongue-and-groove region of the rim, allowing cooling air to cool the rim, spokes, hub, and blades in the tongue-and-groove region of the rim, thereby improving the cooling efficiency of the turbine disk.

Description

Technical Field

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

[0002] Turbine discs are key components in aircraft engines. They typically utilize a single-spoke, solid structure, resulting in heavy weight that impacts the engine's thrust-to-weight ratio and efficiency. Turbine discs operate under high temperatures, high speeds, and high pressures. These demanding operating conditions demand higher strength and rigidity in aircraft engine turbine disc design, placing higher demands on the structural specific strength and stiffness. Furthermore, the disc structure must be lightweight to significantly reduce engine weight.

[0003] In order to prevent the aircraft engine turbine disk from bursting along the meridian plane, the aircraft engine turbine disk in the existing technology is designed to have a thick disk body during design. The thickening and weighting of the disk body is used to increase the disk bursting speed. Apart from this, there are no other measures. This type of improvement makes the turbine disk too heavy, increases the overall deadweight load of the aircraft engine, and reduces the overall performance of the aircraft. Summary of the Invention

[0004] Based on the problems existing in the above background technology, the present invention aims to provide a honeycomb bionic structure turbine disk for an aircraft engine, which solves the problem in the prior art that the turbine disk for an aircraft engine is too heavy, increases the overall deadweight load of the aircraft engine, and reduces the overall performance of the aircraft.

[0005] The embodiment of the present invention is achieved as follows:

[0006] An embodiment of the present invention provides a honeycomb bionic structure turbine disk of an aero-engine, which includes a disk rim, a spoke plate and a hub; the hub is located inside the disk rim, the spoke plate is located between the disk rim and the hub, the outer side surface of the spoke plate is fixedly connected to the inner annular surface of the disk rim, and the inner side surface of the spoke plate is fixedly connected to the outer annular surface of the hub; a plurality of honeycomb annular cavities are arranged inside the spoke plate and the hub, and each of the honeycomb annular cavities is arranged in a circumferential direction with the symmetric center line of the turbine disk as the axis; adjacent honeycomb annular cavities are arranged at intervals and connected through inter-honeycomb cooling holes; a plurality of cooling air inlets are provided on the inner annular surface of the hub, and the plurality of cooling air inlets are connected to a plurality of honeycomb annular cavities close to the inner annular surface of the hub.

[0007] A disk rim air collecting cavity is provided in a circle on the top of the disk rim with the symmetric center line of the turbine disk as the axis, and the disk rim air collecting cavity is connected with multiple honeycomb annular cavities located at the outer edge of the spoke plate; a disk rim tongue and groove area is provided on the outer ring surface of the disk rim with the symmetric center line of the turbine disk as the axis; the disk rim tongue and groove area is connected to the disk rim air collecting cavity through a cooling air outlet hole.

[0008] The basic principle of the honeycomb biomimetic turbine disk for aircraft engines described in this invention is that cooling air enters the honeycomb annular cavity from the cooling air inlet at the center of the turbine disk. It then passes through the intercellular cooling holes within the honeycomb annular cavity and enters the cooling disk body within the hub and spokes. Finally, it enters the turbine disk rim through the cooling air outlet. The cooling air enters the tongue and groove area of the rim and further enters the turbine rotor blades to cool them. Multiple interconnected honeycomb annular cavities are arranged within the spokes and hub. This not only reduces the weight of the turbine disk, achieving a weight reduction of 20% to 40% compared to traditional solid turbine disks, but also, because the honeycomb annular cavity has a much higher specific strength than traditional solid structures, it can withstand greater loads for the same weight. Its specific stiffness is also much higher than that of traditional solid structures, providing better deformation resistance for the same weight. When impacted, the honeycomb annular cavity absorbs significant energy through its own deformation, resulting in excellent impact resistance. The performance of the honeycomb annular cavity can be flexibly adjusted by varying parameters such as its shape, size, and wall thickness to meet the needs of different application scenarios, providing high design flexibility.

[0009] To sum up, multiple honeycomb annular cavities are provided inside the spokes and the hub, so that the entire turbine disc can better ensure the stiffness, damage tolerance and lightweight design of the turbine disc while meeting the strength requirements, thereby solving the problem in the prior art that the turbine disc of the aircraft engine is too heavy, which increases the overall deadweight load of the aircraft engine and reduces the overall performance of the aircraft.

[0010] As an optional solution to the above embodiment, the spoke plate includes a symmetrically arranged left spoke plate and a right spoke plate, the left spoke plate and the right spoke plate are arranged at an angle, an angle is set between the left spoke plate and the right spoke plate, the outer side surfaces of the left spoke plate and the right spoke plate are connected, and the inner side surfaces of the left spoke plate and the right spoke plate are forked and separated; in the present application, by arranging the spoke plate into a structure of a left spoke plate and a right spoke plate, compared with the traditional single-spoke turbine disc, a weight reduction of 10% to 15% can be achieved, further reducing the weight of the turbine disc and achieving a lightweight design.

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

[0012] In the optional scheme of the above embodiment, the shape of the cross section 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 and avoid waste. It can evenly disperse external forces in all directions and avoid stress concentration. While realizing the lightweight design of the entire turbine disk, it also makes the turbine disk have high strength and stability.

[0013] As an optional solution of the above embodiment, the two adjacent honeycomb annular cavities are spaced apart by a honeycomb partition.

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

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

[0016] As an optional solution of the above embodiment, each honeycomb annular cavity is evenly provided with a plurality of inter-honeycomb cooling holes in an circumferential direction, and the diameter of each inter-honeycomb cooling hole is greater than 1 mm, ensuring that the inter-honeycomb cooling holes have better manufacturability.

[0017] As an optional solution to the above embodiment, the disc rim, spoke plate and hub are integrally formed by additive manufacturing.

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

[0019] 1. The honeycomb biomimetic structure turbine disk of an aircraft engine in the present invention can achieve a weight reduction of 10% to 15% compared with a traditional single-spoke turbine disk by arranging the spokes into a structure of left and right spokes. At the same time, multiple interconnected honeycomb annular cavities are arranged inside the left and right spokes and the hub, further reducing the weight of the turbine disk. Compared with a traditional solid turbine disk, the structural weight reduction is 20% to 40%. Furthermore, since the specific strength of the honeycomb annular cavity is much higher than that of the traditional solid structure, it can withstand a greater load at the same weight, thereby realizing a lightweight design of the turbine disk and solving the problem of excessive weight of aircraft engine turbine disks in the prior art, which increases the overall deadweight load of the aircraft engine and reduces the overall performance of the aircraft.

[0020] 2. The honeycomb bionic structure turbine disk of an aircraft engine in the present invention connects the center of the turbine disk with the tongue and groove area of the disk rim by providing a cooling air inlet, a disk rim air collecting cavity and a cooling air outlet, so that the cooling air can cool the disk rim, the spoke plate, the hub and the blades in the tongue and groove area of the disk rim, thereby improving the cooling efficiency of the turbine disk and reducing the operating temperature of the turbine disk by 50°C to 100°C.

[0021] 3. In the honeycomb bionic structure turbine disk of an aircraft engine in the present invention, the cross-section of the honeycomb annular cavity is preferably a regular hexagon. The honeycomb annular cavity with a regular hexagonal structure can absorb a large amount of energy through its own deformation when impacted, and has good impact resistance. Under the premise of meeting the strength requirements of the entire turbine disk, the weight of the turbine disk is effectively reduced, the cooling efficiency is improved, and the performance of the aircraft engine is thereby enhanced. BRIEF DESCRIPTION OF THE DRAWINGS

[0022] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following is a brief introduction to the drawings required for use in the embodiments. Obviously, the drawings described below are only some embodiments of the present invention. For those of ordinary skill in the art, other drawings can be obtained based on these drawings without inventive work. The above and other objects, features and advantages of the present invention will become more apparent through the accompanying drawings. The same reference numerals indicate the same parts throughout the drawings. The drawings are not intentionally scaled to their actual sizes, and the focus is on illustrating the main purpose of the present invention.

[0023] Figure 1 This is a schematic diagram of the meridian disk structure of the honeycomb bionic structure turbine disk of an aircraft engine.

[0024] Figure 2 for Figure 1 Schematic diagram of the structure in the AA direction.

[0025] Figure 3 for Figure 1 Schematic diagram of the enlarged structure at point B in the middle.

[0026] Figure 4 This is a schematic diagram of the three-dimensional cross-sectional structure of the honeycomb bionic structure turbine disk of an aircraft engine.

[0027] Among them, 1. disk rim; 2. left spoke plate; 3. right spoke plate; 4. wheel hub; 5. cooling air inlet; 6. disk rim air collecting cavity; 7. turbine disk symmetry center line; 8. cooling air outlet; 9. disk rim tongue and groove area; 10. honeycomb annular cavity; 11. inter-honeycomb cooling holes; 12. honeycomb partition. DETAILED DESCRIPTION

[0028] To make the objectives, technical solutions, and advantages of the embodiments of the present invention more clear, the technical solutions of the embodiments of the present invention will be clearly and completely described below in conjunction with the accompanying drawings of the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Generally, the components of the embodiments of the present invention described and shown in the drawings herein can be arranged and designed in various different configurations.

[0029] Therefore, 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 invention as claimed, but rather merely represents selected embodiments of the present invention. All other embodiments derived by persons of ordinary skill in the art based on the embodiments of the present invention without creative effort shall fall within the scope of protection of the present invention.

[0030] It should be noted that similar reference numerals and letters denote similar items in the following drawings, and therefore, once an item is defined in one drawing, it does not need to be further defined or explained in subsequent drawings.

[0031] like Figures 1 to 4 As shown, the present invention provides a honeycomb biomimetic turbine disk for an aircraft engine, comprising a rim 1, a spoke, and a hub 4. The hub 4 is located within the rim 1, and the spoke is located between the rim 1 and the hub 4. The outer side of the spoke is fixedly connected to the inner annular surface of the rim 1, and the inner side of the spoke is fixedly connected to the outer annular surface of the hub 4. Multiple honeycomb annular cavities 10 are provided within the spoke and hub 4, each of which is circumferentially arranged about the turbine disk's symmetric centerline 7. Adjacent honeycomb annular cavities 10 are spaced apart and connected by inter-honeycomb cooling holes 11. Multiple cooling air inlets 5 are provided on the inner annular surface of the hub 4, and the multiple cooling air inlets 5 are connected to the multiple honeycomb annular cavities 10 near the inner annular surface of the hub 4.

[0032] A circle of disk rim air collecting cavity 6 is arranged in a ring direction with the turbine disk symmetry center line 7 as the axis on the top of the disk rim 1, and the disk rim air collecting cavity 6 is connected with multiple honeycomb annular cavities 10 located at the outer edge of the spoke plate; a disk rim tongue and groove area 9 is arranged in a ring direction with the turbine disk symmetry center line 7 as the axis on the outer ring surface of the disk rim 1; the disk rim tongue and groove area 9 is connected to the disk rim air collecting cavity 6 through a cooling air outlet hole 8.

[0033] The basic principle of the honeycomb biomimetic turbine disk for aircraft engines is as follows: cooling air enters the honeycomb annular cavity 10 from the cooling air inlet 5 located at the center of the turbine disk. It then passes through the intercellular cooling holes 11 within the honeycomb annular cavity 10, into the hub 4 and the inner cooling disk body, and finally into the turbine disk rim 1 through the cooling air outlet 8. The cooling air enters the tongue-and-groove area 9 of the rim, and further into the turbine rotor blades, cooling them. This improves the cooling efficiency of the turbine disk and can reduce its operating temperature by 50°C to 100°C. Multiple interconnected honeycomb annular cavities 10 are provided within the inner surfaces of the spokes and hub 4. This not only reduces the weight of the turbine disk, achieving a weight reduction of 20% to 40% compared to traditional solid turbine disks, but also, because the honeycomb annular cavities 10 have a much higher specific strength than traditional solid structures, they can withstand greater loads for the same weight. The honeycomb annular cavities 10 also have a much higher specific stiffness than traditional solid structures, providing improved deformation resistance for the same weight. When impacted, the honeycomb annular cavities 10 absorb significant energy through their own deformation, resulting in excellent impact resistance. 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 needs of different application scenarios, and has strong designability.

[0034] To sum up, a plurality of honeycomb annular cavities 10 are provided inside the spoke plate and the hub 4, so that the stiffness, damage tolerance and lightweight design of the entire turbine disc can be better guaranteed while meeting the strength requirements, thereby solving the problem in the prior art that the turbine disc of an aircraft engine is too heavy, which increases the overall deadweight load of the aircraft engine and reduces the overall performance of the aircraft.

[0035] Preferably, but not limited to, the disk rim 1, spokes, and hub 4 are integrally formed by additive manufacturing. The provision of multiple honeycomb annular cavities 10 complicates the structure and manufacturing process of the turbine disk. Additive manufacturing technology, by layering materials to create a three-dimensional structure, can produce complex structural components, such as flow channels and thin walls, that are difficult to achieve using traditional processing methods. This makes it particularly suitable for the manufacture of aircraft engine turbine disks.

[0036] Preferably, the spoke plate includes a symmetrically arranged left spoke plate 2 and a right spoke plate 3, the left spoke plate 2 and the right spoke plate 3 are arranged at an angle, an angle is set between the left spoke plate 2 and the right spoke plate 3, the outer side surfaces of the left spoke plate 2 and the right spoke plate 3 are connected, and the inner side surfaces of the left spoke plate 2 and the right spoke plate 3 are forked and separated; in the present application, by arranging the spoke plate into a structure of a left spoke plate 2 and a right spoke plate 3, compared with the traditional single-spoke turbine disc, a weight reduction of 10% to 15% can be achieved, further reducing the weight of the turbine disc and realizing a lightweight design.

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

[0038] In the optional scheme of the above embodiment, the shape of the cross section of the honeycomb annular cavity 10 can be preferably a regular hexagon, so that multiple honeycomb annular cavities 10 form a honeycomb structure. The regular hexagon can utilize space more efficiently than other polygons to avoid waste, and can evenly disperse external forces in all directions to avoid stress concentration. While realizing the lightweight design of the entire turbine disk, it also makes the turbine disk have high strength and stability.

[0039] Adjacent honeycomb annular cavities 10 are separated by a honeycomb partition 12. The thickness of the honeycomb partition 12 is greater than or equal to 2 mm. 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 multiple honeycomb annular cavities 10 is increased; in areas with low stress, the thickness of the honeycomb partition 12 is reduced, and the distribution density of the multiple honeycomb annular cavities 10 is reduced.

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

[0041] Specifically, each honeycomb annular cavity 10 is uniformly provided with a plurality of inter-honeycomb cooling holes 11 in an circumferential direction, and the diameter of each inter-honeycomb cooling hole 11 is greater than 1 mm, ensuring that the inter-honeycomb cooling hole 11 has good manufacturability.

[0042] In summary, the honeycomb bionic structure turbine disk of an aircraft engine in the present invention has a structural weight reduction of 20% to 40% compared with a traditional solid turbine disk, and can withstand a greater load at 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 disk edge air collecting cavity 6, the cooling air outlet 8, the disk edge tongue and groove area 9, the honeycomb annular cavity 10 and the honeycomb inter-honeycomb cooling hole 11, which can guide the cooling air to the blades in the disk edge tongue and groove area 9, so that the turbine disk has a higher cooling efficiency and can reduce the operating temperature of the turbine disk by 50°C to 100°C.

[0043] The above are merely preferred embodiments of the present invention and are not intended to limit the present invention. Those skilled in the art will readily appreciate that the present invention is susceptible to various modifications and variations. Any modifications, equivalent substitutions, or improvements made within the spirit and principles of the present invention shall be included within the scope of protection of the present invention.

Claims

1. An aircraft engine honeycomb bionic structure turbine disk, characterized in that: Including rim, spokes and hub; The hub is located inside the disc rim, the spoke is located between the disc rim and the hub, the outer side of the spoke is fixedly connected to the inner annular surface of the disc rim, and the inner side of the spoke is fixedly connected to the outer annular surface of the hub; A plurality of honeycomb annular cavities are provided inside the spoke plate and the hub, and each of the honeycomb annular cavities is arranged in an annular direction with the symmetrical center line of the turbine disk as the axis; Two adjacent honeycomb annular cavities are spaced apart and connected through inter-honeycomb cooling holes; A plurality of cooling air inlets are provided on the inner annular surface of the hub, and the plurality of cooling air inlets are communicated with a plurality of honeycomb annular cavities close to the inner annular surface of the hub; A disk rim air collecting cavity is provided in a circle on the top of the disk rim with the symmetric center line of the turbine disk as the axis, and the disk rim air collecting cavity is connected with multiple honeycomb annular cavities located at the outer edge of the spoke plate; a disk rim tongue and groove area is provided on the outer ring surface of the disk rim with the symmetric center line of the turbine disk as the axis; the disk rim tongue and groove area is connected to the disk rim air collecting cavity through a cooling air outlet hole.

2. The aviation engine honeycomb bionic structure turbine disk according to claim 1, characterized in that: The spoke plate includes a symmetrically arranged left spoke plate and a right spoke plate, the left spoke plate and the right spoke plate are tilted, an angle is set between the left spoke plate and the right spoke plate, the outer side surfaces of the left spoke plate and the right spoke plate are connected, and the inner side surfaces of the left spoke plate and the right spoke plate are forked and separated.

3. The aviation engine honeycomb biomimetic structure turbine disk according to claim 1, characterized in that: The cross section of each honeycomb annular cavity is in the shape of a regular pentagon, a regular hexagon, a circle or an ellipse.

4. The aviation engine honeycomb biomimetic structure turbine disk according to claim 1, characterized in that: The two adjacent honeycomb annular cavities are spaced apart by a honeycomb partition.

5. The aviation engine honeycomb bionic structure turbine disk according to claim 4, characterized in that: The thickness of the honeycomb baffle is greater than or equal to 2 mm. The thickness of the honeycomb baffle is adjusted according to the strength design requirements of the turbine disk. In areas with high stress, the thickness of the honeycomb baffle is increased, the side length is reduced, and the distribution density of the multiple honeycomb annular cavities is increased; in areas with low stress, the thickness of the honeycomb baffle is reduced, and the distribution density of the multiple honeycomb annular cavities is reduced.

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

7. The aviation engine honeycomb biomimetic structure turbine disk according to claim 1, characterized in that: Each honeycomb annular cavity is uniformly provided with a plurality of inter-honeycomb cooling holes in an circumferential direction, and the diameter of each inter-honeycomb cooling hole is greater than 1 mm.

8. The aviation engine honeycomb bionic structure turbine disk according to any one of claims 1 to 7, characterized in that: The disc rim, spoke plate and hub are integrally formed by additive manufacturing.

Images