A cooling enhancement device, a turbine component comprising the same, and an aeroengine or gas turbine

Abstract

A cooling enhancement device comprising a first stage turbine disc flange and a baffle between the first stage turbine disc flange and a second stage turbine disc fixed to the second stage turbine disc in a rotating-to-rotating disc cavity connected together by a connecting device, characterised in that a second flange is connected to the first stage turbine disc flange on a side distal from the second stage turbine disc, a cooling ring parallel to the disc cantilever is fixed to the second flange to form a first guide channel to direct a cooling air flow within the disc cavity through the first guide channel, cooling the disc cantilever and to a dovetail slot area on the second stage turbine disc, the first flange being part of a first stage turbine disc.

Description

Technical Field

[0001] This invention relates to the field of turbines, and more particularly to a cooling enhancement device, turbine components including the device, and aircraft engines or gas turbines. Background Technology

[0002] Because turbine disks operate under harsh thermal and centrifugal loads for extended periods, thermal management of the turbine disks is a key design consideration to ensure the reliability and lifespan of the engine throughout its service life. A suitable cooling solution is essential for designers. Currently, turbine disks are primarily cooled by blowing cooling air from the compressor. This cooling air, after being introduced from the compressor, reaches the turbine disk chamber, where it is cooled before flowing through baffles or other structures to the tenon and mortise area or into the main flow.

[0003] The original cooling solution for the turbine rotor (such as...) Figure 1 As shown, cooling air 7 introduced by the compressor flows into the first-stage turbine disk 1 through the center of the first-stage turbine disk 1. The second-stage turbine disk 6 forms a rotor-to-rotor disk cavity 2, which cools the disk cavity 2 and the tenon area 4. However, this scheme has limited cooling effect on the cantilever 3 of the first-stage turbine disk. The main reason is that the total pressure decreases as the cooling airflow flows along the process, and the gas flow effect is not obvious near the disk cantilever position. The disk cantilever structure is generally designed to be thin. Under the influence of thermal load and centrifugal load, the cantilever structure will have a large "barrel-shaped" deformation and generate excessive thermal stress. In severe cases, the sealing structure at the rotor and stator will rub against each other, which will greatly affect the sealing effect between the rotor stages and thus have an adverse effect on the working state of the turbine components. In addition, excessive thermal stress is also detrimental to the rotor strength.

[0004] Under such circumstances, if the cooling effect at the cantilever is to be enhanced, the amount of cooling air drawn from the compressor can be increased. However, this will inevitably affect the performance of the whole machine. Therefore, the present invention improves the cooling effect without increasing the amount of air drawn from the compressor. Summary of the Invention

[0005] The key point to be protected by this patent is a cooling method for a turbine disk cantilever structure and a cooling device for a turbine disk cantilever structure. Any scheme modified based on the concept of this patent shall be within the scope of protection of this patent and may be used in aircraft engines or gas turbines.

[0006] A cooling enhancement device includes a first-stage turbine disk flange fixed to a second-stage turbine disk in a rotary disk cavity and connected together by a connecting device, and a baffle between the first-stage turbine disk flange and the second-stage turbine disk. The device is characterized in that a second flange is connected to the side of the first-stage turbine disk flange away from the second-stage turbine disk, and a cooling ring parallel to the disk cantilever is fixed on the second flange to form a first guide channel, thereby guiding cooling air in the disk cavity to flow through the first guide channel, cooling the disk cantilever, and reaching the tenon region on the second-stage turbine disk.

[0007] According to a preferred embodiment of the present invention, fins with a hardness lower than that of the first-stage turbine disk cantilever are provided on the cooling ring.

[0008] According to a preferred embodiment of the present invention, through holes are provided in the first-stage turbine disk flange and the baffle to form a second guiding channel for guiding cooling air.

[0009] According to a preferred embodiment of the present invention, a venting groove is provided between the second-stage turbine disk and the baffle to form a third guiding channel for guiding cooling air.

[0010] According to a preferred embodiment of the present invention, the fins, the first channel and the second channel are arranged along the axis of the engine.

[0011] According to a preferred embodiment of the present invention, the connecting device is a bolt.

[0012] According to a preferred embodiment of the present invention, the rotary-rotor cavity is formed by a first-stage turbine disk, a second-stage turbine disk, and a first-stage turbine disk cantilever connecting the first-stage turbine disk and the second-stage turbine disk.

[0013] The present invention also relates to a turbine component, characterized in that it includes the cooling enhancement device according to the above description.

[0014] Preferably, it includes a turbine disk and blades tenoned to the turbine disk.

[0015] The present invention also relates to an aircraft engine or gas turbine, characterized in that it includes the turbine component described above.

[0016] Turbine components are one of the main components of aero engines. They are turbine devices that convert the potential energy of high-temperature and high-pressure gas combustion into mechanical energy to drive fans, compressors, propellers, helicopter rotors, and accessory systems. Their working environment is extremely harsh. As the performance of aero engines improves, the requirements for turbine components become increasingly stringent. In addition to selecting appropriate materials, proper cooling methods and cooling flow paths play an important role in improving the turbine's working environment and have a significant impact on the lifespan, reliability, and performance of aero engines throughout their service life.

[0017] This invention has a simple structure and is easy to implement. It improves the cooling efficiency of the turbine disk rotation chamber without increasing the cooling air volume, enhances the cooling effect inside the turbine disk chamber, and plays a positive role in improving the "barrel-shaped" deformation and thermal stress state of the turbine disk cantilever, thereby increasing the safety and stability of engine operation. Attached Figure Description

[0018] Figure 1 This is a cooling structure based on existing technology;

[0019] Figure 2 and Figure 3 Views of the cooling enhancement device of the present invention are shown respectively. Detailed Implementation

[0020] like Figure 2 and 3 As shown, a cooling enhancement device 100 includes a second flange 12 fixed to a second-stage turbine disk 106 in a disk cavity 102 and connected together by a connecting device 1061, and a baffle 1062 between a first-stage turbine disk flange 1063 and a second-stage turbine disk 106. The second flange 12 is connected to the side of the first-stage turbine disk flange 1063 away from the second-stage turbine disk 106. A cooling ring 11 parallel to the disk cantilever 103 is fixed on the second flange 12 to form a first guide channel, thereby guiding the cooling air in the disk cavity to flow through the first guide channel, cooling the disk cantilever 103, and reaching the tenon area 104 on the second-stage turbine disk.

[0021] like Figure 2 and 3 As shown, according to a preferred embodiment of the present invention, the cooling ring 11 is provided with fins 10 with a hardness lower than that of the wheel cantilever 103.

[0022] like Figure 2 and 3 As shown, according to a preferred embodiment of the present invention, through holes are provided on the first-stage turbine disk flange 1063 and the baffle 1062 to form a second guide channel for guiding cooling air.

[0023] like Figure 2 and 3 As shown, according to a preferred embodiment of the present invention, a gap is provided between the second-stage turbine disk 106 and the baffle 1062 to form a third guide channel for guiding cooling air.

[0024] According to a preferred embodiment of the present invention, the fins, the first channel and the second channel are arranged along the axis of the engine.

[0025] like Figure 3 As shown, according to a preferred embodiment of the present invention, the connecting device is a bolt 1061.

[0026] According to a preferred embodiment of the present invention, the disk cavity is formed by a first-stage turbine disk, a second-stage turbine disk, and a first-stage turbine disk cantilever connecting the first-stage turbine disk and the second-stage turbine disk.

[0027] The present invention also relates to a turbine component, characterized in that it includes the cooling enhancement device according to the above description.

[0028] Preferably, it includes a turbine disk and blades tenoned to the turbine disk.

[0029] The present invention also relates to an aircraft engine or gas turbine, including the aforementioned turbine component.

[0030] The cooling device of the turbine disk cantilever structure of the present invention (e.g.) Figure 2 (As shown) It has an air system hole 8 opened in the baffle, an air system hole 9 opened in the first-stage turbine disk flange, fins 10 arranged along the cooling ring 11, and a flange 12 connecting the entire structure to the rotor system. The air system holes 8, 9 and fins 10 are evenly distributed along the engine axis. The hardness of the fin material is lower than that of the first-stage turbine disk cantilever 3, so that the first-stage turbine disk cantilever will not wear after the fins come into contact with it. 10 can also be omitted as it is not a necessary part, but the gap between the cooling cantilever 11 and the first-stage turbine disk cantilever must be ensured by other means. The cooling device is bolted to the turbine rotor, which will not affect the original assembly. The cooling device is an independent device, and can be removed separately if damaged, making maintenance convenient. The entire cooling device changes the original cooling channel (such as...). Figure 3 As shown in the figure, based on the original cooling scheme, the direction of the cooling airflow will be guided by the cooling device. The cooling airflow passing through the cooling device can further improve the cooling efficiency. In this way, the deformation on the turbine disk cantilever caused by thermal load will be further reduced, and the thermal stress state will also be improved.

[0031] The technical principles and concepts of the present invention have been illustrated above with reference to the accompanying drawings, but this does not limit the scope of protection of the present invention. As long as they are within the scope of protection of the claims, those skilled in the art can make any improvements or combinations.

Claims

1. A cooling enhancement device, comprising a first-stage turbine disk flange fixed to a second-stage turbine disk in a rotary disk cavity and connected together by a connecting device, and a baffle between the first-stage turbine disk flange and the second-stage turbine disk, characterized in that, A second flange is connected to the side of the first-stage turbine disk flange away from the second-stage turbine disk. A cooling ring parallel to the disk cantilever is fixed on the second flange to form a first guide channel, thereby guiding the cooling air in the disk cavity to flow through the first guide channel, cooling the disk cantilever and reaching the tenon area on the second-stage turbine disk. The first-stage turbine disk flange is part of the first-stage turbine disk.

2. The cooling enhancement device according to claim 1, characterized in that, The cooling ring is provided with fins with a hardness lower than that of the wheel cantilever.

3. The cooling enhancement device according to claim 2, characterized in that, Through holes are provided in the first-stage turbine disk flange and the baffle to form a second guide channel for guiding cooling air.

4. The cooling enhancement device according to claim 3, characterized in that, A ventilation groove is provided between the second-stage turbine disk and the baffle to form a third guiding channel for guiding cooling air.

5. The cooling enhancement device according to claim 4, characterized in that, The screw blades, the first channel, and the second channel are arranged along the axis of the engine.

6. The cooling enhancement device according to claim 5, characterized in that, The connecting device is a bolt.

7. The cooling enhancement device according to claim 6, characterized in that, The rotary disk cavity is formed by a first-stage turbine disk, a second-stage turbine disk, and a first-stage turbine disk cantilever connecting the first-stage turbine disk and the second-stage turbine disk.

8. A turbine component, characterized in that, Includes the cooling enhancement device according to any one of claims 1-7.

9. The turbine component according to claim 8, characterized in that, It includes a turbine disk and blades that are tenoned to the turbine disk.

10. An aircraft engine or gas turbine, characterized in that, Includes the turbine component as described in claim 8 or 9.

Images