A double-wall structure of a high-pressure turbine guide vane

Abstract

The application belongs to the technical field of aero-engines, and particularly relates to a high-pressure turbine guide vane double-wall structure, which is divided into a film cooling area and a vortex cooling area according to the external gas heat exchange parameter of the outer side of an outer wall; the outer wall in the film cooling area is provided with film holes communicating between the outside and a panel cavity; the outer wall in the vortex cooling area is provided with recessed vortices, and the partition ribs are provided with communication holes communicating between adjacent panel cavities; the outer wall at the boundary of the vortex cooling area is provided with film holes communicating between the outside and the panel cavity; the impact holes, the recessed vortices and the disturbance ribs synergistically act to improve the thermal protection capability of the vane in a high-temperature gas environment.

Description

Technical Field

[0001] This application belongs to the field of aero-engine technology, and specifically relates to a double-walled structure for a high-pressure turbine guide vane. Background Technology

[0002] The turbine inlet temperature of the new generation engine is nearly 300K higher than that of the previous generation. Considering the temperature non-uniformity at the combustion chamber outlet and the deterioration of engine use, the temperature that the blades endure far exceeds the melting point of the blade material. Meanwhile, to avoid power loss due to excessive cooling air usage, the amount of available cooling air in future high thrust-to-weight ratio engines will be further reduced. For advanced aero-engines currently under development, the temperature resistance of single-crystal turbine air-cooled blades has become a bottleneck and constraint on achieving engine technical targets. A double-walled structure, as an efficient cooling solution, has been proposed and applied to advanced aero-engine turbine blades. Summary of the Invention

[0003] To address the aforementioned issues, this application provides a double-walled structure for a high-pressure turbine guide vane. The double-walled structure includes an inner wall and an outer wall located outside the inner wall. The inner wall forms a cooling inlet cavity. The interlayer between the outer and inner walls is formed by ribs distributed circumferentially along the blade, creating multiple circumferentially distributed laminated cavities. The inner wall has impact holes connecting the cooling inlet cavity and the laminated cavities. Based on the heat transfer temperature or coefficient on the outer side of the outer wall, the double-walled structure is divided into a film cooling region and a vortex cooling region. Regions with heat transfer temperatures or coefficients higher than a preset threshold are designated as vortex cooling regions; regions with heat transfer temperatures or coefficients lower than a preset threshold are designated as film cooling regions.

[0004] The outer wall of the air film cooling area has air film holes that connect the outside to the plate cavity;

[0005] The outer wall of the vortex cooling area has recessed vortices, and the partition rib has a connecting hole that connects to the adjacent layer cavity. The outer wall at the boundary between the vortex cooling area and the air film cooling area has an air film hole that connects to the outside and the layer cavity.

[0006] Preferably, the cavity of the shelf is provided with turbulence ribs.

[0007] Preferably, the turbulence ribs in the vortex cooling area are divided into multiple rows along the direction from the blade root to the blade tip, and a single row of recessed vortices is set between two adjacent rows of turbulence ribs.

[0008] Preferably, the trailing edge of the blade has a slit that connects the outside to the cooling air intake chamber.

[0009] Preferably, the friction ribs are staggered in the direction from the leaf root to the leaf tip.

[0010] The advantage of this application is that the outer wall of the vortex cooling zone does not need to be provided with air film holes. The cooling gas enters the next layer of plate cavity through the connecting hole, realizing step-by-step cooling and flow control, significantly improving the internal cooling efficiency, saving the amount of cooling gas, and improving the utilization rate of cooling gas.

[0011] The combined effect of impact, vortex structure and turbulence ribs cools the high heat load area and improves the thermal protection capability of the blades in the high temperature gas environment.

[0012] Due to improved cooling efficiency and uniformity, the temperature distribution on the blade surface is more reasonable, the temperature difference between the blade and the back side is smaller, and the thermal stress is less, thus extending the blade's lifespan. Attached Figure Description

[0013] Figure 1 This is a schematic diagram of the double-walled structure of the turbine guide vane in this design.

[0014] Figure 2 yes Figure 1 A partially enlarged schematic diagram of the vortex cooling region.

[0015] Figure 3 yes Figure 2 A cross-sectional view of the cavity of the double-walled plate of the turbine guide vane. Detailed Implementation

[0016] The embodiments of this application will now be described in detail with reference to the accompanying drawings. Figures 1-3 As shown, this application provides a double-wall structure for a high-pressure turbine guide vane. The double-wall structure includes an inner wall 2 and an outer wall 1 located outside the inner wall 2. The inner wall 2 forms a cooling inlet cavity 3. The interlayer between the outer wall 1 and the inner wall 2 is formed by partition ribs 10 distributed along the circumference of the blade, which form multiple circumferentially distributed layered cavities 9. The inner wall 2 has impact holes 5 that connect the cooling inlet cavity 3 and the layered cavities 9. According to the external gas heat exchange parameters outside the outer wall 1, the double-wall structure is divided into a film cooling region and a vortex cooling region.

[0017] In some alternative embodiments, the film cooling region is located on the blade basin side and the leading edge of the blade back side, and the vortex cooling region is located on the trailing edge of the blade back side. The outer wall 1 within the film cooling region has film holes 8 that connect the outside to the plate cavity 9.

[0018] The outer wall 1 within the vortex cooling region has a recessed vortex 7, and the partition rib has a connecting hole 6 that connects to the adjacent layer cavity 9. The outer wall 1 at the boundary between the vortex cooling region and the air film cooling region has an air film hole 8 that connects to the outside and the layer cavity 9.

[0019] In some alternative embodiments, the laminate cavity 9 is provided with turbulence ribs 4 to enhance the disturbance and uniformity of cooling airflow.

[0020] In some alternative embodiments, the turbulence ribs 4 within the vortex cooling region are divided into multiple rows along the direction from the blade root to the blade tip. A single row of recessed vortices 7 is arranged between adjacent rows of turbulence ribs 4, and a single row of recessed vortices 7 is arranged between two rows of transverse turbulence ribs 4. The single row of recessed vortices enhances heat transfer while having a relatively small impact on pressure drop. This structure can be achieved through a casting process, without the need for complex machining methods, making it suitable for mass production.

[0021] In some alternative embodiments, the friction ribs 4 are staggered in the direction from the blade root to the blade tip to enhance the disturbance and uniformity of the cooling airflow. After the cooling airflow passes through the transverse friction ribs 4, Figure 1 The arrows in the diagram indicate the direction of the cooling airflow, which further impacts the recessed vortices 7 arranged on the outer wall 1 and inner wall 2. These recessed vortices 7 induce a strong secondary flow of the cooling airflow, forming a large-scale vortex, thereby significantly enhancing the local heat transfer intensity and improving the impact cooling efficiency. After completing the impact cooling of the recessed vortexes 7, the cooling airflow enters the next layer cavity through the connecting holes 6 on the partition ribs set between adjacent layer cavity 9, realizing the stepwise distribution and flow control of the cooling airflow. Finally, the cooling airflow is discharged through the film cooling holes 8 on the outer wall 1 or the slits on the blade trailing edge that connect the outside to the cooling inlet cavity 3, achieving blade back cooling.

[0022] In some alternative implementations, the heat transfer parameters of the external gas include heat transfer temperature and heat transfer coefficient.

[0023] In some alternative implementations, regions with heat transfer temperatures and / or heat transfer coefficients exceeding a preset threshold are designated as vortex cooling regions. By analyzing the distribution of gas heat transfer temperature and heat transfer coefficient outside the blades, recessed vortices 7 are arranged inside the laminate cavity corresponding to regions with high gas heat transfer intensity and the "throat" area where the amount of cold air needs to be controlled. Figure 3 The diagram shows a structure in which recessed vortices 7 are arranged in the cavity 9 of the back layer of the blade. The position of the recessed vortex 7 and the position of the cold gas outlet can be selected according to the heat exchange characteristics of the external gas of the blade.

[0024] The advantage of this application is that the outer wall of the vortex cooling zone does not need to be provided with air film holes 8. The cooling gas enters the next layer plate cavity 9 through the connecting hole, realizing step-by-step cooling and flow control, which significantly improves the internal cooling efficiency, saves the amount of cooling gas, and improves the utilization rate of cooling gas.

[0025] The impact hole 5, the recessed vortex 7 structure and the turbulence rib 4 work together to enhance the cooling effect in the high heat load area and improve the thermal protection capability of the blades in the high temperature combustion environment.

[0026] Due to improved cooling efficiency and uniformity, the surface temperature distribution of the blades is more reasonable, the temperature difference between the blade back side and the blade tip is smaller, and the thermal stress is less, thereby extending the blade life and improving the reliability and safety of turbine components.

[0027] The above description is merely a specific embodiment of this application, but the scope of protection of this application is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in this application should be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.

Claims

1. A double-walled structure for a high-pressure turbine guide vane, the double-walled structure comprising an inner wall (2) and an outer wall (1) located outside the inner wall (2), the inner wall (2) forming a cooling intake cavity (3), the interlayer between the outer wall (1) and the inner wall (2) forming a plurality of circumferentially distributed layered cavities (9) by circumferentially distributed ribs, the inner wall (2) having impact holes (5) connecting the cooling intake cavity (3) and the layered cavities (9), characterized in that, Based on the external gas heat exchange parameters on the outer side of the outer wall (1), the double-wall structure is divided into a gas film cooling region and a vortex cooling region. The outer wall (1) within the air film cooling area has air film holes (8) that connect the outside to the plate cavity (9). The outer wall (1) within the vortex cooling area has a recessed vortex (7), and the partition rib has a connecting hole (6) that connects to the adjacent layer cavity (9). The outer wall (1) at the boundary of the vortex cooling area has a film hole (8) that connects to the outside and the layer cavity (9). A flow-tightening rib (4) is provided inside the cavity of the laminate (9); The turbulence ribs (4) in the vortex cooling area are divided into multiple rows along the radial direction, and a single row of recessed vortices (7) is set between two adjacent rows of turbulence ribs (4).

2. The double-walled structure of the high-pressure turbine guide vane as described in claim 1, characterized in that, The trailing edge of the blade has a slit that connects the outside to the cooling intake chamber (3).

3. The double-walled structure of the high-pressure turbine guide vane as described in claim 2, characterized in that, The ribs (4) are staggered in the radial direction.

4. The double-walled structure of the high-pressure turbine guide vane as described in claim 1, characterized in that, The heat transfer parameters of external gas include heat transfer temperature and heat transfer coefficient.

5. The double-walled structure of the high-pressure turbine guide vane as described in claim 1, characterized in that, Regions with heat exchange temperatures and / or heat exchange coefficients higher than a preset threshold are designated as vortex cooling regions.

Images