Cooling structure for blade top of high-pressure turbine blade

Abstract

The invention discloses a cooling structure of a blade top of a high-pressure turbine blade, and belongs to the technical field of cooling structures, the cooling structure comprises a main flow channel, the bottom of the main flow channel is connected with a cold air channel, a double-layer wall is arranged in the cold air channel, air film holes are formed in the top of the double-layer wall in an array mode, and turbulent flow columns are arranged in the double-layer wall in an array mode. Impact holes are formed in the turbulent flow column in an array manner; according to the cooling structure for the blade top of the high-pressure turbine blade, the cold air coverage area of the blade top can be effectively increased, the blade top is prevented from being directly scoured by blade top gap leakage flow, separation bubbles formed by attachment of the blade top gap leakage flow and separation are weakened or even eliminated, the flow loss of the blade top area is reduced, and the temperature distribution of the blade top is improved; the cooling effect of the blade top close to the trailing edge area is effectively improved, and the service life of the turbine blade is prolonged.

Description

Technical Field

[0001] The present invention specifically relates to a cooling structure for the tip of a high-pressure turbine blade, belonging to the technical field of cooling structures. Background Art

[0002] With the increasing development of the field of aero-engines, the performance requirements for aero-engines are getting higher and higher. Increasing the turbine inlet temperature and the compressor pressure ratio can effectively improve the thrust-to-weight ratio and thermal efficiency of the engine.

[0003] Patent document CN119467008A discloses a gas turbine blade based on a partitioned composite cooling mechanism. "In the blade body, several different cooling channels are partitioned, mainly including double-layer S-shaped ribs in the leading edge area of the blade, double-layer wedge-shaped matrix ribs in the mid-chord area, and double-channel exhaust grooves in the trailing edge area; gas film holes are longitudinally arranged on the peripheral wall of the thin-walled outer layer corresponding to the positions of the double-layer S-shaped ribs and the double-layer wedge-shaped matrix ribs; jet holes are longitudinally arranged on the peripheral wall of the inner layer of the solid body, and the cooling gas passing through the double-layer S-shaped ribs and the double-layer wedge-shaped matrix ribs is sprayed onto the inner wall surface of the thin-walled outer layer through the jet holes. The present invention increases the overall heat transfer area inside the turbine blade, enhances turbulence and increases the disturbance of the airflow boundary layer; cooler air is introduced from inside or around the blade through the gas film holes and sprayed onto the blade surface to form a gas film, isolating the direct contact between the high-temperature airflow and the blade material surface, thereby reducing the heat load on the blade surface and improving the cooling effect of the cooling gas on the high-temperature blade." However, for a gas turbine blade based on a partitioned composite cooling mechanism in the above-mentioned published literature, it mainly considers partitioning several different areas of the blade, and mainly sets structures such as turbulator ribs, jet holes, and gas film holes. The advantage is that it increases the overall heat transfer area inside the turbine blade, enhances turbulence and the disturbance of the airflow boundary layer, and improves the comprehensive cooling efficiency. The disadvantage is that it does not well improve the flow loss at the blade tip.

[0004] Due to the increase in the turbine inlet temperature, the cooling demand for the hot-end components of the engine has risen. On the other hand, with the increase in the pressure ratio, the temperature of the high-pressure air has risen, resulting in a continuous decline in the quality of the cooling air. It has become very difficult to meet the application requirements of hot-end components such as turbines in aero-engines by relying on the development of the high-temperature resistance performance of the material itself. The thermal protection problem has become a key problem restricting the development of aero-engines. For this reason, we propose a cooling structure for the tip of a high-pressure turbine blade. Summary of the Invention

[0005] The purpose of the present invention is to provide a cooling structure for the tip of a high-pressure turbine blade in view of the deficiencies of the prior art, so as to achieve the purpose of cooling and reducing the temperature of the tip of the high-pressure turbine blade.

[0006] A cooling structure for the tip of a high-pressure turbine blade, comprising a mainstream channel, a cold air channel is connected to the bottom of the mainstream channel, a double-layer wall is arranged inside the cold air channel, air film holes are arranged in an array at the top of the double-layer wall, spoiler columns are arranged in an array inside the double-layer wall, and impact holes are arranged in an array inside the spoiler columns.

[0007] Further, an adiabatic surface is arranged at the top of the mainstream channel.

[0008] Further, the left side of the mainstream channel is the mainstream inlet, and the right side of the mainstream channel is the mainstream outlet.

[0009] Further, the bottom of the cold air channel is the cold air inlet

[0010] Beneficial effects:

[0011] The present invention can effectively increase the coverage area of the cold air at the tip, avoid the direct erosion of the tip clearance leakage flow on the tip, weaken or even eliminate the attachment of the tip clearance leakage flow and the separation bubble formed by separation, reduce the flow loss in the tip region, improve the temperature distribution at the tip, effectively enhance the cooling effect in the region near the trailing edge of the tip, and extend the service life of its turbine blade. Description of the drawings

[0012] Figure 1 is the overall structural schematic diagram of the present invention;

[0013] Figure 2 is the structural schematic diagram of the cold air channel of the present invention;

[0014] Figure 3 is the structural schematic diagram of the double-layer wall of the present invention;

[0015] Figure 4 is the structural schematic diagram of the spoiler column of the present invention.

[0016] In the figure: 1, mainstream channel; 2, cold air channel; 3, double-layer wall; 4, air film hole; 5, spoiler column; 6, impact hole; 7, adiabatic surface. Specific embodiments

[0017] Next, the technical solutions in the embodiments of the present invention will be clearly and completely described in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are only a part of the embodiments of the present invention, rather than all the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative efforts shall fall within the protection scope of the present invention.

[0018] Please refer to Figures 1-4As shown in the figure, a cooling structure for the tip of a high-pressure turbine blade includes a main flow channel 1. The bottom of the main flow channel 1 is connected to a cold air channel 2. Inside the cold air channel 2, there is a double-layer wall 3. At the top of the double-layer wall 3, there are film holes 4 arranged in an array. Inside the double-layer wall 3, there are turbulators 5 arranged in an array. Inside the turbulators 5, there are impingement holes 6 arranged in an array. High-temperature gas (1500K) enters from the left flow inlet of the main flow channel 1, passes through the tip surface, and then exits from the right outlet. 750K cooling air is injected from the cold air inlet of the bottom cold air channel 2, and a cooling air film is formed after passing through the internal structure. The adiabatic surface 7 can reduce heat dissipation and concentrate the cooling effect. The double-layer wall 3 is a composite structure composed of an inner and an outer wall surface. The inner wall is in contact with the main flow channel 1. The film holes 4 are arranged in an array, and the cold air is evenly ejected through the film holes 4 to form a low-temperature protective air film between the high-temperature gas and the wall surface. The turbulators 5 enhance the heat exchange efficiency between the cold air and the wall surface through turbulent perturbation and extend the residence time of the cold air. The cold air is ejected vertically at high speed through the impingement holes 6 onto the inner wall surface of the main flow channel 1 to achieve local impingement cooling (high heat transfer coefficient area). When the cold air flows through the cold air channel 2, it sequentially passes through the impingement holes 6 to efficiently impinge and cool the wall surface. The turbulators 5 strengthen the turbulent heat transfer of the cold air, and the film holes 4 form a heat-insulating air film to cover the high-temperature wall surface. Finally, it realizes: impingement cooling, film cooling, and turbulator strengthening triple protection, increases the residence time of the cold air, strengthens the convective heat transfer, can effectively increase the coverage area of the cold air at the tip, avoid the direct scouring of the tip by the leakage flow in the tip clearance, weaken or even eliminate the separation bubble formed due to the attachment and separation of the leakage flow in the tip clearance, reduce the flow loss in the tip area, improve the temperature distribution at the tip, effectively enhance the cooling effect in the area near the trailing edge of the tip, and extend the service life of the turbine blade.

[0019] An adiabatic surface 7 is provided at the top of the main flow channel 1. The adiabatic surface 7 can reduce heat dissipation and concentrate the cooling effect.

[0020] The left side of the main flow channel 1 is the main flow inlet, and the right side of the main flow channel 1 is the main flow outlet. The bottom of the cold air channel 2 is the cold air inlet. High-temperature gas (1500K) enters from the left flow inlet of the main flow channel 1, passes through the tip surface, and then exits from the right outlet. 750K cooling air is injected from the cold air inlet of the bottom cold air channel 2, and a cooling air film is formed after passing through the internal structure.

[0021] As a technical optimization solution of the present invention, when the cooling structure at the tip of the high-pressure turbine blade is first used, the high-temperature gas enters from the left flow inlet of the mainstream channel 1, passes through the tip surface and exits from the right outlet. The cooling air is injected from the bottom cooling air channel 2 through the cold air inlet, and forms a cooling air film after passing through the internal structure. The adiabatic surface 7 can reduce heat dissipation and concentrate the cooling effect. The double-layer wall 3 is a composite structure composed of an inner and an outer wall surface. The inner wall is in contact with the mainstream channel 1, and the film holes 4 are arranged in an array. The cold air is evenly ejected through the film holes 4 to form a low-temperature protective air film between the high-temperature gas and the wall surface. The turbulator 5 enhances the heat exchange efficiency between the cold air and the wall surface through turbulent disturbance and prolongs the residence time of the cold air. The cold air is ejected vertically at high speed through the impingement holes 6 onto the inner wall surface of the mainstream channel 1 to achieve local impingement cooling (high heat transfer coefficient area). When the cold air flows through the cold air channel 2, it sequentially impinges on the wall surface through the impingement holes 6 for efficient impingement cooling. The turbulator 5 strengthens the turbulent heat transfer of the cold air, and the film holes 4 form a heat-insulating air film to cover the high-temperature wall surface, finally achieving: triple protection of impingement cooling, film cooling, and turbulator strengthening, increasing the residence time of the cold air, and strengthening convective heat transfer.

[0022] For those skilled in the art, it is obvious that the present invention is not limited to the details of the above exemplary embodiments, and without departing from the spirit or basic characteristics of the present invention, the present invention can be implemented in other specific forms. Therefore, from any point of view, the embodiments should be regarded as exemplary and non-limiting. The scope of the present invention is defined by the appended claims rather than the above description. Therefore, all changes falling within the meaning and scope of the equivalent elements of the claims are intended to be included in the present invention. Any reference signs in the claims should not be regarded as limiting the claimed rights.

[0023] In addition, it should be understood that although this specification is described according to embodiments, not every embodiment only contains an independent technical solution. This narrative way of the specification is only for clarity. Those skilled in the art should regard the specification as a whole, and the technical solutions in each embodiment can also be appropriately combined to form other embodiments that can be understood by those skilled in the art.

Claims

1. A cooling structure for a high-pressure turbine blade tip, comprising a main flow channel (1), characterized in that: The bottom of the mainstream channel (1) is connected to a cold air channel (2), a double wall (3) is arranged inside the cold air channel (2), an array of air film holes (4) are arranged on the top of the double wall (3), an array of spoiler columns (5) are arranged inside the double wall (3), and an array of impact holes (6) are arranged inside the spoiler columns (5).

2. A cooling structure for a high-pressure turbine blade tip according to claim 1, characterized in that: A heat insulating surface (7) is provided on the top of the main flow channel (1).

3. A cooling structure for a high-pressure turbine blade tip according to claim 1, characterized in that: The left side of the mainstream channel (1) is a mainstream inlet, and the right side of the mainstream channel (1) is a mainstream outlet.

4. A cooling structure for a high-pressure turbine blade tip according to claim 1, characterized in that: The bottom of the cold air channel (2) is a cold air inlet.

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