WO2020199370A1

WO2020199370A1 - Aero-engine turbine blade having cooling structure

Info

Publication number: WO2020199370A1
Application number: PCT/CN2019/091937
Authority: WO
Inventors: 高晟钧
Original assignee: 高晟钧
Priority date: 2019-04-02
Filing date: 2019-06-19
Publication date: 2020-10-08
Also published as: CN109973154B; CN109973154A

Abstract

An aero-engine turbine blade having a cooling structure, comprising a leading end (1), a pressure surface (2), a suction surface (3) and a trailing end (4); the leading end (1), the pressure surface (2), the suction surface (3) and the trailing end (4) are enclosed to form a hollow structure, and the hollow structure provides a flow channel for a cold flow. At least one of the leading end (1), the pressure surface (2) and the suction surface (3) is a thin-wall structure provided, at a central portion thereof, with a heat exchange cavity, and the heat exchange cavity is in communication with the hollow structure. The heat exchange cavity is in communication with the outside of the aero-engine turbine blade through a variable-diameter air outlet hole, and the area of an inlet through which the variable-diameter air outlet hole is connected to the heat exchange cavity is smaller than the area of an outlet through which the variable-diameter air outlet hole is connected to the outside of the aero-engine turbine blade.

Description

Aeroengine turbine blade with cooling structure

Technical field

The invention relates to the technical field of aeroengine refrigeration, in particular to an aeroengine turbine blade with a cooling structure.

Background technique

With the rapid development of aero-engine technology, the turbocharger ratio of the aero-engine compressor and the inlet temperature in front of the turbine have greatly increased, which will inevitably lead to an increase in the thermal load on the turbine blades, which will cause them to bear very serious thermal stress. In order to solve this problem, in addition to the continuous development of new materials and new processes, one of the decisive factors is the use of advanced and efficient enhanced cooling technology for the turbine blades. The cooling technology of turbine blades is mainly carried out from two aspects: one is to strengthen the disturbance of the cooling air inside the turbine blade, and to increase the heat exchange area inside the turbine blade; the other is to use film cooling on the blade surface to effectively block high-temperature gas from the turbine blade. Convective heat transfer. However, the cooling effect of turbine blades is limited in either way.

Summary of the invention

In view of the above analysis, the present invention aims to provide an aero-engine turbine blade with a cooling structure to solve the problem of insufficient cooling of existing aero-engine turbine blades.

The purpose of the present invention is mainly achieved through the following technical solutions:

In the technical scheme of the present invention, an aero-engine turbine blade with a cooling structure. The aero-engine turbine blade includes a front end, a pressure surface, a suction surface and a tail end. The front end, pressure surface, suction surface and tail end are enclosed in a hollow structure, The hollow structure provides flow channels for cold flow;

At least one of the front end, the pressure surface and the suction surface is a thin-walled structure with a heat exchange cavity in the middle, and the heat exchange cavity is respectively connected with the hollow structure;

The heat exchange cavity and the exterior of the aero engine turbine blade are connected through an outlet reducing hole; the area of the inlet connecting the outlet reducing hole and the heat exchange cavity is smaller than the area of the outlet reducing hole and the outside of the aero engine turbine blade.

In the technical scheme of the present invention, a heat exchange cavity is provided at the front end, and the heat exchange cavity is communicated with the outside of the aero-engine turbine blade through the air outlet reducing hole;

The axis of the air outlet reducing hole is perpendicular to the plane tangent to the outer surface of the front end where the air outlet reducing hole is located.

In the technical scheme of the present invention, the diameter of the outlet reducing hole increases uniformly from the inlet to the outlet along the axial direction of the outlet reducing hole.

In the technical scheme of the present invention, the diameter of the gas outlet reducing hole along the axial direction of the gas outlet reducing hole changes from inlet to outlet first, and then uniformly increases.

In the technical solution of the present invention, the pressure surface and/or the suction surface are provided with a heat exchange cavity, and the heat exchange cavity communicates with the outside of the aero-engine turbine blade through the outlet reducing hole;

The projection of the inlet of the outlet reducing hole on the outer surface of the aero-engine turbine blade is closer to the front end than the outlet of the outlet reducing hole.

In the technical scheme of the present invention, a heat exchange cavity is provided at the tail end, and the heat exchange cavity at the tail end communicates with the outside of the aero-engine turbine blade through a straight air outlet;

The pressure surface and the suction surface are tangent at the tail end, and both are tangent to the same plane, and the axis of the straight air outlet is parallel or coincident with the plane.

In the technical scheme of the present invention, the straight air outlet holes are uniformly arranged along the extension direction of the aero-engine turbine blades, and are evenly arranged on a plane perpendicular to the extension direction of the aero-engine turbine blades;

The heat exchange cavity at the tail end is provided with a supporting spoiler structure connecting the inner and outer surfaces of the aeroengine turbine blades.

In the technical solution of the present invention, the air outlet reducing holes are evenly arranged along the extension direction of the aero-engine turbine blades, and are evenly arranged on a plane perpendicular to the extension direction of the aero-engine turbine blades;

The heat exchange cavity is provided with a supporting spoiler structure connecting the inner and outer surfaces of the aeroengine turbine blades.

In the technical scheme of the present invention, an aeroengine includes: an air inlet, a compressor, a combustion chamber, a turbine, and a tail nozzle;

The blades of the turbine adopt the aero-engine turbine blades with cooling structure in the above technical solution;

The compressor is connected to the combustion chamber and shares the shell;

The intake port is arranged at the front end of the aero engine, and the inside of the compressor communicates with the outside through the intake port;

The tail nozzle is arranged at the tail of the combustion chamber, and the turbine is arranged in the tail nozzle.

The technical solution of the present invention can achieve at least one of the following effects:

1. In the present invention, the air outlet is set as a diameter-reducing hole with a small inside and a large outside, which can change the volume of the gas to lower the temperature inside and increase the cooling effect;

2. In the present invention, the air vents on the pressure surface and the suction side are biased toward the tail, and it is easier to generate an air film when the turbine blades rotate, thereby ensuring the thermal insulation effect of the air film. In addition, it can avoid the air outlet with the pressure surface and the suction surface. Toward the forward direction of the turbine, thereby increasing the energy utilization rate of the turbine operation.

In the present invention, the above technical solutions can also be combined with each other to realize more preferred combination solutions. Other features and advantages of the present invention will be described in the following specification, and part of the advantages may become obvious from the specification or be understood by implementing the present invention. The purpose and other advantages of the present invention can be realized and obtained through the content specified in the specification, claims and drawings.

Description of the drawings

The drawings are only used for the purpose of illustrating specific embodiments, and are not considered to be a limitation of the present invention. Throughout the drawings, the same reference signs represent the same components.

Figure 1 is an overall cross-sectional view in an embodiment of the present invention;

Figure 2-1 is a schematic diagram of Example 1 enlarged at A of the embodiment of the present invention;

2-2 is a schematic diagram of Example 2 enlarged at A of the embodiment of the present invention;

Figure 3-1 is a schematic diagram of Example 1 enlarged at B of the embodiment of the present invention;

3-2 is a schematic diagram of Example 2 enlarged at B of the embodiment of the present invention;

Figure 4-1 is a schematic diagram of Example 1 enlarged at C of the embodiment of the present invention;

4-2 is a schematic diagram of Example 2 enlarged at C of the embodiment of the present invention;

FIG. 5 is a schematic diagram of an enlarged example at D of the embodiment of the present invention;

Reference number

1- Front end; 2- Pressure surface; 3- Suction surface; 4- Tail end.

detailed description

The preferred embodiments of the present invention will be described below in detail with reference to the accompanying drawings. The accompanying drawings constitute a part of the application and are used together with the embodiments of the present invention to explain the principle of the present invention, and are not used to limit the scope of the present invention.

Example 1

In the field of aero-engines, because the turbine blades are in contact with the hot gas after combustion for a long time, the temperature is usually quite high, and no matter which alloy material is used for the turbine blades, its structural performance at high temperatures will often deteriorate. Regardless of the material, considering the high temperature mechanical properties of the alloy, the cooling effect of the turbine blade is limited in the end, so more and more research and development are turning to specific structures to make the air film formed on the surface of the page to realize the cooling or cooling of the blade. Insulation. However, once too many structural improvements are made to the turbine, its fluid mechanics performance will often be affected, thereby affecting the working efficiency of the turbine.

In the embodiment of the present invention, an air film is formed by arranging air outlets on the surface of the turbine blade, and at the same time, the direction of the air outlet is set reasonably, which can reduce the influence of the air outlet on the fluid performance of the turbine blade as much as possible while forming the air film. Combined with the small and large cross-sectional dimensions of the vent hole, it can further ensure the high temperature outside and low temperature inside, thereby significantly improving the cooling and heat insulation effects without affecting the working efficiency of the turbine.

Specifically, an embodiment of the present invention provides an aeroengine turbine blade with a cooling structure. As shown in FIG. 1, the aeroengine turbine blade includes: a front end 1, a pressure surface 2, a suction surface 3, and a tail end 4. The front end 1 , Pressure surface 2, suction surface 3 and tail end 4 are enclosed in a hollow structure, which provides a flow channel for cold flow; at least one of front end 1, pressure surface 2 and suction surface 3 is a thin-walled structure with a heat exchange cavity in the middle , And the heat exchange cavity is respectively connected with the hollow structure; the heat exchange cavity and the outside of the aero-engine turbine blade are connected through the outlet reducing hole; the area of the inlet connecting the outlet reducing hole and the heat exchange cavity is smaller than the outlet reducing hole and the aero engine turbine blade The area of the external exit.

In the embodiment of the present invention, the specific shape of the turbine blade is not limited, and a common turbine blade shape may be used. When at least two of the front end 1, the pressure surface 2, and the suction surface 3 are provided with heat exchange chambers, the flow rate and pressure of the high-temperature gas are different when the turbine blades are in operation. The pressure and flow rate of the flow also need to be adjusted accordingly, so they need to be isolated from each other. The cross-sectional size of the outlet diameter reducing hole is small inside and large outside. According to the ideal gas state equation, because the outside cross-sectional size becomes larger, the volume of the cold flow gas becomes larger and the temperature rises, so as to avoid the formation of too large with the high temperature gas outside the turbine blade The temperature difference can avoid excessive heat transfer of the high-temperature gas to the gas film, and avoid affecting the working efficiency of the turbine. At the same time, it can form a temperature-stable gas film to play a certain heat insulation effect and ensure the cooling effect.

As shown in Figures 2-1 and 2-2, in the embodiment of the present invention, the front end 1 is provided with a heat exchange cavity, and the heat exchange cavity communicates with the outside of the aero-engine turbine blade through the outlet reducing hole; the axis of the outlet reducing hole is The position of the air outlet reducing hole is perpendicular to the plane tangent to the outer surface of the front end 1. Considering that the front end 1 is on the windward side, the cold flow in the outlet hole after heat exchange needs a relatively large pressure to flow out of the outlet hole. In addition, as long as the cold flow can flow out, due to the flow of external high-temperature gas, the outflow cold flow can be A gas film is directly formed at the front end 1, so the axis of the outlet reducing hole is perpendicular to the plane tangent to the outer surface of the front end 1 where the outlet reducing hole is located, that is, parallel to the flow direction of the external high temperature gas.

In order to ensure that the reducing structure of the air outlet reducing hole can have a better heat insulation effect, the reducing part of the air outlet reducing hole needs to be uniformly changed. According to the ideal gas state equation, the larger the cone angle, the more obvious the temperature reduction effect will be. However, when the cone angle increases to a certain extent, the temperature reduction effect will not continue to increase significantly. On the contrary, the processing difficulty of the reduced diameter part will be greatly increased. To improve, considering the above two points comprehensively, the taper angle of the side wall of the reducing part of the outlet reducing hole is 30°-120°. In order to maximize the balance between the processing difficulty and the heat insulation effect, the reducing part of the outlet reducing hole The taper angle of the side wall is preferably 60°-90°.

Two specific examples are given. The diameter of the outlet reducing hole increases uniformly from the inlet to the outlet along the axis of the outlet reducing hole, that is, the overall diameter is uniform; the diameter of the outlet reducing hole is from the inlet to the outlet along the axis of the outlet reducing hole. The outlet is unchanged first, and then increased evenly, that is, part of the diameter is uniformly reduced and the rest is non-reducing.

As shown in Figures 3-1, 3-2, 4-1, 4-2, in the field of turbine blades, the pressure surface 2 usually refers to the convex curved surface, and the suction surface 3 usually refers to the concave curved surface, although one of the two is concave. It is convex, but the external high-temperature gas will not blow vertically to the outer surface of the impeller, but flow obliquely to the tail end 4 of the turbine blade. Therefore, if the vertical section of the gas outlet reducing hole is still used, a uniform gas film cannot be formed. On the contrary, it will increase the windward area, thereby reducing the efficiency of the turbine. In the embodiment of the present invention, the pressure surface 2 and/or the suction surface 3 are provided with a heat exchange cavity, and the heat exchange cavity is connected to the outside of the aero-engine turbine blade through an outlet reducing hole; the inlet of the outlet reducing hole is outside the aero engine turbine blade The projection of the surface is closer to the front end 1 than the outlet of the air outlet reducing hole. The air outlet reducing hole inclined to the tail end 4 can make the outflowing cold flow flow toward the tail end 4, which facilitates the formation of a uniform air film, and can reduce the windward area without excessive cold flow pressure. Considering the difficulty of processing and the uniformity of the gas film, the angle between the axis of the gas outlet reducing hole and the tangent plane where it is located is 45°-90°, preferably 60°-75°.

Similarly, in order to ensure that the reducing structure of the outlet reducing hole can have a better heat insulation effect, the reducing parts of the outlet reducing hole on the pressure surface 2 and the suction surface 3 need to be uniformly changed. Two specific examples are given. The diameter of the outlet reducing hole increases uniformly from the inlet to the outlet along the axis of the outlet reducing hole, that is, the overall diameter is uniform; the diameter of the outlet reducing hole is from the inlet to the outlet along the axis of the outlet reducing hole. The outlet is unchanged first, and then increased evenly, that is, part of the diameter is uniformly reduced and the rest is non-reducing.

The temperature of the end 4 of the turbine blade is not as high as other parts, but the end 4 can still be provided with a heat exchange cavity to further improve the cooling effect. Specifically, as shown in Figure 5, in the embodiment of the present invention, the tail end 4 is provided with a heat exchange cavity, and the heat exchange cavity of the tail end 4 communicates with the outside of the aero-engine turbine blade through the straight air outlet; the pressure surface 2 and the suction surface 3 is tangent to the tail end 4, and all tangent to the same plane, and the axis of the straight air outlet is parallel or coincident with the plane. The cold flow out of the tail end 4 will directly follow the external high temperature gas to flow away in the tangential direction of the tail end 4, so there is no need to consider the problem of gas film, just ensure that the outflow direction is the same as the flow direction of the high temperature gas. Due to the difficulty of processing, the tail 4 uses straight holes directly due to its relatively low temperature.

In order to ensure that the tail end 4 is more uniform to ensure the cooling effect, in the embodiment of the present invention, the straight air outlet holes are evenly arranged along the extension direction of the aero-engine turbine blades, and are evenly arranged on a plane perpendicular to the extension direction of the aero-engine turbine blades.

For the same consideration, and to ensure the uniformity of the air film on the surface of the turbine blade, in the embodiment of the present invention, the air outlet reducing holes are uniformly arranged along the extension direction of the aero-engine turbine blade, and are perpendicular to the extension direction of the aero-engine turbine blade. Set evenly on the plane of the

Since too many openings or too large openings will affect the overall structural strength of the aero-engine turbine blades, and too few openings or too small openings will affect the cooling effect of the aero-engine turbine blades. In the embodiment of the present invention, the outgassing The area of the variable diameter hole and the straight hole of the air outlet is 5%-20% of the area of the aero-engine turbine blade. In order to maximize the thickness and uniformity of the gas film without affecting the structural strength of the engine turbine blade, The sum of the area of the air outlet variable diameter hole and the air outlet straight hole is 5%-20% of the area of the aero engine turbine blade, preferably 7%-12%.

In addition, in order to further improve the heat exchange efficiency between the cold flow in the heat exchange cavity and the turbine blades, in the embodiment of the present invention, the heat exchange cavity is provided with a supporting spoiler structure that connects the inner and outer surfaces of the aeroengine turbine blades. , Can ensure the structural strength of the turbine blades, prevent structural damage during operation, and its turbulence function can also make the cold flow contact with the turbine blades more fully and improve the heat exchange efficiency.

Example 2

In an embodiment of the present invention, an aeroengine includes: an air inlet, a compressor, a combustion chamber, a turbine, and a tail nozzle;

The blades of the turbine adopt the aeroengine turbine blades with cooling structure in embodiment 1;

The compressor is connected to the combustion chamber and shares the shell;

In summary, the embodiment of the present invention provides an aero-engine turbine blade with a cooling structure. The present invention sets the air outlet as a diameter-reducing hole with a small inside and a large outside, which can change the gas volume and lower the temperature inside. The outer height further improves the cooling effect; in the present invention, the vent holes of the pressure surface 2 and the suction surface 3 are biased toward the tail, and it is easier to generate an air film when the turbine blades rotate, thereby ensuring the heat insulation effect of the air film. In addition, it can avoid pressure The air outlets of the surface 2 and the suction surface 3 face the advancing direction of the turbine, thereby improving the energy utilization rate of the turbine operation.

The above are only preferred specific embodiments of the present invention, but the protection scope of the present invention is not limited thereto. Any person skilled in the art can easily think of changes or changes within the technical scope disclosed by the present invention. All replacements shall be covered within the protection scope of the present invention.

Claims

An aero-engine turbine blade with a cooling structure, characterized in that the aero-engine turbine blade includes a front end (1), a pressure surface (2), a suction surface (3) and a tail end (4). (1) The pressure surface (2), the suction surface (3) and the tail end (4) are enclosed in a hollow structure, which provides a flow channel for cold flow;

At least one of the front end (1), the pressure surface (2), and the suction surface (3) is a thin-walled structure with a heat exchange cavity in the middle, and the heat exchange cavity is connected with the hollow structure respectively;

The heat exchange cavity is communicated with the outside of the aeroengine turbine blade through an outlet reducing hole; the inlet area of the outlet reducing hole connected with the heat exchange cavity is smaller than the outlet connecting the outlet reducing hole with the outside of the aeroengine turbine blade area.
The aero-engine turbine blade with a cooling structure according to claim 1, wherein the front end (1) is provided with a heat exchange cavity, and the heat exchange cavity is communicated with the outside of the aero-engine turbine blade through an outlet reducing hole;

The axis of the air outlet reducing hole is perpendicular to the plane tangent to the outer surface of the front end (1) at the location of the air outlet reducing hole.
The aero-engine turbine blade with a cooling structure according to claim 2, wherein the diameter of the outlet reducing hole increases uniformly from the inlet to the outlet along the axial direction of the outlet reducing hole.
The aero-engine turbine blade with a cooling structure according to claim 2, wherein the diameter of the outlet reducing hole is constant from the inlet to the outlet along the axis of the outlet reducing hole, and then uniformly increases.
The aero-engine turbine blade with a cooling structure according to claim 1, characterized in that the pressure surface (2) and/or the suction surface (3) is provided with a heat exchange cavity, and the heat exchange cavity is reduced in diameter through the outlet air The hole communicates with the outside of the aero engine turbine blade;

The projection of the inlet of the outlet reducing hole on the outer surface of the aero-engine turbine blade is closer to the front end (1) than the outlet of the outlet reducing hole.
The aero-engine turbine blade with a cooling structure according to claim 5, wherein the diameter of the outlet reducing hole increases uniformly from the inlet to the outlet along the axial direction of the outlet reducing hole.
The aero-engine turbine blade with a cooling structure according to claim 5, wherein the diameter of the outlet reducing hole is constant from the inlet to the outlet along the axial direction of the outlet reducing hole, and then uniformly increases.
The aero-engine turbine blade with a cooling structure according to claim 1, characterized in that the tail end (4) is provided with a heat exchange cavity, and the heat exchange cavity at the tail end (4) communicates with the aerospace through a straight air outlet hole. External communication of engine turbine blades;

The pressure surface (2) and the suction surface (3) are tangent at the tail end (4), and both are tangent to the same plane, and the axis of the straight air outlet hole is parallel or coincident with the plane.
The aero-engine turbine blade with a cooling structure according to claim 8, wherein the straight air outlet holes are evenly arranged along the extension direction of the aero-engine turbine blade, and are arranged perpendicular to the aero-engine turbine blade. Evenly set on the plane of the extension direction;

The heat exchange cavity of the tail end (4) is provided with a supporting spoiler structure connecting the inner and outer surfaces of the aero-engine turbine blade; the outlet reducing holes are evenly arranged along the extending direction of the aero-engine turbine blade, and Evenly arranged on a plane perpendicular to the extending direction of the aero-engine turbine blade;

The heat exchange cavity is provided with a supporting spoiler structure connecting the inner and outer surfaces of the aeroengine turbine blades.
An aeroengine, characterized in that, the aeroengine includes: an air inlet, a compressor, a combustion chamber, a turbine, and a tail nozzle;

The blade of the turbine adopts the aeroengine turbine blade with a cooling structure according to any one of claims 1 to 9;

The compressor is connected to the combustion chamber and shares a shell;

The air intake passage is arranged at the front end of the aero engine, and the inside of the compressor communicates with the outside through the air intake passage;

The tail nozzle is provided at the tail of the combustion chamber, and the turbine is provided in the tail nozzle.