Turbine stator blade with circumferentially assembled cooling structure

Abstract

The invention discloses a turbine stator blade provided with a circumferential assembly cooling structure, and relates to the technical field of gas turbines. The turbine stationary blade comprises a blade body and edge plates connected with the top end and the bottom end of the blade body, the two ends of each edge plate in the circumferential direction are circumferential assembling faces, the multiple turbine stationary blades are sequentially installed in the circumferential direction of a turbine main shaft, a sealing groove and a sealing piece are arranged between the circumferential assembling faces of the adjacent edge plates, and composite cooling units are further arranged on the circumferential assembling faces. The composite cooling unit communicates with a cold air channel on the outer side of the margin plate and the circumferential assembling face. The composite cooling unit comprises an impact hole, an impact cavity, an air entraining channel, a cooling cavity and an exhaust hole, a turbulent flow structure is arranged in the cooling cavity, cold air impacts the impact cavity through the impact hole, enters the cooling cavity through the air entraining channel, and flows out from the exhaust hole after turbulent flow heat exchange is completed. According to the turbine stator blade, cooling and sealing effects are achieved at the same time, safe and reliable operation of the circumferential sealing edge of the stator blade is guaranteed, the cold air utilization rate of the composite cooling unit is higher than that of a traditional structure, and the circulation efficiency of a gas turbine is improved.

Description

Technical Field

[0001] The present invention relates to the technical field of turbine equipment, and in particular to a turbine stator blade provided with a circumferentially assembled cooling structure. Background Art

[0002] With the continuous growth of energy demand and increasingly stringent environmental protection requirements, the development and application of heavy-duty gas turbines, as efficient and clean energy conversion equipment, have attracted widespread attention. The performance and economic efficiency of gas turbines depend largely on the efficiency of their thermodynamic cycle and the service life of their high-temperature components. Based on the thermodynamic cycle principles of gas turbines, increasing the turbine inlet temperature and reducing the amount of cooling air for turbine components are effective ways to improve gas turbine efficiency.

[0003] As a core component of a gas turbine, turbine blades account for a significant portion of the total cost of manufacturing. Traditional turbine blade manufacturing processes are complex, and achieving lightweight, high-performance cooling structures is difficult. In recent years, the rapid development of additive manufacturing technology has brought new opportunities for turbine blade manufacturing. This technology enables the precise manufacture of complex structures, paving the way for the development of novel cooling structures.

[0004] In the design of heavy-duty gas turbines, the assembly structure of the turbine stator blades is particularly important. The turbine stator blades are installed circumferentially around the turbine main shaft. The assembly structure between adjacent blades is the key to isolating the gas from the cold air. A reasonable assembly structure can effectively ensure the safe operation of the gas while improving the overall performance and reliability of the gas turbine.

[0005] Existing patent CN119754868A discloses a turbine stator blade, a turbine stator ring, and a gas turbine. The blade body is connected to an edge plate and arranged along the height direction of the blade body. The edge plate has periodic surfaces that are opposite to each other along the thickness direction of the blade body. The periodic surfaces are provided with a sealing groove. The inner cavity of the sealing groove includes a sealing section and a cooling section arranged along the depth direction of the sealing groove. The sealing section is suitable for mating with a sealing sheet. The side of the edge plate facing away from the blade body is provided with a first cooling hole connected to the cooling section, and the periodic surface is also provided with a second cooling hole connected to the cooling section. The combination of the sealing groove and the cooling unit in this solution results in insufficient reliability of the sealing structure, and gas leaks through the sealing groove, posing a safety hazard.

[0006] Existing patent CN113153447A discloses a pre-swirl structure for enhancing the cooling of leakage flow from a turbine stator endwall. The structure comprises a turbine stator, a stator endwall, a disc cavity gap upstream of the stator endwall, and a pre-swirl structure. The pre-swirl structure is a contraction channel formed by shrinking the turbine stator several times and then mounting it in reverse on the endwall in front of the cascade channel inlet. As the leakage flow flows from the stator endwall disc cavity gap to the cascade channel, it first generates a velocity component in the pre-swirl structure that is opposite to the direction of the endwall's lateral pressure gradient. This component is then accelerated to form a leakage flow with greater momentum directed toward the pressure side of the endwall channel, thereby facilitating the expansion of the leakage flow's cooling coverage of the endwall. The introduction of the pre-swirl structure in this patent alters the aerodynamic characteristics of the cascade channel, resulting in additional aerodynamic losses and aerodynamic instability in localized areas.

[0007] In summary, none of the aforementioned patents fully address the issues of safe and reliable cooling and sealing of the circumferential assembly structure of turbine vanes in the prior art. Therefore, developing cooling structures for turbine vanes suitable for additive manufacturing is of great significance for improving the performance and economic efficiency of heavy-duty gas turbines. Summary of the Invention

[0008] To better address the sealing and cooling safety issues of the circumferential assembly structure of turbine vanes in the prior art, this invention, based on additive manufacturing technology, proposes a turbine vane with a circumferential assembly cooling structure. This improves the thermal cycle efficiency of the gas turbine, extends the service life of high-temperature components, reduces manufacturing costs, and enhances the overall performance and market competitiveness of the gas turbine. The specific technical solution is as follows:

[0009] A turbine stator blade provided with a circumferentially assembled cooling structure comprises a blade and an edge plate connected to the blade top and blade bottom of the blade, the two ends of the edge plate along the circumference being circumferential assembly surfaces, a plurality of turbine stator blades are installed in sequence along the circumference of the turbine main shaft, sealing grooves and sealing members are provided between the circumferential assembly surfaces of adjacent edge plates, a composite cooling unit is provided on the edge plate near the circumferential assembly surface, the composite cooling unit connects the cold air channel outside the edge plate and the circumferential assembly surface, the composite cooling unit comprises a cooling cavity located inside the edge plate, the cooling cavity is located on the side of the sealing groove near the blade.

[0010] Furthermore, the composite cooling unit also includes an impact cavity and an air bleed channel located inside the edge plate. The impact cavity is located on the side of the cooling cavity away from the circumferential assembly surface. The impact cavity and the cooling cavity are connected through the air bleed channel.

[0011] Furthermore, the air bleed channel has a bending portion, one end of the air bleed channel is connected to an end of the cooling cavity away from the circumferential assembly surface, and the other end is connected to an end of the impact cavity away from the blade.

[0012] Furthermore, the composite cooling unit includes a plurality of air induction channels, and the plurality of air induction channels are distributed at equal intervals along the axial direction.

[0013] Furthermore, the composite cooling unit further comprises an impact hole, one end of which is connected to a side of the impact cavity away from the circumferential assembly surface, and the other end of which is connected to a side of the edge plate away from the blade.

[0014] Furthermore, the composite cooling unit includes a plurality of impact holes, and the plurality of impact holes are distributed at equal intervals along the axial direction.

[0015] Furthermore, the impact hole is arranged at an angle, and the inclination angle of the impact hole is 30°-60°.

[0016] Furthermore, the composite cooling unit further includes an exhaust hole, one end of which is connected to a side of the cooling cavity close to the circumferential assembly surface, and the other end of which is connected to the circumferential assembly surface.

[0017] Furthermore, the composite cooling unit includes a plurality of exhaust holes, and the plurality of exhaust holes are distributed at equal intervals along the axial direction.

[0018] Furthermore, the distance between the cooling cavity and the sealing groove is 0.5-2 times the radial width of the sealing groove, and the axial width of the cooling cavity is the same as that of the sealing groove.

[0019] Furthermore, a plurality of flow-disturbing structures are provided in the cooling cavity, and the flow-disturbing structures are distributed in an array.

[0020] Furthermore, the spoiler structure is a spoiler column, which radially connects two surfaces of the cooling cavity. The spoiler columns are arranged in multiple rows along the circumferential direction, and the spoiler columns in two adjacent rows are staggered.

[0021] Furthermore, the composite cooling unit is arranged on a single side of the circumferential assembly of the adjacent edge plates.

[0022] Furthermore, the composite cooling unit is mirror-imaged and arranged on both sides of the circumferential assembly of adjacent edge plates.

[0023] Based on the above technical solution, the present invention has at least the following beneficial effects:

[0024] 1. The present invention proposes a turbine stator blade with a circumferentially assembled cooling structure. The composite cooling unit combines an impingement cavity and a cooling cavity. The cold air utilization rate is higher than that of the traditional structure, and the cold air volume is lower than that of the traditional structure, thereby improving the cycle efficiency of the combustion turbine.

[0025] 2. The present invention proposes a turbine stator blade provided with a circumferentially assembled cooling structure. A composite cooling unit is provided at the circumferential assembly surface of the edge plates of adjacent turbine stator blades and is provided separately from the sealing groove. This does not affect the sealing effect of the sealing groove while protecting the sealing groove and the sealing component.

[0026] 3. The present invention proposes a turbine stator blade provided with a circumferentially assembled cooling structure. The composite cooling unit can be arranged on one side or in a mirrored manner on both sides of the circumferential assembly surface of the edge plate of the adjacent turbine stator blade to ensure safe and reliable operation at the circumferential sealing edge of the turbine stator blade.

[0027] 4. The present invention proposes a turbine stator blade with a circumferentially assembled cooling structure. Based on additive manufacturing technology, it ensures the structural dimensional accuracy of the composite cooling unit, reduces manufacturing costs, and improves the economic efficiency of gas turbine operation. BRIEF DESCRIPTION OF THE DRAWINGS

[0028] The accompanying drawings, which constitute part of the present invention, are intended to provide a further understanding of the present invention. The exemplary embodiments of the present invention and their descriptions are intended to explain the present invention and do not constitute an undue limitation of the present invention. In the accompanying drawings:

[0029] Figure 1 The present invention provides a schematic structural diagram of a turbine stator blade with a circumferentially assembled cooling structure;

[0030] Figure 2 The present invention provides a schematic diagram of a main shaft installation structure of a turbine stator blade with a circumferential assembly cooling structure;

[0031] Figure 3 The present invention provides a schematic radial cross-sectional view of a circumferential mounting surface of a turbine stator blade provided with a circumferential mounting cooling structure;

[0032] Figure 4 The present invention provides a three-dimensional schematic diagram of a circumferential mounting surface of a turbine stator blade provided with a circumferential mounting cooling structure;

[0033] Figure 5 The present invention provides a three-dimensional schematic diagram of a composite cooling unit for turbine stator blades provided with a circumferentially assembled cooling structure.

[0034] Figure numerals: 10-blade, 20-upper edge plate, 30-lower edge plate, 40-compound cooling unit, 11-blade top, 12-blade bottom, 13-suction surface, 14-pressure surface, 41-impact hole, 42-impact cavity, 43-bleed air channel, 44-cooling cavity, 45-exhaust hole. DETAILED DESCRIPTION

[0035] The present invention is further described in detail below with reference to specific examples. These examples should not be construed as limiting the scope of protection claimed in the present invention.

[0036] It should be noted that, in the absence of conflict, the embodiments and features of the embodiments of the present invention can be combined with each other. The present invention will be described in detail below with reference to the accompanying drawings and in combination with the embodiments.

[0037] The present invention proposes a turbine stator blade provided with a circumferentially assembled cooling structure based on additive manufacturing technology, which improves the cooling performance of the turbine stator blade, reduces the manufacturing cost, improves the cycle efficiency of the gas turbine, and improves the reliability of the gas turbine operation.

[0038] To achieve the above purpose, see Figure 1-Figure 5 The present invention provides a specific embodiment of a turbine static blade with a circumferentially assembled cooling structure.

[0039] See Figure 1 As shown, the turbine stator blade includes a blade 10, an upper edge plate 20, and a lower edge plate 30. The blade 1 has a blade top 11 and a blade bottom 12 that are opposite to each other in the height direction, and the blade body 10 has a suction surface 13 and a pressure surface 14 that are opposite to each other in the thickness direction. The upper edge plate 20 is connected to the blade top 11, and the two ends of the upper edge plate 20 along the circumferential direction are opposite circumferential assembly surfaces. The lower edge plate 30 is connected to the blade bottom 12, and the two ends of the lower edge plate 30 along the circumferential direction are opposite circumferential assembly surfaces. Axial sealing grooves are provided on the circumferential assembly surfaces of the upper edge plate 20 and the lower edge plate 30. A first seal is provided between the upper plate edges 20 of adjacent turbine stator blades, and the two ends of the first seal are disposed in the sealing grooves of the circumferential assembly surfaces of adjacent upper plate edges 20. A second seal is provided between the lower plate edges 30 of adjacent turbine stator blades, and the two ends of the second seal are disposed in the sealing grooves of the circumferential assembly surfaces of adjacent lower plate edges 30. The circumferential assembly surfaces of the upper edge plate 20 and the lower edge plate 30 are further provided with a composite cooling unit 40. Figure 2 As shown, a plurality of turbine blades are sequentially installed along the circumference of the turbine main shaft. The area between the upper edge plate 20 and the lower edge plate 30 is the gas side, and the outer sides of the upper edge plate 20 and the lower edge plate 30 are the cold air side.

[0040] See Figure 3 As shown, a composite cooling unit 40 is provided at the circumferential assembly surface of the upper plate edge 20 and the lower edge plate 30 (not shown) of adjacent turbine stator blades. The composite cooling unit 40 can be provided on a single side of the adjacent circumferential assembly surface, or can be arranged in a mirror image along the circumferential assembly surface, that is, adjacent upper edge plates 20 all include composite cooling units 40 symmetrically along the sealing groove.

[0041] The following mainly uses the upper edge plate 20 as an example for specific description. The lower edge plate 30 is the same as the upper edge plate 20 .

[0042] Specifically, the composite cooling unit 40 includes an impact hole 41, an impact cavity 42, an air bleed channel 43, a cooling cavity 44 and an exhaust hole 45. The impact hole 41 is arranged on the side of the upper edge plate 20 away from the blade 10 and is connected to the cold air side. The exhaust hole 45 is arranged on the circumferential assembly surface between adjacent upper edge plates 20. The impact cavity 42, the air bleed channel 43, and the cooling cavity 44 are arranged inside the upper edge plate 20. The impact hole 41, the impact cavity 42, the air bleed channel 43, the cooling cavity 44 and the exhaust hole 45 are connected in sequence.

[0043] See Figure 4 As shown, the cooling chamber 44 is opened on the side of the sealing groove close to the blade 10, and is spaced a certain distance from the sealing groove and parallel to the sealing groove; the impact chamber 42 is opened on the side of the cooling chamber 44 away from the circumferential assembly surface, and is spaced a certain distance from the cooling chamber 44; a plurality of exhaust holes 45 are arranged at equal intervals along the axial direction on the side of the cooling chamber 44 close to the circumferential assembly surface, and the exhaust holes 45 connect the cooling chamber 44 and the circumferential assembly surface. A plurality of array-arranged turbulent structures are arranged in the cooling chamber 44. The cooling chamber 44 and the impact chamber 42 are connected by a plurality of bleed air channels 43. One end of the bleed air channel 43 is connected to the side of the cooling chamber 44 close to the impact chamber 42, and the other end is connected to the side of the impact chamber 42 away from the blade 10 or to the upper end of the side of the impact chamber 42 close to the cooling chamber 44. The purpose is to fully utilize the cooling effect of the impact vortex on the gas side of the impact chamber 42, and at the same time effectively prevent the high-temperature gas from invading the cold gas side. Multiple impact holes 41 are arranged at equal intervals along the axial direction on the side of the impact cavity 42 away from the cooling cavity 44. The impact holes 41 are arranged at an angle. The inlet of the impact hole 41 is connected to the cold air side, and the outlet of the impact hole 42 extends to the interior of the upper edge plate 20 and is connected to the lower end of the impact cavity 42 away from the cooling cavity 44.

[0044] Optionally, the spacing between the cooling cavity 44 and the sealing groove is 0.5-2 times the radial width of the sealing groove. The cooling cavity 44 is a rectangular cavity, the axial width of which is the same as that of the sealing groove, and the circumferential width of which is greater than the circumferential depth of the sealing groove.

[0045] Optionally, the spoiler structure is a spoiler column, which is connected to both ends of the cooling cavity 44 along the axial direction. The spoiler columns are arranged in 3-5 rows along the circumferential direction, and the spoiler columns in two adjacent rows are staggered.

[0046] Optionally, the impact cavity 43 is a rectangular cavity, and its axial width is the same as that of the cooling cavity 44 .

[0047] Optionally, the inclination angle of the impact hole 41 is 30°-60°, ensuring that the surface of the impact cavity 42 close to the blade 10 obtains a larger heat transfer coefficient, thereby improving the cooling effect of the composite cooling unit.

[0048] See Figure 5As shown, the impact hole 41 is a circular or square hole with a diameter-to-length ratio greater than 1. Multiple air bleed channels 43 are equidistantly spaced axially. These channels are rectangular slits, with the cross-sectional area of each rectangular slit channel being the product of its radial width and axial length. This cross-sectional area is 1.0-1.2 times the area of the corresponding impact hole 41. Multiple exhaust holes 45 are equidistantly spaced axially on the circumferential assembly surface of the upper edge plate 20. These exhaust holes are square or circular, with a length-to-diameter ratio greater than 1. The ratio of the radial cross-sectional area of the exhaust hole 45 to the radial cross-sectional area of the corresponding impact hole 41 is 0.9-1.1.

[0049] When using the above-mentioned turbine stator blade with a circumferentially assembled cooling structure, cold air enters the impact cavity 42 from the cold air side along the impact hole 41 for impact cooling. The impact hole 41 ensures that the cold air is evenly distributed and impacts the impact cavity 42 at high speed, effectively enhancing the heat exchange efficiency; the cold air enters the downstream cooling cavity 44 through the air inlet channel 43. The design of the air inlet channel 43 optimizes the cold air flow path, further improving the cooling effect; the turbulent flow structure in the cooling cavity 44 can increase the contact area and time between the cold air and the wall, thereby improving the heat exchange efficiency; finally, it flows out from the gap between the circumferential assembly surfaces through the exhaust hole 45. The turbine stator blade with a circumferentially assembled cooling structure provided by the present invention does not affect the sealing effect of the blade, while achieving an efficient cooling effect, ensuring the safe and reliable operation of the circumferential sealing edge of the turbine stator blade.

[0050] The foregoing description is merely a preferred embodiment of the present invention and is not intended to limit the present invention. Those skilled in the art will readily appreciate that various modifications and variations of the present invention are possible. Any modifications, equivalent substitutions, or improvements made within the spirit and principles of the present invention are intended to be within the scope of protection of the present invention.

[0051] It should be noted that, in this document, relational terms such as first and second, etc., are used only to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any actual relationship or order between these entities or operations. Moreover, the terms "comprises," "comprising," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or device comprising a series of elements includes not only those elements, but also other elements not explicitly listed, or elements inherent to such process, method, article, or device. In the absence of further limitations, an element defined by the phrase "comprising a ..." does not exclude the presence of other identical elements in the process, method, article, or device comprising the element.

[0052] It should be noted that, in the description of this specification, the reference terms "one embodiment", "some embodiments", "example", "specific example", or "some examples" mean that the specific features, structures, materials or characteristics described in conjunction with the embodiment or example are included in at least one embodiment or example of the present invention. In this specification, the schematic expressions of the above terms do not necessarily refer to the same embodiment or example. Moreover, the specific features, structures, materials or characteristics described can be combined in any one or more embodiments or examples in a suitable manner. In addition, those skilled in the art can combine and combine different embodiments or examples described in this specification and the features of different embodiments or examples, unless they are contradictory.

Claims

1. A turbine stator blade with a circumferentially mounted cooling structure, comprising a blade and an edge plate connected to the blade tip and blade foot, wherein both ends of the edge plate along the circumferential direction form circumferential mounting surfaces. A plurality of turbine stator blades are sequentially mounted along the circumference of a turbine main shaft, and sealing grooves and sealing members are provided between the circumferential mounting surfaces of adjacent edge plates, characterized in that: A composite cooling unit is provided on the edge plate near the circumferential assembly surface, and the composite cooling unit connects the cold air channel on the outside of the edge plate and the circumferential assembly surface. The composite cooling unit includes a cooling cavity located inside the edge plate, and the cooling cavity is located on the side of the sealing groove close to the blade.

2. The turbine vane with a circumferentially mounted cooling structure according to claim 1, characterized in that: The composite cooling unit further includes an impact cavity and an air bleed channel located inside the edge plate. The impact cavity is located on a side of the cooling cavity away from the circumferential assembly surface. The impact cavity and the cooling cavity are connected via the air bleed channel.

3. The turbine vane with a circumferentially mounted cooling structure according to claim 2, characterized in that: The air bleed channel has a bent portion, one end of the air bleed channel is connected to an end of the cooling cavity away from the circumferential assembly surface, and the other end is connected to an end of the impact cavity away from the blade.

4. The turbine vane with a circumferentially mounted cooling structure according to claim 3, characterized in that: The composite cooling unit includes a plurality of air induction channels, and the plurality of air induction channels are distributed at equal intervals along the axial direction.

5. The turbine vane with a circumferentially mounted cooling structure according to claim 3, characterized in that: The composite cooling unit further comprises an impact hole, one end of which is connected to a side of the impact cavity away from the circumferential assembly surface, and the other end of which is connected to a side of the edge plate away from the blade.

6. The turbine vane with a circumferentially mounted cooling structure according to claim 5, characterized in that: The composite cooling unit includes a plurality of impact holes, and the plurality of impact holes are distributed at equal intervals along the axial direction.

7. The turbine vane with a circumferentially mounted cooling structure according to claim 6, characterized in that: The impact hole is arranged obliquely, and the inclination angle of the impact hole is 30°-60°.

8. The turbine vane with a circumferentially mounted cooling structure according to claim 5, characterized in that: The composite cooling unit further includes an exhaust hole, one end of which is connected to a side of the cooling cavity close to the circumferential assembly surface, and the other end of which is connected to the circumferential assembly surface.

9. The turbine vane with a circumferentially mounted cooling structure according to claim 8, characterized in that: The composite cooling unit includes a plurality of exhaust holes, and the plurality of exhaust holes are distributed at equal intervals along the axial direction.

10. The turbine vane with a circumferentially mounted cooling structure according to claim 1, characterized in that: The distance between the cooling cavity and the sealing groove is 0.5-2 times the radial width of the sealing groove, and the axial width of the cooling cavity is the same as that of the sealing groove.

11. The turbine vane with a circumferentially mounted cooling structure according to any one of claims 1 to 10, characterized in that: A plurality of flow-disturbing structures are provided in the cooling cavity, and the flow-disturbing structures are distributed in an array.

12. The turbine vane with a circumferentially mounted cooling structure according to claim 11, characterized in that: The spoiler structure is a spoiler column, which connects two surfaces of the cooling cavity in a radial direction. The spoiler columns are arranged in multiple rows along the circumferential direction, and the spoiler columns in two adjacent rows are staggered.

13. The turbine vane with a circumferentially mounted cooling structure according to claim 11, characterized in that: The composite cooling unit is arranged on a single side of the circumferential assembly of the adjacent edge plates.

14. The turbine vane with a circumferentially mounted cooling structure according to claim 11, characterized in that: The composite cooling unit is mirror-imaged and arranged on both sides of the circumferential assembly of adjacent edge plates.

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