High-temperature turbine blade based on complex matrix rib double-layer wall cooling structure and gas turbine

Abstract

The invention relates to a high-temperature turbine blade based on a complex matrix rib double-wall cooling structure and a gas turbine, the high-temperature turbine blade comprises a blade body which is composed of an outer thin-wall layer and an inner solid layer, wedge-shaped matrix ribs are arranged in the inner solid layer from the front edge to the middle chord area of the inner solid layer, and an exhaust groove penetrating through the outer thin-wall layer is formed in the tail edge area of the inner solid layer; the exhaust grooves are communicated with the rear end surfaces of the wedge-shaped matrix ribs; impact holes which impact the inner wall face of the inner solid layer through cooling gas jet flow of the wedge-shaped matrix ribs are formed in the inner solid layer in an arrayed mode, and gas film holes which form a cooling gas film on the outer surface of the outer thin wall layer are formed in the outer thin wall in an arrayed mode. According to the turbine blade, the wedge-shaped matrix ribs and the exhaust grooves are arranged in the turbine blade, so that the cooling effect on the interior of the blade is achieved, jet flow generated by the impact holes impacts on the inner wall face of the inner solid layer, the air film holes form a cooling air film on the outer surface of the outer thin-wall layer, the heat exchange capacity in the turbine blade is better improved, and the service life of the turbine blade is prolonged. And the cooling effect of the turbine blade is improved.

Description

Technical Field

[0001] The present invention relates to the field of gas turbine turbine technology, in particular to a high-temperature turbine blade and a gas turbine based on a complex matrix rib double-wall cooling structure. Background Art

[0002] Increasing the turbine inlet temperature is the key to improving the output power and thermal efficiency of modern gas turbines. Today, the ambient temperature of turbine blades has exceeded 2200K, far exceeding the melting point of blade materials. Therefore, it is necessary to design a new internal cooling structure to protect the blades.

[0003] Currently, the cooling structure inside the blade is widely implemented using pin-fin arrays, ribbed channels or column ribs, but in actual use, the enhanced heat exchange capacity of these cooling structures is still limited and cannot meet the operating conditions of advanced gas turbines with higher heat loads. The matrix rib is a cooling structure with extremely high enhanced heat transfer capacity and can provide higher structural strength for the blade.

[0004] The double-wall cooling structure is a composite cooling method that integrates internal impingement cooling and external air film cooling, that is, the cold air jets from the impingement holes on the inner wall to the inner surface of the outer wall, and then flows out from the air film holes on the outer surface to form air film cooling. This cooling method gives full play to the internal impingement jet to enhance heat exchange and the air film hole overflow to enhance heat exchange. It is an effective method to improve the local heat transfer coefficient. Under the same opening area and hot and cold airflow conditions, it is more efficient than traditional divergent cooling.

[0005] However, since the leading edge of the blade is subject to an extremely high heat load and the trailing edge is extremely thin and easily causes thermal failure, it is necessary to design different cooling structures in different areas. Summary of the invention

[0006] The technical problem to be solved by the present invention is: in order to overcome the deficiencies in the prior art, the present invention provides a high-temperature turbine blade based on a complex matrix rib double-wall cooling structure to further improve the overall cooling efficiency of the turbine blade and maintain good convective heat transfer uniformity; and provides a gas turbine having the high-temperature turbine blade.

[0007] The technical solution adopted by the present invention to solve its technical problems is: a high-temperature turbine blade based on a complex matrix rib double-wall cooling structure, including a blade body, the outer convex arc surface on the rear side of the blade body is the suction surface, and the inner concave arc surface on the front side of the blade body is the pressure surface, the blade body is composed of an outer thin-walled layer and an inner solid layer located in the outer thin-walled layer, and wedge-shaped matrix ribs are arranged in the inner solid layer from its leading edge to the mid-chord area, and the front and rear side surfaces of the wedge-shaped matrix ribs are respectively corresponding to the front and rear inner side surfaces of the inner solid layer, and the trailing edge area of ​​the inner solid layer is provided with an exhaust groove that penetrates the outer thin-walled layer, and the exhaust groove is connected to the rear end surface of the wedge-shaped matrix rib; the inner solid layer is provided with impact holes for the cooling gas jet passing through the wedge-shaped matrix ribs to impact the inner wall surface of the inner solid layer, and the outer thin wall is provided with air film holes for forming a cooling air film on the outer surface of the outer thin-walled layer.

[0008] The wedge-shaped matrix ribs are composed of two rows of ribs, front and rear. The front and rear rows of ribs are staggered at 45°, and the wedge angle of the wedge-shaped matrix ribs is 10°. In this way, a matrix channel can be formed between the front and rear rows of ribs to regulate the entry of cooling airflow, so that the cooling airflow can more quickly form a more complex vortex structure using the structural form of the front and rear rows of matrix ribs.

[0009] The ribs are provided with through holes to reduce the overall weight of the blades.

[0010] In order to better improve the cooling effect of the inner wall of the inner solid layer, the impact holes corresponding to the suction surface and the pressure surface are distributed in an array on the inner solid layer, and the center of the impact hole corresponds to the center of the square formed by the staggered front and rear rows of ribs; the impact holes located at the front edge of the inner solid layer are distributed in four rows at intervals, so as to utilize the jet impact force generated by the impact holes to form a stronger cooling force on the inner wall of the solid layer.

[0011] Preferably, the impact hole is a cylindrical hole, and the diameter of the impact hole is 1 mm.

[0012] The air film holes are interconnected with the impact holes on the suction surface and the pressure surface, and the air film holes corresponding to the leading edge of the inner solid layer are interlaced with the impact holes, so that the cooling gas inside the outer thin wall is guided to the outer wall of the outer thin wall by the air film holes to form a gas protective film, thereby achieving good heat insulation and airflow cooling effects.

[0013] Preferably, the air film hole is a cylindrical hole with a diameter of 1 mm.

[0014] A gas turbine has the high-temperature turbine blades described above.

[0015] The beneficial effects of the present invention are as follows: the present invention achieves cooling effect on the inside of the turbine blade by arranging wedge-shaped matrix ribs and exhaust grooves inside the turbine blade, and utilizes the jet generated by the impact hole to impact the inner wall surface of the inner solid layer and the air film hole to form a cooling air film on the outer surface of the outer thin wall layer, thereby better improving the heat exchange capacity inside the turbine blade and improving the cooling effect of the turbine blade. BRIEF DESCRIPTION OF THE DRAWINGS

[0016] The present invention is further described below in conjunction with the accompanying drawings and embodiments.

[0017] Figure 1 It is a schematic diagram of the three-dimensional structure of the present invention.

[0018] Figure 2 It is a schematic diagram of the top structure of the present invention.

[0019] Figure 3 It is a schematic diagram of the three-dimensional structure after the outer thin wall layer is removed according to the present invention.

[0020] Figure 4 It is a schematic diagram of the three-dimensional structure of the wedge-shaped matrix ribs of the present invention.

[0021] Figure 5 It is a schematic diagram of the main structure of the wedge-shaped matrix ribs of the present invention.

[0022] In the figure: 1. blade body, 2. suction surface, 3. pressure surface, 4. outer thin wall layer, 5. inner solid layer, 6. wedge-shaped matrix ribs, 7. exhaust grooves, 8. impact holes, 9. air film holes, 10. through holes. DETAILED DESCRIPTION

[0023] The present invention will now be described in further detail with reference to the accompanying drawings. These drawings are simplified schematic diagrams, which only illustrate the basic structure of the present invention in a schematic manner, and therefore only show the components related to the present invention.

[0024] like Figure 1 to Figure 5 As shown, a high-temperature turbine blade based on a complex matrix rib double-layer wall cooling structure includes a blade body 1, which is composed of an outer thin-wall layer 4 and an inner solid layer 5 located inside the outer thin-wall layer 4. The convex arc surface on the rear side of the blade body 1 forms a suction surface 2 when the turbine blade is working, and the concave arc surface on the front side of the blade body 1 forms a pressure surface 3 when the turbine blade is working.

[0025] A cavity extending from the leading edge of the inner solid layer 5 to the mid-chord area is opened in the inner solid layer 5, and a wedge-shaped matrix rib 6 is installed in the cavity. An exhaust groove 7 is opened in the trailing edge area of ​​the inner solid layer 5. The inner end of the exhaust groove 7 is communicated with the rear end surface of the wedge-shaped matrix rib 6, and the outer end of the exhaust groove 7 penetrates the outer thin-walled layer 4 on one side of the pressure surface 3, and there is a certain inclination angle between the exhaust groove 7 and the pressure surface 3. After the cooling gas passes through the wedge-shaped matrix rib 6, it is smoothly discharged outward from the exhaust groove 7, thereby improving the internal cooling effect of the trailing edge area of ​​the blade and enhancing the structural strength of the trailing edge area.

[0026] The wedge-shaped matrix ribs 6 are composed of two rows of ribs, front and rear. The front row of ribs and the rear row of ribs are staggered at 45° to form a staggered grid flow channel. The wedge angle of the wedge-shaped matrix ribs 6 is 10°. The front side surface of the wedge-shaped matrix ribs 6 corresponds to the front inner side surface of the inner solid layer 5, and the rear side surface of the wedge-shaped matrix ribs 6 corresponds to the rear side surface of the inner solid layer 5, so that the suction surface 2 and the pressure surface 3 are cooled.

[0027] Through holes 10 are provided on both the front row ribs and the rear row ribs to reduce the overall weight of the blades and achieve overall lightweighting of the turbine blades.

[0028] The inner wall of the inner solid layer 5 is provided with cylindrical impact holes 8, and the diameter of the impact holes 8 is 1 mm. The specific distribution of the impact holes 8 is as follows: the impact holes 8 corresponding to the suction surface 2 and the pressure surface 3 are distributed in an array, and the center of the impact hole 8 corresponds to the center of the square grid flow channel formed by the staggered front and rear rows of ribs; the impact holes 8 located at the front edge of the inner solid layer 5 are distributed in four rows at intervals, and each row of impact holes 8 has nine holes arranged from top to bottom. After the cooling gas passing through the wedge-shaped matrix ribs 6 passes through the impact holes 8, it is sprayed on the inner wall of the inner solid layer 5 in the form of a jet.

[0029] The outer thin wall 4 is provided with air film holes 9 for forming a cooling air film on the outer surface of the outer thin wall layer 4. The air film holes 9 are cylindrical holes with a diameter of 1 mm. The air film holes 9 are interconnected with the impact holes 8 on the suction surface 2 and the pressure surface 3, and the air film holes 9 and the impact holes 8 are interlaced with each other at the leading edge of the inner solid layer 5.

[0030] At the same time, the present invention provides a gas turbine having the high-temperature turbine blades described above.

[0031] The present invention optimizes the design of the turbine blade structure of the gas turbine. By arranging wedge-shaped matrix ribs 6 and exhaust grooves 7 inside the turbine blade, a comprehensive cooling effect on the inside of the blade is achieved. The jet generated by the impact hole 8 impacts the inner wall surface of the inner solid layer 5 and the air film hole 9 forms a cooling air film on the outer surface of the outer thin wall layer 4, thereby better improving the heat exchange capacity inside the turbine blade and improving the overall cooling effect of the turbine blade.

[0032] Based on the above ideal embodiments of the present invention, the relevant staff can make various changes and modifications without departing from the technical concept of the present invention through the above description. The technical scope of the present invention is not limited to the contents of the specification, and its technical scope must be determined according to the scope of the claims.

Claims

1. A high-temperature turbine blade based on a complex matrix rib double-wall cooling structure, comprising a blade body (1), wherein the outer convex arc surface on the rear side of the blade body (1) is a suction surface (2), and the inner concave arc surface on the front side of the blade body (1) is a pressure surface (3), wherein the blade body (1) is composed of an outer thin wall layer (4) and an inner solid layer (5) located inside the outer thin wall layer (4), and wherein: The inner solid layer (5) is provided with wedge-shaped matrix ribs (6) from its leading edge to the mid-chord region, and the front and rear side surfaces of the wedge-shaped matrix ribs (6) are respectively corresponding to the front and rear inner side surfaces of the inner solid layer (5), and the trailing edge region of the inner solid layer (5) is provided with an exhaust groove (7) penetrating the outer thin-walled layer (4), and the exhaust groove (7) is communicated with the rear end surface of the wedge-shaped matrix ribs (6); the inner solid layer (5) is provided with impact holes (8) for the cooling gas jet passing through the wedge-shaped matrix ribs (6) to impact the inner wall surface of the inner solid layer (5), and the outer thin wall (4) is provided with air film holes (9) for forming a cooling air film on the outer surface of the outer thin-walled layer (4).

2. The high temperature turbine blade based on the complex matrix rib double wall cooling structure according to claim 1 is characterized in that: The wedge-shaped matrix ribs (6) are composed of two rows of ribs, front and rear, the front row of ribs and the rear row of ribs are staggered at 45 degrees, and the wedge angle of the wedge-shaped matrix ribs (6) is 10 degrees.

3. The high temperature turbine blade based on the complex matrix rib double wall cooling structure according to claim 2 is characterized in that: The ribs are provided with through holes (10) for reducing the weight of the blades.

4. The high temperature turbine blade based on the complex matrix rib double wall cooling structure according to claim 2, characterized in that: The impact holes (8) corresponding to the suction surface (2) and the pressure surface (3) are distributed in an array on the inner solid layer (5), and the center of the impact hole (8) corresponds to the center of the square formed by the front and rear rows of ribs; the impact holes (8) located at the front edge end of the inner solid layer (5) are distributed in four rows at intervals.

5. The high temperature turbine blade based on the complex matrix rib double wall cooling structure according to claim 4, characterized in that: The impact hole (8) is a cylindrical hole, and the diameter of the impact hole (8) is 1 mm.

6. The high temperature turbine blade based on the complex matrix rib double wall cooling structure according to claim 4, characterized in that: The air film holes (9) are interlacedly connected with the impact holes (8) on the suction surface (2) and the pressure surface (3), and the air film holes (9) and the impact holes (8) corresponding to the front edge of the inner solid layer (5) are interlaced with each other.

7. The high temperature turbine blade based on the complex matrix rib double wall cooling structure according to claim 6, characterized in that: The air film hole (9) is a cylindrical hole, and the diameter of the air film hole (9) is 1 mm.

8. A gas turbine, characterized in that: A high-temperature turbine blade according to any one of claims 1 to 7.

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