Gas turbine blade based on partitioned composite cooling structure

Abstract

According to the gas turbine blade based on the partitioned composite cooling structure, a plurality of different cooling channels are arranged in a blade body in a partitioned mode, and the cooling channels mainly comprise double-layer S-shaped ribs in the front edge area of the blade, double-layer wedge-shaped matrix ribs in the middle chord area of the blade and double-channel exhaust grooves in the tail edge area of the blade; air film holes are longitudinally formed in the peripheral wall of the outer layer of the thin wall corresponding to the positions of the double-layer S-shaped ribs and the double-layer wedge-shaped matrix ribs; jet holes are longitudinally formed in the circumferential wall of the solid inner layer, and cooling gas passing through the double-layer S-shaped ribs and the double-layer wedge-shaped matrix ribs is jetted to the inner wall face of the thin-wall outer layer through the jet holes. The overall heat exchange area in the turbine blade is increased, turbulence is enhanced, and boundary layer disturbance of airflow is increased; cold air is introduced from the interior or the periphery of the blade through the air film holes and sprayed to the surface of the blade to form an air film, direct contact between high-temperature airflow and the surface of a blade material is isolated, and therefore the thermal load on the surface of the blade is reduced, and the cooling effect of cooling gas on the high-temperature blade is improved.

Description

Technical Field

[0001] The invention relates to the technical field of gas turbines, in particular to a gas turbine blade based on a zoned composite cooling structure. Background Art

[0002] The cooling structure design of gas turbine turbine blades is one of the key technologies to improve their cooling efficiency and life. Since turbine blades operate in a high temperature and high pressure working environment for a long time, it is necessary to further strengthen and improve the cooling effect of turbine blades.

[0003] From the structural analysis of the turbine blades, different areas face similar but not completely consistent high temperature effects. The leading edge of the blade is the area most severely impacted by the high-temperature and high-speed fluid. In this area, the blade needs to withstand great impact and ensure the structural integrity and working stability of the blade; the chord area of ​​the blade is the central part of the blade, connecting the leading edge and trailing edge areas of the blade, and carries important aerodynamic and heat transfer functions. Its structural design must be reasonable and smooth to ensure that the cooling airflow can be evenly distributed and effectively take away the heat; and the trailing edge area of ​​the blade is thinner and needs to have sufficient strength and rigidity to resist the shear force, deflection and vibration of the high-speed airflow on the blade. In general, the gas turbine blade needs to carry out targeted cooling structure optimization in different areas to improve the overall durability and performance of the blade.

[0004] The existing cooling methods for turbine blades include external cooling and internal cooling. The external cooling technology generally adopts film cooling, and the internal cooling technology adopts cooling channels set inside the blades, such as pin-fin arrays and ribbed channels. However, its enhanced heat exchange capacity is still limited and cannot meet the operating conditions of advanced gas engines with higher heat loads. Therefore, it is necessary to further transform and upgrade the blade cooling structure to meet the use requirements of turbine blades in high heat load environments. Summary of the invention

[0005] 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 gas turbine turbine blade based on a partitioned composite cooling structure with better cooling effect.

[0006] The technical solution adopted by the present invention to solve the technical problem is: a combustion engine turbine blade based on a partitioned composite cooling structure, having a blade body, the blade body comprising a thin-walled outer layer, an inner wall of the thin-walled outer layer being embedded with a solid inner layer, the solid inner layer having a rectangular cavity, a wedge-shaped cavity and an exhaust cavity in sequence from the leading edge to the trailing edge of the blade body, the rectangular cavity being provided with a double-layer S-shaped rib, the wedge-shaped cavity being provided with a double-layer wedge-shaped matrix rib, the exhaust cavity being provided with a double-channel exhaust groove formed by setting a partition, a cooling channel for cooling gas flow being formed between the rectangular cavity and the wedge-shaped cavity; the double-layer S-shaped rib is composed of two rows of curved front and rear rows. The double-layer wedge-shaped matrix ribs are composed of two rows of wedge-shaped ribs that are cross-staggered, and the thin-walled outer layer is provided with exhaust holes that are connected to the double-channel exhaust grooves; air 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 solid inner layer, 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.

[0007] Specifically, the convex arc surface on the rear side of the blade body is the suction surface, the concave arc surface on the front side of the blade body is the pressure surface, the front side surface of the double-layer wedge-shaped matrix ribs fits the pressure surface, and the rear side surface of the double-layer wedge-shaped matrix ribs fits the suction surface.

[0008] Preferably, the jet holes are elliptical holes, and the jet holes corresponding to the pressure surface and the suction surface are distributed in an array on the solid inner wall, and the jet holes on the solid inner wall at the leading edge of the blade body are distributed in three rows at intervals.

[0009] Preferably, the air film holes are regular hexagonal holes with a side length of 0.8 mm, and the air film holes are staggeredly connected with the jet holes in the arrangement direction on the pressure surface, the suction surface and the leading edge area of ​​the blade body.

[0010] Furthermore, the front and rear rows of S-shaped ribs of the double-layer S-shaped ribs are provided with interconnected diamond-shaped holes at the overlapping curved parts, so as to reduce the overall weight of the turbine blades; and the upper and lower planes of each wedge-shaped rib of the double-layer wedge-shaped matrix ribs are spaced apart with concave arc grooves, so as to increase the contact area between the wedge-shaped ribs and the cooling gas.

[0011] Furthermore, a connecting column is fixedly connected between the rear wall surface of the partition and the rear inner wall surface of the solid inner layer.

[0012] Preferably, the cross section of the cooling channel is an irregular trapezoid, and the width of the cooling channel gradually shrinks from the rectangular cavity to the wedge-shaped cavity.

[0013] Specifically, the exhaust hole includes a central hole with a diameter of 2 mm, and four small holes with a diameter of 1 mm are evenly distributed circumferentially on the contour line of the central hole. The central hole and the small holes form a plum blossom-shaped exhaust hole.

[0014] The beneficial effects of the present invention are as follows: the present invention increases the overall heat exchange area inside the turbine blade, enhances turbulence and increases boundary layer disturbance of the airflow by arranging double-layer S-shaped ribs, double-layer wedge-shaped matrix ribs and double-channel exhaust grooves in partitions inside the turbine blade; at the same time, the design of the air film holes can introduce cooler air from inside or around the blade, and spray it onto the blade surface to form an air film, isolating the direct contact between the high-temperature airflow and the surface of the blade material, thereby reducing the heat load on the blade surface, and the dual combination of internal cooling and external cooling realizes an all-round cooling effect on the turbine blade. BRIEF DESCRIPTION OF THE DRAWINGS

[0015] The present invention is further described below in conjunction with the accompanying drawings and embodiments.

[0016] Figure 1 It is a schematic diagram of the three-dimensional structure of the present invention.

[0017] Figure 2 It is a schematic diagram of the top structure of the present invention.

[0018] Figure 3 It is a schematic diagram of the three-dimensional structure of the thin-walled outer layer of the present invention.

[0019] Figure 4 It is a schematic diagram of the three-dimensional structure of the inner layer of the entity described in the present invention.

[0020] Figure 5 It is a schematic diagram of the three-dimensional structure of the double-layer S-shaped ribs of the present invention.

[0021] Figure 6 It is a schematic diagram of the three-dimensional structure of the double-layer wedge-shaped matrix ribs of the present invention.

[0022] Figure 7 It is a schematic diagram of the enlarged structure of the exhaust hole of the present invention.

[0023] In the figure: 1. blade body, 2. thin-walled outer layer, 3. solid inner layer, 3-1. rectangular cavity, 3-2. wedge-shaped cavity, 3-3. exhaust cavity, 4. double-layer S-shaped ribs, 5. double-layer wedge-shaped matrix ribs, 6. partition, 7. dual-channel exhaust grooves, 8. cooling channels, 9. cylindrical protrusions, 10. exhaust holes, 10-1. center holes, 10-2. small holes, 11. air film holes, 12. jet holes, 13. diamond holes, 14. concave arc grooves, 15. connecting columns, 16. suction surfaces, 17. pressure surfaces. DETAILED DESCRIPTION

[0024] 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.

[0025] like Figure 1 to Figure 4 A combustion engine turbine blade based on a partitioned composite cooling structure is shown, comprising a blade body 1, wherein the blade body 1 is composed of a thin-walled outer layer 2 and a solid inner layer 3, wherein the solid inner layer 3 is embedded in the thin-walled outer layer 2 and fits with the inner wall of the thin-walled outer layer 2.

[0026] The blade body 1 is divided into a leading edge area, a mid-chord area and a trailing edge area from front to back, the outer convex arc surface on the rear side of the blade body 1 is a suction surface 16, and the inner concave arc surface on the front side of the blade body 1 is a pressure surface 17. The solid inner layer 3 is provided with a rectangular cavity 3-1 corresponding to the leading edge area, a wedge-shaped cavity 3-2 corresponding to the mid-chord area and an exhaust cavity 3-3 corresponding to the trailing edge area, a double-layer S-shaped rib 4 is provided in the rectangular cavity 3-1, and a double-layer wedge-shaped matrix rib 5 is provided in the wedge-shaped cavity 3-2. The front side surface of the double-layer wedge-shaped matrix rib 5 is in contact with the pressure surface 17, and the rear side surface of the double-layer wedge-shaped matrix rib 5 is in contact with the suction surface 16; a partition 6 extending along the length direction of the blade body is provided in the middle position of the exhaust cavity 3-4, and a connecting column 15 is fixedly connected between the rear wall surface of the partition 6 and the rear inner wall surface of the solid inner layer 3, and the exhaust cavity 3-4 is divided into a double-channel exhaust groove 7 by the partition 6. A double-channel exhaust groove 7 is arranged in the trailing edge area of ​​the blade body 1, especially with the help of the turbulence effect of the connecting column 15 on the gas, which can effectively increase the effective heat exchange area of ​​the trailing edge area, weaken the anti-kidney vortex that affects the heat exchange effect, and enhance the cooling effect of the trailing edge area.

[0027] A cooling channel 8 for cooling gas flow is formed between the rectangular cavity 3-1 and the wedge-shaped cavity 3-2. The cross-section of the cooling channel 8 is an irregular trapezoid. The width of the cooling channel 8 gradually shrinks from the rectangular cavity 3-1 to the wedge-shaped cavity 3-2, which can play a better cooling gas diversion role.

[0028] like Figure 5 As shown, the double-layer S-shaped rib 4 is composed of two rows of S-shaped ribs with opposite bending directions, each row of S-shaped ribs is arranged at equal intervals along the height direction of the blade body 1, and the upper and lower surfaces of each S-shaped rib are densely covered with cylindrical protrusions 9. When the cooling gas passes through the cylindrical protrusions 9, the gas will produce a turbulent effect due to the irregular shape caused by the protrusions on the surface of the S-shaped rib, thereby increasing the kinetic energy, and changing the gas flow speed and direction to generate a pressure difference, thereby promoting the passage of the cooling gas.

[0029] The S-shaped ribs can significantly disturb the airflow. When the cooling gas passes through, a complex flow pattern will be generated, breaking the laminar state of the airflow and enhancing the heat exchange. In particular, the design of the double-layer S-shaped ribs 4, which further increases the formation of vortices through the arrangement of two layers of S-shaped ribs in different directions, helps to improve the heat transfer coefficient of the cooling gas and take away more heat.

[0030] The front and rear rows of the double-layer S-shaped ribs 4 are provided with interconnected diamond-shaped holes 13 at their respective curved overlapping positions, so as to reduce the overall weight of the turbine blade.

[0031] like Figure 6 As shown, the double-layer wedge-shaped matrix rib 5 is composed of two rows of wedge-shaped ribs that are cross-staggered in front and back. Inwardly concave arc grooves 14 are arranged at intervals on the upper and lower planes of each wedge-shaped rib of the double-layer wedge-shaped matrix rib 5. The inwardly concave arc grooves 14 generate local disturbances on the cooling gas, reduce the thickness of the boundary layer, and thus help to improve the cooling effect of the cooling gas. At the same time, the inwardly concave arc grooves 14 can increase the effective surface area of ​​the wedge-shaped rib in contact with the cooling gas, thereby providing more heat exchange surface area.

[0032] The double-layer wedge-shaped matrix ribs 5 can increase the surface area of ​​the cooling channel inside the blade, disturb the airflow and enhance heat exchange, significantly improving the cooling efficiency and high temperature resistance of the turbine blade, thereby enhancing the overall performance and reliability of the turbine blade, helping to improve the efficiency of the gas turbine, extend its service life, and improve the stability of the power system.

[0033] like Figure 4 As shown, jet holes 12 are longitudinally arranged on the peripheral wall of the solid inner layer 3, and the jet holes 12 are elliptical holes. The jet holes 12 corresponding to the pressure surface 17 and the suction surface 16 are distributed in an array on the wall of the solid inner layer 3, and the jet holes 12 located on the wall of the solid inner layer 3 at the leading edge of the blade body 1 are distributed in three rows at intervals.

[0034] like Figure 3 As shown, air film holes 11 are longitudinally arranged on the peripheral wall of the thin-walled outer layer 2 corresponding to the positions of the double-layer S-shaped ribs 4 and the double-layer wedge-shaped matrix ribs 5. The air film holes 11 are regular hexagonal holes with a side length of 0.8 mm. The air film holes 11 are staggered and connected with the jet holes 12 in the arrangement direction on the pressure surface 17, the suction surface 16 and the leading edge area of ​​the blade body 1. The cooling gas passing through the double-layer S-shaped ribs 4 and the double-layer wedge-shaped matrix ribs 5 is sprayed onto the inner wall surface of the thin-walled outer layer 2 through the jet holes 12.

[0035] The thin-walled outer layer 2 is provided with exhaust holes 10 connected to the dual-channel exhaust grooves 7; Figure 7As shown, the exhaust hole 10 includes a central hole 10-1 with a diameter of 2 mm, and four small holes 10-2 with a diameter of 1 mm are evenly distributed circumferentially on the contour line of the central hole 10-1. The central hole 10-1 and the small holes 10-2 form a plum blossom-shaped exhaust hole 10.

[0036] The above-mentioned air film holes 11, jet holes 12 and exhaust holes 10 all adopt a special-shaped hole structure, so as to optimize the penetration and heat dissipation path of the cooling gas, so that the turbine blades can better cope with high temperature loads and improve high temperature resistance; on the other hand, the design structure of the above-mentioned special-shaped holes can flexibly adapt to the space requirements of different areas, and through the reasonable arrangement of holes of different shapes, it can provide the best cooling effect in a limited space.

[0037] The present invention increases the overall heat exchange area inside the turbine blade, enhances turbulence and increases boundary layer disturbance of the airflow by arranging double-layer S-shaped ribs 4, double-layer wedge-shaped matrix ribs 5 and double-channel exhaust grooves 7 in partitions inside the turbine blade. At the same time, the design of the air film holes 11 can introduce cooler air from inside or around the blade and spray it onto the blade surface to form an air film, isolating the direct contact between the high-temperature airflow and the surface of the blade material, thereby reducing the heat load on the blade surface. The dual combination of internal cooling and external cooling realizes an all-round cooling effect on the turbine blade.

[0038] 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 combustion engine turbine blade based on a zoned composite cooling structure, comprising a blade body (1), wherein the blade body (1) comprises a thin-walled outer layer (2), and a solid inner layer (3) is embedded in the inner wall of the thin-walled outer layer (2), wherein: The solid inner layer (3) has a rectangular cavity (3-1), a wedge-shaped cavity (3-2) and an exhaust cavity (3-3) in sequence from the leading edge to the trailing edge of the blade body (1); the rectangular cavity (3-1) is provided with a double-layer S-shaped rib (4); the wedge-shaped cavity (3-2) is provided with a double-layer wedge-shaped matrix rib (5); the exhaust cavity (3-4) has a double-channel exhaust groove (7) formed by arranging a partition (6); a cooling channel (8) for cooling gas to flow is formed between the rectangular cavity (3-1) and the wedge-shaped cavity (3-2); The double-layer S-shaped ribs (4) are composed of two front and rear rows of S-shaped ribs with opposite bending directions, each row of S-shaped ribs is arranged at equal intervals along the height direction of the blade body (1), and the upper and lower surfaces of each S-shaped rib are densely covered with cylindrical protrusions (9); the double-layer wedge-shaped matrix ribs (5) are composed of two front and rear rows of wedge-shaped ribs that are cross-staggered, and the thin-walled outer layer (2) is provided with exhaust holes (10) that are connected to the dual-channel exhaust grooves (7); Air film holes (11) are arranged longitudinally on the peripheral wall of the thin-walled outer layer (2) corresponding to the positions of the double-layer S-shaped ribs (4) and the double-layer wedge-shaped matrix ribs (5); jet holes (12) are arranged longitudinally on the peripheral wall of the solid inner layer (3), and the cooling gas passing through the double-layer S-shaped ribs (4) and the double-layer wedge-shaped matrix ribs (5) is sprayed onto the inner wall surface of the thin-walled outer layer (2) through the jet holes (12).

2. The combustion engine turbine blade based on the zoned composite cooling structure according to claim 1 is characterized in that: The convex arc surface on the rear side of the blade body (1) is the suction surface (16), the concave arc surface on the front side of the blade body (1) is the pressure surface (17), the front side surface of the double-layer wedge-shaped matrix rib (5) is in contact with the pressure surface (17), and the rear side surface of the double-layer wedge-shaped matrix rib (5) is in contact with the suction surface (16).

3. The combustion engine turbine blade based on the zoned composite cooling structure according to claim 2 is characterized in that: The jet holes (12) are elliptical holes. The jet holes (12) corresponding to the pressure surface (17) and the suction surface (16) are distributed in an array on the wall of the solid inner layer (3). The jet holes (12) located on the wall of the solid inner layer (3) at the leading edge of the blade body (1) are distributed in three rows at intervals.

4. The combustion engine turbine blade based on the zoned composite cooling structure according to claim 3 is characterized in that: The air film holes (11) are regular hexagonal holes with a side length of 0.8 mm. The air film holes (11) are staggeredly connected with the jet holes (12) in the arrangement direction on the pressure surface (17), the suction surface (16) and the leading edge area of ​​the blade body (1).

5. The combustion engine turbine blade based on the zoned composite cooling structure according to claim 1, characterized in that: The front and rear rows of the double-layer S-shaped ribs (4) are provided with interconnected diamond-shaped holes (13) at the overlapping curved portions.

6. The combustion engine turbine blade based on the zoned composite cooling structure according to claim 1, characterized in that: Inwardly concave arc grooves (14) are arranged at intervals on the upper and lower planes of each wedge-shaped rib of the double-layer wedge-shaped matrix rib (5).

7. The combustion engine turbine blade based on the zoned composite cooling structure according to claim 1, characterized in that: A connecting column (15) is fixedly connected between the rear wall surface of the partition plate (6) and the rear inner wall surface of the solid inner layer (3).

8. The combustion engine turbine blade based on the zoned composite cooling structure according to claim 1, characterized in that: The cross section of the cooling channel (8) is an irregular trapezoid, and the width of the cooling channel (8) gradually shrinks from the rectangular cavity (3-1) to the wedge-shaped cavity (3-2).

9. The combustion engine turbine blade based on the zoned composite cooling structure according to claim 1, characterized in that: The exhaust hole (10) comprises a central hole (10-1) with a diameter of 2 mm, and four small holes (10-2) with a diameter of 1 mm are evenly distributed circumferentially on the contour line of the central hole (10-1). The central hole (10-1) and the small holes (10-2) form a plum blossom-shaped exhaust hole (10).

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