CN219262455U

CN219262455U - Turbine blade leading edge cooling structure of gas turbine

Info

Publication number: CN219262455U
Application number: CN202320344990.5U
Authority: CN
Inventors: 肖俊峰; 高松; 段静瑶; 李园园; 于飞龙; 刘战胜; 何伟; 张浩浩; 武耀族; 伍赫
Original assignee: Xian Thermal Power Research Institute Co Ltd
Current assignee: Xian Thermal Power Research Institute Co Ltd
Priority date: 2023-02-28
Filing date: 2023-02-28
Publication date: 2023-06-27
Anticipated expiration: 2033-02-28

Abstract

The application discloses a cooling structure of the front edge of a turbine blade of a gas turbine, wherein the cooling structure is arranged inside the front edge of the turbine blade and comprises an air supply cavity and an impact cavity which are separated by a partition plate, and a plurality of impact holes are formed in the partition plate and are communicated with the air supply cavity and the impact cavity; the cross section area of the air supply cavity is gradually reduced from the blade root of the turbine blade to the blade top of the turbine blade, the thickness of the partition plate is also gradually reduced, and the distance between the same impact hole and the pressure surface of the turbine blade is smaller than the distance between the same impact hole and the suction surface of the turbine blade; the cooling structure further comprises a plurality of air film cooling holes formed in the side wall of the front edge of the turbine blade, and the air film cooling holes are communicated with the impact cavity. The flow loss of the cold air can be generated from the blade root to the blade tip direction, but because the height of the impact hole at the blade tip is correspondingly reduced and the cross section area of the air supply cavity is correspondingly reduced, the flow speed and the flow rate of the cold air after entering the impact cavity at the blade tip can be ensured, and then the impact cooling effect and the air film cooling effect of the cold air at the blade tip are ensured.

Description

Turbine blade leading edge cooling structure of gas turbine

Technical Field

The application relates to the technical field of impeller machines, in particular to a turbine blade leading edge cooling structure of a gas turbine.

Background

In order to develop more advanced gas turbines, the gas temperature of the turbine inlet is continuously improved, and the performance of turbine blade materials can not meet the increasing turbine inlet temperature, so that it is necessary to develop turbine blade cooling technology, reduce the thermal load born by the blades on the premise of ensuring the aerodynamic performance of the blades, prolong the service life of the blades, and improve the reliability and safety of the gas turbine.

For a turbine blade, the leading edge of the turbine blade is one of the regions of highest turbine blade heat load and is also one of the regions of the turbine blade that are most susceptible to erosion. The practical gas turbine blade generally uses an impact-gas film composite cooling structure to reduce the heat load of the front edge of the blade, and the cooling design improves the heat exchange coefficient of the internal structure and simultaneously ensures that the gas supply temperature of the gas film cooling is not too high and the flow distribution is reasonable.

The existing turbine blade impact-air film composite cooling device comprises an air supply cavity and an impact cavity which are arranged in the turbine blade, wherein the air supply cavity and the impact cavity are separated by a partition plate, impact holes are formed in the partition plate, the impact holes are communicated with the air supply cavity and the impact cavity, the partition plate is a straight plate, and the thicknesses of the partition plate and the impact cavity are the same.

The turbine blade impact-air film composite cooling device has better cooling effect, but still has some defects: for example, the cooling air must have a flow loss during the flow from the blade root to the blade tip of the turbine blade, which can affect the impingement cooling effect and the film cooling effect at the blade tip and thus lead to an excessively high temperature in the region of the blade tip at the blade leading edge.

Disclosure of Invention

Therefore, the technical problem to be solved by the application is to overcome the defect that after flow loss is experienced in the process of flowing cool air from the blade root to the blade tip of a turbine blade in the prior art, the impact cooling effect and the air film cooling effect at the blade tip of the front edge are affected, so that the cooling structure of the front edge of the turbine blade of the gas turbine is provided.

In order to solve the technical problems, the technical proposal of the application is as follows,

the cooling structure is arranged in the front edge of the turbine blade and comprises an air supply cavity and an impact cavity which are separated by a partition plate, and a plurality of impact holes are formed in the partition plate so as to be communicated with the air supply cavity and the impact cavity; the cross-sectional area of the air supply cavity is gradually reduced from the blade root of the turbine blade to the blade top direction of the turbine blade, and the thickness of the partition plate is also gradually reduced; the cooling structure further comprises a plurality of air film cooling holes formed in the side wall of the front edge of the turbine blade, and the air film cooling holes are communicated with the impact cavity.

Further, the impingement holes are located on a side of the gas supply chamber and the impingement chamber adjacent to the pressure face of the turbine blade.

Further, the cross section of the impact hole is in a runway shape, the ratio of the long axis of the impact hole to the short axis of the impact hole is a, and the value range of a is 1-2.

Further, the angle between the long axis of the impingement hole and the blade height direction of the turbine blade is alpha, and the value of alpha is in the range of 0 DEG to 8 deg.

Further, the impact hole is provided with a rounding at the positions adjacent to the air supply cavity and the impact cavity respectively; the ratio of the radius of the rounding to the short axis of the impact hole is b, and the value range of b is 0.5-2.5.

Further, the thickness of the spacer is 1% -5% of the blade chord of the turbine blade.

Further, the impact holes are equidistantly distributed along the height direction of the blade.

Further, at least one row of air film cooling holes is formed along the blade height direction, and the number of the impact holes is not less than the number of the air film cooling holes in the same row.

Further, the axis of the film cooling hole is disposed obliquely to the normal of the leading edge sidewall of the turbine blade.

Further, the impingement cavity, the gas supply cavity, and the impingement hole are integrally cast with the turbine blade.

The technical scheme of the application has the following advantages:

1. according to the cooling structure of the turbine blade front edge of the gas turbine, the thickness of the partition plate is gradually reduced from the blade root to the blade top direction, so that the height of the impact hole is also gradually reduced, and therefore, although the air conditioner can generate flow loss from the blade root to the blade top direction, the air conditioner can ensure the flow velocity and flow rate after entering the impact cavity at the blade top because the height of the impact hole through which the air conditioner needs to pass at the blade top is correspondingly reduced and the cross section area of the air supply cavity is correspondingly reduced, and further the impact cooling effect and the air film cooling effect of the air conditioner at the blade top are ensured.

Drawings

In order to more clearly illustrate the embodiments of the present application or the technical solutions in the prior art, the drawings that are needed in the description of the embodiments or the prior art will be briefly described below, and it is obvious that the drawings in the following description are some embodiments of the present application, and other drawings may be obtained according to these drawings without inventive effort for a person skilled in the art.

FIG. 1 is a schematic view of the internal structure of a gas turbine bucket of the present application;

FIG. 2 is an enlarged partial schematic view of FIG. 1;

FIG. 3 is a cross-sectional view of section B of FIG. 1;

fig. 4 is a schematic plan view of the impingement holes in the diaphragm of the present application.

The reference numerals are used to describe the components,

1. turbine blades; 11. blade root; 12. leaf tops; 15. a leading edge; 2. an air supply chamber; 3. an impingement cavity; 23. a partition plate; 4. an impingement hole; 5. air film cooling holes; 6. rounding; 7. a pressure surface; 8. a suction surface; 9. a long axis of the impingement hole; 10. short axis of the impact hole.

Detailed Description

The following description of the embodiments of the present application will be made apparent and fully in view of the accompanying drawings, in which some, but not all embodiments of the utility model are shown. All other embodiments, which can be made by one of ordinary skill in the art without undue burden from the present disclosure, are within the scope of the present disclosure.

In the description of the present application, it should be noted that the directions or positional relationships indicated by the terms "center", "upper", "lower", "left", "right", "vertical", "horizontal", "inner", "outer", etc. are based on the directions or positional relationships shown in the drawings, are merely for convenience of description of the present application and to simplify the description, and do not indicate or imply that the devices or elements referred to must have a specific orientation, be configured and operated in a specific orientation, and thus should not be construed as limiting the present application. Furthermore, the terms "first," "second," and "third" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance.

In the description of the present application, it should be noted that, unless explicitly specified and limited otherwise, the terms "mounted," "connected," and "connected" are to be construed broadly, and may be either fixedly connected, detachably connected, or integrally connected, for example; can be mechanically or electrically connected; can be directly connected or indirectly connected through an intermediate medium, and can be communication between two elements. The specific meaning of the terms in this application will be understood by those of ordinary skill in the art in a specific context.

In addition, the technical features described below in the different embodiments of the present application may be combined with each other as long as they do not collide with each other.

Examples

In the following, embodiments of the present utility model will be described in terms of a heavy duty gas turbine blade, the leading edge of which is one of the regions of the turbine blade where the thermal load is highest, and also one of the regions of the blade where it is most susceptible to erosion, typically using an impingement-film composite cooling structure. Thus, the leading edge cooling design should consider the rationality of film cooling flow distribution in addition to the elevation of the internal structural heat transfer coefficient.

As shown in fig. 1 to 4, the present embodiment provides a turbine blade leading edge cooling structure of a gas turbine, including a gas supply chamber 2, an impingement chamber 3, an impingement hole 4, and a film cooling hole 5 provided in a leading edge 15 of a turbine blade 1. In this embodiment, the gas supply chamber 2, the impingement chamber 3, and the impingement holes 4 are cast integrally with the turbine blade 1.

The air supply cavity 2 and the impact cavity 3 are respectively positioned at two sides of the partition plate 23, but are communicated through a plurality of impact holes 4 penetrating through the partition plate 23. As shown in fig. 1, the cross-sectional area of the air supply cavity 2 gradually decreases from the blade root 11 of the turbine blade 1 to the blade tip 12 of the turbine blade 1, and the thickness L of the partition plate 23 separating the impingement cavity 3 from the air supply cavity 2 gradually decreases, wherein b is the blade chord length, and the height of the impingement hole 4 (as indicated by the arrow C-C in fig. 1) also gradually decreases due to the gradual decrease of the thickness L of the partition plate 23. In the present embodiment, the chord length of the blade at the blade root 11 is preferably 132mm, the chord length of the blade at the blade tip 12 is preferably 126mm, the thickness L of the partition 23 at the blade root 11 is preferably 4.65mm, and the thickness L of the partition 23 at the blade tip 12 is preferably 2.15mm, and the coupling effect of impingement cooling and film cooling is comprehensively considered.

The cross-sectional shape of the impingement holes 4 is circular, oval or racetrack. As shown in fig. 4, in the present embodiment, the cross-sectional shape of the impingement hole 4 is a racetrack shape, the ratio of the major axis 9 of the impingement hole to the minor axis 10 of the impingement hole is a, the value of a is in the range of 1-2, and preferably the value of a is 1.3. The angle between the long axis 9 of the impingement hole and the direction of the blade height of the turbine blade 1 (indicated by the arrow A-A in fig. 1) is alpha, which has a value in the range of 0 deg. -8 deg.. The impact holes 4 are provided with a round 6 at the connection with the air supply cavity 2 and the impact cavity 3, the ratio of the radius of the round 6 to the short axis 10 of the impact hole is b, and the value of b is in the range of 0.5-2.5, and in the embodiment, the value of b is preferably 1.5. The cross section of the impact hole 4 is track-shaped, and the connection parts of the impact hole 4, the air supply cavity 2 and the impact cavity 3 are provided with round-down circles 6, so that the flow loss of cold air can be reduced, and the cooling efficiency of the front edge 15 of the turbine blade 1 can be improved.

In this embodiment, the impingement holes 4 are disposed near the pressure surface 7, that is, the distance between the impingement holes 4 and the outer wall surface of the turbine blade 1 on the side of the pressure surface 7 is smaller than the distance between the impingement holes and the outer wall surface of the turbine blade 1 on the side of the suction surface 8, so that the cooling gas can generate swirling flow impingement on the cooling wall surface to be cooled in the impingement cavity 3, and the impingement cooling efficiency is improved. The impingement holes 4 are equally distributed along the height direction of the blade, the number of impingement holes 4 should be not less than the number of film cooling holes 5 arranged in a row at the leading edge 15, in this embodiment the number of impingement holes 4 is 23.

Three exhaust film cooling holes 5 are formed in the cooling wall surface, close to the front edge 15, of the turbine blade 1 along the blade height direction, of course, more rows of the three exhaust film cooling holes can be formed, and the impact cold air can be sprayed to the outer wall surface of the turbine blade 1 through the air film cooling holes 5, so that the design can enable the cold air coming out of the air film cooling holes 5 to cover more outer wall surfaces of the turbine blade 1, better air film cooling is formed, and the surface temperature of the turbine blade 1 is reduced. In addition, the axis of the air film cooling hole 5 and the normal line of the wall surface to be cooled are arranged at an included angle, so that cooling gas can better cover the outer wall surface of the turbine blade 1 after coming out of the air film cooling hole 5, and the air film cooling efficiency is improved.

In the present embodiment, since the thickness of the partition 23 gradually decreases from the blade root 11 toward the blade tip 12, so that the height of the impingement holes 4 also gradually decreases, and thus, although the flow loss of the cold air from the blade root 11 toward the blade tip 12 occurs, since the height of the impingement holes 4 through which the cold air needs to pass at the blade tip 12 decreases correspondingly and the cross-sectional area of the air supply chamber 2 decreases correspondingly, the flow velocity and flow rate of the cold air after entering the impingement chamber 3 at the blade tip 12 can be ensured, and further the impingement cooling effect and the film cooling effect of the cold air at the blade tip 12 can be ensured.

In a word, the cooling structure for the front edge of the turbine blade of the gas turbine provided by the embodiment can lead the distribution of the cold air flow to be more reasonable in terms of the whole, further effectively improve the cooling efficiency of the front edge 15 of the turbine blade 1, reduce the temperature of the front edge 15 area of the turbine blade 1, prolong the service life of the turbine blade 1 and improve the operation safety.

It is apparent that the above examples are given by way of illustration only and are not limiting of the embodiments. Other variations or modifications of the above teachings will be apparent to those of ordinary skill in the art. It is not necessary here nor is it exhaustive of all embodiments. While nevertheless, obvious variations or modifications may be made to the embodiments described herein without departing from the scope of the utility model.

Claims

1. The turbine blade leading edge cooling structure of the gas turbine is characterized in that the cooling structure is arranged in the front edge (15) of the turbine blade (1) and comprises an air supply cavity (2) and an impact cavity (3) which are separated by a partition plate (23), and a plurality of impact holes (4) are formed in the partition plate (23) to be communicated with the air supply cavity (2) and the impact cavity (3); the cross-sectional area of the air supply cavity (2) gradually decreases from the blade root (11) of the turbine blade (1) to the blade tip (12) of the turbine blade (1), and the thickness of the partition plate (23) also gradually decreases; the cooling structure further comprises a plurality of air film cooling holes (5) formed in the side wall of the front edge of the turbine blade (1), and the air film cooling holes (5) are communicated with the impact cavity (3).

2. The turbine blade leading edge cooling structure of a gas turbine according to claim 1, characterized in that the impingement holes (4) are located on the side of the gas supply chamber (2) and the impingement chamber (3) close to the pressure surface (7) of the turbine blade (1).

3. The turbine blade leading edge cooling structure of a gas turbine according to claim 1, wherein the cross-sectional shape of the impingement hole (4) is a racetrack shape, and the ratio of the long axis (9) of the impingement hole to the short axis (10) of the impingement hole is a, and the value of a ranges from 1 to 2.

4. A turbine blade leading edge cooling structure of a gas turbine according to claim 3, characterized in that the angle between the long axis (9) of the impingement hole and the blade height direction of the turbine blade (1) is α, the value of α being in the range of 0 ° to 8 °.

5. A turbine blade leading edge cooling structure of a gas turbine according to claim 3, wherein the impingement holes (4) are provided with a radius (6) adjacent to the gas supply chamber (2) and the impingement chamber (3), respectively; the ratio of the radius of the rounding (6) to the short axis (10) of the impact hole is b, and the value range of b is 0.5-2.5.

6. The turbine blade leading edge cooling structure of a gas turbine according to claim 1, wherein the thickness value of the partition plate (23) is 1% -5% of the blade chord value of the turbine blade (1).

7. Turbine blade leading edge cooling structure of a gas turbine according to claim 1, characterized in that the impingement holes (4) are equally distributed in the blade height direction.

8. The turbine blade leading edge cooling structure of a gas turbine according to claim 7, wherein the film cooling holes (5) are provided in at least one row in the blade height direction, and the number of the impingement holes (4) is not less than the number of the film cooling holes (5) in the same row.

9. The turbine blade leading edge cooling structure of a gas turbine according to claim 8, characterized in that the axis of the film cooling hole (5) is disposed obliquely to the normal of the leading edge side wall of the turbine blade (1).

10. Turbine blade leading edge cooling structure of a gas turbine according to claim 1, characterized in that the impingement cavity (3), the gas supply cavity (2) and the impingement hole (4) are cast integrally with the turbine blade (1).