CN112459849B

CN112459849B - Cooling structure for turbine blade of gas turbine

Info

Publication number: CN112459849B
Application number: CN202011162279.5A
Authority: CN
Inventors: 刘庆鹤; 朱凯迪; 刘文文; 刘宇
Original assignee: Harbin Guanghan Gas Turbine Co ltd
Current assignee: Harbin Guanghan Gas Turbine Co ltd
Priority date: 2020-10-27
Filing date: 2020-10-27
Publication date: 2022-08-30
Anticipated expiration: 2040-10-27
Also published as: CN112459849A

Abstract

The invention aims to provide a cooling structure for a turbine blade of a gas turbine, wherein a front edge independent chamber, a middle chamber and a tail edge chamber are arranged in a blade body, the front edge independent chamber is separated from the middle chamber, the middle chamber is communicated with the tail edge chamber at the blade top position, a front edge air film hole and a tail edge split seam are arranged on the blade body, the front edge air film hole is communicated with the front edge independent chamber, the tail edge split seam is communicated with the tail edge chamber, a pressurization bleed air pipe communicated with the front edge independent chamber is arranged below the front edge independent chamber, an end face air inlet hole is arranged on the end face of a tenon, a first bottom face air inlet hole and a second bottom face air inlet hole are arranged on the bottom face of the tenon, and cooling air from the end face air inlet hole, the first bottom face air inlet hole and the second bottom face air inlet hole enters the middle chamber to be converged. The independent cooling cavity structure of blade leading edge directly faces high temperature gas thermal shock, and the blade produces showing pressure boost effect in rotatory process. The blade has the vortex effect at rotatory in-process tenon terminal surface inlet port, has the function of improving air flow rate, does benefit to and strengthens heat transfer effect.

Description

Cooling structure for turbine blade of gas turbine

Technical Field

The invention relates to a cooling structure of a gas turbine, in particular to a cooling structure of a gas turbine.

Background

The gas turbine has the characteristics of small volume, high single-machine power, high thermal efficiency and the like. The high-pressure turbine stage 1 movable blade is a part with the highest working bearing temperature, complex stress and the worst working environment in a gas turbine rotor component, and is a key component with the highest risk in the running process of the gas turbine. The bearing temperature of the 1 st stage movable blade of the high-pressure turbine directly determines the performance of the gas turbine, and at present, two ways for improving the temperature resistance are provided, wherein one way is to adopt high-quality high-temperature-resistant alloy materials, and the other way is to adopt an internal cooling structure for the blade. The turbine blade cooling structure is limited by the temperature resistance limitation of the current material under the high-temperature condition, breakthrough improvement cannot be realized in a short time, and the development of the turbine blade cooling structure technology is an important means for improving the temperature resistance of the turbine blade. The blade adopts hollow structure more, and the internal design supplies the passageway that cooling air flows, and the wall is equipped with the rib structure and reaches and cool off the best heat transfer effect of air, reduces blade temperature, has the cooling hole at the wall design simultaneously, forms the air film on blade surface and plays thermal-insulated effect. Therefore, the design quality of the blade cooling structure is a key factor influencing the cooling effect.

Disclosure of Invention

The invention aims to provide a cooling structure for a turbine blade of a gas turbine, which increases the heat exchange area between cooling air and a blade body and improves the heat exchange efficiency.

The purpose of the invention is realized as follows:

the invention relates to a cooling structure for a turbine blade of a gas turbine, which comprises a blade body and a tenon and is characterized in that: the blade body is internally provided with a front edge independent cavity, a middle cavity and a tail edge cavity, the front edge independent cavity is separated from the middle cavity, the middle cavity is communicated with the tail edge cavity at the blade top position, the rest part is separated, the blade body is provided with a front edge air film hole and a tail edge split seam, the front edge air film hole is communicated with the front edge independent cavity, the tail edge split seam is communicated with the tail edge cavity, a pressurizing air-entraining pipe communicated with the front edge independent cavity is arranged below the front edge independent cavity, the end face of the tenon is provided with an end face air inlet hole, the bottom face of the tenon is provided with a first bottom face air inlet hole and a bottom face second air inlet hole, and cooling air from the end face air inlet hole, the first air inlet hole and the second air inlet hole enters the middle cavity to be converged.

The present invention may further comprise:

1. turbulence ribs are arranged in the independent front edge cavity, the middle cavity and the tail edge cavity.

2. The turbulence ribs in the front edge independent chambers form an angle of 45 degrees with the horizontal direction, the turbulence ribs in the middle chambers form an angle of 135 degrees with the horizontal direction, and the turbulence ribs in the tail edge chambers form an angle of 115 degrees with the horizontal direction.

3. Staggered ribs are arranged on one side, close to the tail edge cleft seam, of the tail edge cavity.

4. The air inlet hole of the tenon end face is a trapezoidal air inlet hole, and the first bottom surface air inlet hole and the second bottom surface air inlet hole are square air inlet holes.

The invention has the advantages that: the independent cooling cavity structure of blade leading edge is direct to high temperature gas thermal shock, and through ingenious design bleed air pipeline, the blade produces the effect of showing the pressure boost in rotatory in-process, leads to the pressure of cooling cavity to be greater than leading edge incoming flow gas pressure, and the wall adopts the design of large size vortex rib heat transfer structure in addition, and the cooling air sprays through leading edge air film hole after abundant heat transfer and goes out, forms the cooling air film at blade leading edge surface, plays thermal-insulated effect.

The middle cooling channel cavity is provided with a turbulence rib structure for heat exchange, the air flow enters the tail edge cooling channel cavity after being fully exchanged heat, the wall surface is provided with a staggered rib heat exchange structure, and the blades have a turbulence effect on the air inlet hole of the tenon end surface in the rotating process, so that the air flow rate is improved, and the heat exchange effect is favorably enhanced.

Drawings

FIG. 1 is a schematic view of the structure of the present invention;

FIG. 2 is a schematic transverse cross-sectional view;

fig. 3 is a schematic longitudinal cross-section.

Detailed Description

The invention will now be described in more detail by way of example with reference to the accompanying drawings in which:

with reference to fig. 1-3, the invention is formed by investment casting, and the internal cooling structure is located in the space surrounded by the suction surface and the pressure surface of the blade. The blade adopts a non-blade crown structure and comprises an upper cover plate, a blade body, a lower edge plate, a root extension, a tenon 10 and the like. Three cooling channels are sequentially designed from the front edge to the tail edge of the blade body, wherein the cavity 1 of the front edge cooling channel is independent, the middle cavity 2 and the tail edge cavity 3 are communicated at the blade top position, and airflow flows out from the tail edge cleft 9 of the blade body and is converged into main flow fuel gas after fully exchanging heat through the staggered ribs of the tail edge cavity 3. The end surface of the blade tenon 10 is provided with a cooling air bleed pipeline 4 which leads to the blade front edge cooling chamber 1. A trapezoidal air inlet 5 is designed on one side end face of the tenon 10, meanwhile, a square air inlet is formed in the bottom of the tenon 10, and cooling air is respectively introduced through the end face air inlet and the bottom air inlet and finally flows into a cavity of a cooling channel in the middle of the blade body.

A pressurizing air guide pipe 4 is designed at an inlet of an independent cavity 1 at the front edge of a cooling structure, two rows of air film holes 8 are formed in the air inlet side of the front edge of a blade body, and are machined in an electric spark mode, wherein the diameter of each air film hole 8 is phi 0.4mm, and the number of the air film holes is 22. The wall surface of the inner cavity is provided with a large-size turbulence rib structure, and the rib direction and the horizontal direction form an angle of 45 degrees. Play cooling air pressure boost effect at the rotatory in-process of blade, the independent cavity 1 of leading edge is got into after the pressure boost to the air conditioning of pressure boost bleed pipe 4, through vortex rib and wall intensive heat transfer, erupts through leading edge air film hole 8, forms the air film protective layer at blade leading edge, effectively reduces the blade and bears the destruction of thermal shock.

Middle cavity 2 is linked together at the blade top with trailing edge cavity 3, at the 10 terminal surface designs of tenon 2 structures that admit air, the 10 bottom of tenon designs the first inlet port 6 in bottom surface and 7 inlet structures in bottom surface second inlet port, cooling air joins at middle cavity 2 through these three-strand runner, wherein middle cavity 2 structural design is more novel, the blade has the vortex effect at rotatory in-process 10 tenon terminal surface inlet port 5, improve the air input flow, strengthen the heat transfer effect, the requirement to air pressure has been reduced. The direction of the turbulence ribs on the wall surface of the middle chamber 2 and the horizontal plane form an angle of 45 degrees respectively, and the direction is opposite to that of the front edge independent chamber 1. The airflow enters the trailing edge cavity 3 to cool the blade body after effectively exchanging heat with the wall surface through the large-size turbulence ribs of the middle cavity 2, and then flows out from the trailing edge cleft 9 of the blade through the staggered rib structure and is converged into mainstream fuel gas. The direction of the turbulence ribs on the wall surface of the tail edge cavity 3 is 75 degrees with the horizontal direction, the staggered rib structure of the blade body close to the air outlet side is 60 degrees with the horizontal direction, the cooling air at the tail edge cleft joint 9 flows out in the horizontal direction, the air outlet edge of the tail edge of the blade is provided with cleft joint cooling channels, 15 groups are formed in total, and each group comprises 4 cooling channels.

The turbine cooling blade adopts a unique air inlet structure, independent cooling chamber distribution and a multi-form reinforced heat exchange structure, effectively controls the air inlet and outlet pressure difference, ensures the cooling air flow, improves the heat exchange efficiency, and meets the working requirement of the cooling blade in a high-temperature environment.

Claims

1. A cooling structure for a turbine blade of a gas turbine, comprising a blade body and a tenon, characterized in that: the blade body is internally provided with a front edge independent cavity, a middle cavity and a tail edge cavity, the front edge independent cavity is separated from the middle cavity, the middle cavity is communicated with the tail edge cavity at the blade top position, the rest parts are separated, the blade body is provided with a front edge air film hole and a tail edge split seam, the front edge air film hole is communicated with the front edge independent cavity, the tail edge split seam is communicated with the tail edge cavity, a pressurizing air-entraining pipe communicated with the front edge independent cavity is arranged below the front edge independent cavity, the end face of the tenon is provided with an end face air inlet hole, the bottom face of the tenon is provided with a first bottom face air inlet hole and a second bottom face air inlet hole, and cooling air from the end face air inlet hole, the first bottom face air inlet hole and the second bottom face air inlet hole enters the middle cavity to be converged.

2. The cooling structure for a turbine blade of a gas turbine as set forth in claim 1, wherein: turbulence ribs are arranged in the independent cavity at the front edge, the middle cavity and the tail edge cavity.

3. The cooling structure for a turbine blade of a gas turbine as set forth in claim 2, wherein: the turbulence ribs in the front edge independent chambers form an angle of 45 degrees with the horizontal direction, the turbulence ribs in the middle chambers form an angle of 135 degrees with the horizontal direction, and the turbulence ribs in the tail edge chambers form an angle of 115 degrees with the horizontal direction.

4. The cooling structure for a turbine blade of a gas turbine as set forth in claim 1, wherein: staggered ribs are arranged on one side, close to the tail edge cleft seam, of the tail edge cavity.

5. The cooling structure for a turbine blade of a gas turbine according to any one of claims 1 to 4, wherein: the air inlet hole of the tenon end face is a trapezoidal air inlet hole, and the first bottom surface air inlet hole and the second bottom surface air inlet hole are square air inlet holes.