CN213039330U

CN213039330U - Gas turbine engine turbine rotor blade with blade top integrated structure

Info

Publication number: CN213039330U
Application number: CN202022306108.7U
Authority: CN
Inventors: 董瑞佳; 张洪国; 孟广双
Original assignee: TANGSHAN INDUSTRIAL VOCATIONAL TECHNICAL COLLEGE
Current assignee: TANGSHAN INDUSTRIAL VOCATIONAL TECHNICAL COLLEGE
Priority date: 2020-10-16
Filing date: 2020-10-16
Publication date: 2021-04-23
Anticipated expiration: 2030-10-16

Abstract

The utility model discloses a gas turbine engine turbine rotor blade with a blade top integrated structure, which comprises a blade root, a blade profile with a blade leading edge and a blade trailing edge and a blade platform for connecting the blade root and the blade profile; a cooling channel enhanced convection cooling loop for cooling the top of the blade is arranged in the blade profile of the blade, and a connecting hole is formed between the cooling channels; the top of the blade profile is provided with a top sealing cover with a groove, and a top air film hole and a raised rib structure are respectively arranged in the groove; the leakage amount and the leakage loss of the blade top are reduced, the cooling effect of the top of the turbine blade is further enhanced, and the efficiency and the overall performance of the turbine are improved.

Description

Gas turbine engine turbine rotor blade with blade top integrated structure

Technical Field

The utility model relates to a gas turbine engine. More particularly, the present invention relates to a blade structure for a gas turbine engine.

Background

The gas turbine engine has the characteristics of light weight, small volume, large single-machine power, quick start, less pollution, high thermal efficiency, good economy and the like. Gas turbine engines generally include a turbine section downstream of a combustion section that is rotatable with a compressor section to rotate and operate the gas turbine engine to generate power, such as propulsive thrust. Conventional gas turbine engines generally include a fan section, a compression section, a combustion section, and a turbine section that define a high pressure turbine arranged in serial flow with an intermediate pressure turbine, a low pressure turbine.

Currently, the pre-cooling temperature of a gas turbine engine turbine blade has already exceeded the withstand temperature of its material. The blade can be guaranteed to work safely and reliably, and the blade is mainly achieved through two ways, namely the high-temperature resistance of the material is improved, and the temperature of the blade is reduced through a cooling mode with stronger cooling capacity. On the other hand, there are recognized problems due to combustion gas leakage between the rotating tips of the turbine blades and the stationary casing surrounding them. Such leakage is sometimes referred to as "tip leakage".

Disclosure of Invention

An object of the utility model is to provide an improve gas turbine engine turbine blade's heat transfer performance, reduce most advanced leakage, improve a gas turbine engine turbine blade of comprehensive turbine thermal efficiency.

The purpose of the utility model is realized like this:

a gas turbine engine turbine rotor blade having a tip composite, wherein the gas turbine engine is a high bypass turbofan gas turbine engine, comprising: an air intake, a propulsion fan, a low pressure compressor, a high pressure compressor, a combustor including a high pressure turbine, an intermediate pressure turbine and a low pressure turbine and an exhaust diffuser;

so that the air entering the air inlet is accelerated by the propulsion fan, creating two air flows: a first gas flow into the low pressure compressor and a second gas flow providing propulsive thrust; the low-pressure compressor compresses a first airflow, and then conveys the air to the high-pressure compressor for further compression; compressed air discharged from the high pressure compressor is directed into a combustor where the compressed air is mixed with fuel and the mixture is combusted; the resulting hot combustion products are then expanded and drive a high pressure turbine, an intermediate pressure turbine and a low pressure turbine, and then discharged through an exhaust diffuser; the high pressure turbine, the intermediate pressure turbine and the low pressure turbine drive the high pressure compressor and the low pressure compressor, respectively, and the propulsion fan through interconnecting shafts; the high pressure turbine, the intermediate pressure turbine and the low pressure turbine drive the high pressure compressor and the low pressure compressor, respectively, and the propulsion fan through interconnecting shafts; the high-pressure turbine, the medium-pressure turbine and the low-pressure turbine are all provided with turbine rotor blades, and the turbine rotor blades are provided with blade top integrated structures and comprise blade roots, blade profiles with blade leading edges and blade trailing edges and blade platforms for connecting the blade roots and the blade profiles; a cooling channel enhanced convection cooling loop for cooling the top of the blade is arranged in the blade profile of the blade, and a connecting hole is formed between the cooling channels;

the top of the blade profile is provided with a top sealing cover with a groove, the groove is internally provided with a top air film hole and a raised rib structure respectively, the top air film hole is distributed at intervals along the central lines of a pressure surface and a suction surface of the top of the blade profile, and the diameter of the top air film hole is gradually reduced from the front edge to the rear edge;

the raised rib structures are formed in a manner that a plurality of Z shapes are inserted into the middle of the top air film hole and arranged on the top cover, wherein the starting points of the Z-shaped raised rib structures are positioned at the front edge of the blade, the end points of the Z-shaped raised rib structures are positioned at the rear edge of the blade, and the middle turning points of the Z-shaped raised rib structures are respectively connected with the pressure surface and the suction surface of the blade;

the raised rib structure and the top sealing cover are integrally processed; the top cover is separately manufactured and is mounted to the blade tip in a welded fashion to form an integral structure with the blade tip.

The height of the raised rib structures accounts for 50% of the depth of the grooves.

The top closure surface is coated with a thermal barrier coating.

The utility model has the advantages that: the utility model provides a gas turbine engine turbine blade is used for improving the loss that turbine blade tip leakage flow, reduction blade tip leakage flow and leakage flow caused to further strengthen turbine blade top cooling effect, improve the efficiency and the wholeness ability of turbine.

Drawings

A full and enabling disclosure of the present invention, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures, in which:

FIG. 1 is a schematic cross-sectional view of an exemplary gas turbine engine incorporating an exemplary embodiment of a turbine section, in accordance with aspects of the present disclosure;

FIG. 2 is a schematic cross-sectional view of a turbine blade of the turbine section of FIG. 1;

FIG. 3 is a schematic view of a turbine blade tip.

Repeat use of reference characters in the present specification and drawings is intended to represent same or analogous features or elements of the invention.

Detailed Description

Reference now will be made in detail to embodiments of the invention, one or more examples of which are illustrated in the drawings. The examples are provided as illustrations of the invention and not as limitations of the invention. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made in the present invention without departing from the scope or spirit of the invention. For example, features illustrated or described as part of one embodiment can be used with another embodiment to yield still a further embodiment. It is therefore intended that the present invention cover such modifications and variations as come within the scope of the appended claims and their equivalents.

In the context of figure 1 of the drawings,

the gas turbine engine is generally indicated at 10, has an axis X-X and comprises an axial flow series of an air intake 11, a propulsion fan 12, a low pressure compressor 13, a high pressure compressor 14, a combustor 15, including a high pressure turbine 16, an intermediate pressure turbine 17 and a low pressure turbine 18, and an exhaust diffuser 19.

Gas turbine engine 10 is a high bypass turbofan gas turbine engine such that air entering air intake 11 is accelerated by a propulsion fan 12 to produce two airflows: a first flow entering the low pressure compressor 13 and a second flow providing propulsive thrust. The low pressure compressor 13 compresses the first air stream and the low pressure compressor delivers the air to the high pressure compressor 14 for further compression.

The compressed air discharged from the high-pressure compressor 14 is directed into the combustor 15, where the compressed air is mixed with fuel and the mixture is combusted. The resulting hot combustion products then expand and drive the high pressure turbine 16, the intermediate pressure turbine 17 and the low pressure turbine 18, and are then exhausted through the exhaust diffuser 19 to provide additional propulsive force. The high-pressure turbine 16, the intermediate-pressure turbine 17 and the low-pressure turbine 18 drive the high-pressure compressor 13 and the low-pressure compressor 14 and the propulsion fan 12, respectively, via interconnecting

shafts

26, 28, 30.

In the context of figure 2 of the drawings,

the high-pressure turbine 16, the intermediate-pressure turbine 17 and the low-pressure turbine 18 each have a turbine rotor blade with a blade tip complex structure comprising a blade root 1, a blade profile 2 with a blade leading edge 20 and a blade trailing edge 21, and a blade platform 3 connecting the blade root 1 and the blade profile 2; the blade profile 2 is internally provided with cooling channels 5 for cooling the top of the blade, and a connecting hole 6 is formed between the cooling channels 5.

In the context of figure 3 of the drawings,

the top of the blade profile 2 is provided with a top sealing cover 4 with a groove, a top air film hole 8 and a raised rib structure 7 are respectively arranged in the groove, the top air film hole 8 is distributed at intervals along the central lines of a pressure surface and a suction surface at the top of the blade profile 2, and the diameter of the top air film hole 8 is gradually reduced from the front edge to the rear edge;

the raised rib structures 7 are formed in a manner that a plurality of Z-shaped raised rib structures 7 are inserted in the middle of the top air film hole 8 and arranged on the top cover 4, wherein the starting points of the Z-shaped raised rib structures 7 are positioned at the front edge 20 of the blade, the end points of the Z-shaped raised rib structures 7 are positioned at the rear edge 21 of the blade, and the middle turning points of the Z-shaped raised rib structures 7 are respectively connected with the pressure surface and the suction surface of the blade;

the raised rib structures 7 and the top cover 4 are integrally processed; the tip cap 4 is manufactured separately and is mounted to the blade tip in the form of a weld, thereby forming an integral structure with the blade tip.

The top cover 4 and the partition plate inside the blade form a transverse cooling channel in a cooling channel reinforced convection cooling loop, and the transverse cooling channel is provided with at least one opening leading to a main flow of fuel gas outside the blade; the cooling gas in the transverse cooling channel flows between the blade leading edge 20 and the blade trailing edge 21, and finally flows into the main gas flow from the top film hole 8 and the opening of the transverse cooling channel. Through protruding 7 structures of the fin structure, the air film is pressed in the groove, the air film coverage rate of two sides of the groove surface of the blade top is enhanced, the groove surface of the blade top is enabled to obtain an even and reasonable air film cooling effect, the heat exchange area in the groove surface is increased, the local high temperature and the local thermal stress of the groove surface of the blade top are effectively reduced, and reasonable and effective thermal protection capability is provided. And flow separation in the grooves is caused by the raised fin structures 7 and the top air film holes 8, so that flow mixing in the grooves is increased, leakage amount and leakage loss of the blade tops are reduced, the cooling effect of the tops of the turbine blades is further enhanced, and the efficiency and the overall performance of the turbine are improved.

Because the blade is whole by blade leading edge 20 to blade trailing edge 21 direction bending, present certain radian, in order to improve the evenly distributed performance of cooling working medium in the blade top groove face, the utility model discloses in, top air film hole 8's central line direction is unanimous with blade leading edge 20 to the crooked direction of blade trailing edge 21. The distribution directions of the Z-shaped raised rib structures 7 and the top air film holes 8 are the same as the curvature direction of the whole radian of the blade, so that the cooling medium in the grooves at the top of the blade is distributed along the curvature direction, the distribution uniformity of the cooling medium is improved, and the cooling effect on the two sides of the grooves is better.

In order to better press the cooling air film in the groove of the blade tip, the height of the raised rib structure 7 accounts for 50% of the depth of the groove.

In order to further improve the thermal protection capability of the concave surface of the blade top, the surface of the top cover 4 is coated with a thermal insulation coating.

The above description is only a preferred embodiment of the invention, and should not be taken as limiting the invention, and any modifications, equivalents, improvements and the like that are within the spirit and principle of the invention should be included in the scope of the invention.

Claims

1. A gas turbine engine turbine rotor blade having a tip composite,

the gas turbine engine is a high bypass turbofan gas turbine engine comprising: an air intake, a propulsion fan, a low pressure compressor, a high pressure compressor, a combustor including a high pressure turbine, an intermediate pressure turbine and a low pressure turbine and an exhaust diffuser;

so that the air entering the air inlet is accelerated by the propulsion fan, creating two air flows: a first gas flow into the low pressure compressor and a second gas flow providing propulsive thrust; the low-pressure compressor compresses a first airflow, and then conveys the air to the high-pressure compressor for further compression; compressed air discharged from the high pressure compressor is directed into a combustor where the compressed air is mixed with fuel and the mixture is combusted; the resulting hot combustion products are then expanded and drive a high pressure turbine, an intermediate pressure turbine and a low pressure turbine, and then discharged through an exhaust diffuser; the high pressure turbine, the intermediate pressure turbine and the low pressure turbine drive the high pressure compressor and the low pressure compressor, respectively, and the propulsion fan through interconnecting shafts;

it is characterized in that the preparation method is characterized in that,

the high-pressure turbine, the medium-pressure turbine and the low-pressure turbine are all provided with turbine rotor blades, and the turbine rotor blades are provided with blade top integrated structures and comprise blade roots, blade profiles with blade leading edges and blade trailing edges and blade platforms for connecting the blade roots and the blade profiles; a cooling channel enhanced convection cooling loop for cooling the top of the blade is arranged in the blade profile of the blade, and a connecting hole is formed between the cooling channels;

the raised rib structures are formed in a manner that a plurality of Z shapes are inserted into the middle of the top air film hole and arranged on the top cover, the starting points of the Z-shaped raised rib structures are positioned at the front edge of the blade, the ending points of the Z-shaped raised rib structures are positioned at the rear edge of the blade, and the middle turning points of the Z-shaped raised rib structures are respectively connected with the pressure surface and the suction surface of the blade;

2. The gas turbine engine turbine rotor blade with tip composite in accordance with claim 1, wherein the height of the raised rib structure is 50% of the depth of the groove.

3. The gas turbine engine turbine rotor blade with tip composite in accordance with claim 1, wherein the tip cap surface is coated with a thermal barrier coating.