CN110593961B

CN110593961B - Divided cabin type turbine blade

Info

Publication number: CN110593961B
Application number: CN201910932354.2A
Authority: CN
Inventors: 谢剑; 徐进良; 李文霄; 梁聪; 马杨
Original assignee: North China Electric Power University
Current assignee: North China Electric Power University
Priority date: 2019-09-29
Filing date: 2019-09-29
Publication date: 2020-09-15
Anticipated expiration: 2039-09-29
Also published as: CN110593961A

Abstract

The invention belongs to the technical field of gas turbines and aeroengines, and particularly relates to a divided cabin type turbine blade.A plurality of longitudinal screens are arranged in the blade along the direction from the top of the blade to the root of the blade so as to be divided into a plurality of stages of flow passages, each stage of flow passage is divided into a front cabin and a rear cabin, the top of each front cabin is communicated with the bottom of each rear cabin, and the front cabin of the previous stage of flow passage is communicated with the rear cabin of the next stage of flow passage at the root; the transverse screen is provided with a plurality of jet holes which are connected with a plurality of jet cavities at the front edge of the blade, the jet cavities are provided with air film holes on the suction surface and the pressure surface at the two sides of the blade, and the bottom of the first-stage runner rear cabin is connected with a cooling medium inlet at the root part of the blade. Coriolis force generated in rotation of the blades is utilized to generate a thin liquid film and a large velocity gradient on two heating surfaces of a pressure surface and a suction surface simultaneously, heat transfer is enhanced, the thick liquid film is adsorbed on a transverse screen of a partition channel and becomes a source of jet flow impact cooling and air film cooling, a gas-liquid two-phase cooling medium is fully utilized, and the temperature-resistant neck problem of the turbine blade is solved.

Description

Divided cabin type turbine blade

Technical Field

The invention belongs to the technical field of gas turbines and aero-engines, and particularly relates to a divided cabin type turbine blade.

Background

The gas turbine is the key equipment of the natural gas power station, the aircraft engine is the heart of the aircraft, and the core component of the gas turbine and the aircraft engine is the turbine blade. To obtain excellent performance of high efficiency and large thrust, gas turbine engines and aircraft engines need to increase inlet gas temperature. Correspondingly, the design of turbine blades is subject to severe testing. In order to adapt to the high-temperature gas environment, the development of the turbine blade is divided into two directions, one is to develop high-temperature-resistant advanced materials, and the other is to optimally design a turbine blade cooling system. The former direction needs a huge industrial system for support, the development period is long, the innovative margin in the latter direction is large, the performance of the gas turbine and the aero-engine can be improved in a short time, and the method is particularly suitable for developing countries. At present, a turbine blade cooling system usually adopts single-phase working media such as air, steam and the like, and a corresponding channel optimization design method also aims at the single-phase working media, and comprises the steps of expanding a heating surface, increasing a flow-around column and the like. In fact, the gas-liquid two-phase flow is introduced into the turbine blade cooling channel, and the heat transfer performance is better, so that a heat transfer strengthening method aiming at the gas-liquid two-phase flow cooling working medium needs to be developed by combining the operation characteristics of the turbine blade, and the problem of temperature-resistant neck of the turbine blade is solved.

Disclosure of Invention

Aiming at the problems, the invention creatively provides a divided cabin type turbine blade, a plurality of longitudinal screens are arranged in the blade along the direction from the front edge to the rear edge so as to be divided into a plurality of stages of flow passages, each stage of flow passage is divided into a front cabin and a rear cabin which are communicated with each other at the top and are separated at the bottom by a transverse screen, and the front cabin of the previous stage of flow passage is communicated with the rear cabin of the next stage of flow passage at the root; the transverse screen is provided with a plurality of jet holes which are connected with a plurality of jet cavities at the front edge of the blade, the jet cavities are provided with air film holes on the suction surface and the pressure surface at the two sides of the blade, and the bottom of the first-stage runner rear cabin is connected with a cooling medium inlet at the root part of the blade.

The suction surface and the pressure surface are respectively a front surface and a rear surface of the blade along the rotation direction, the front cabin is formed by a longitudinal screen and the suction surface in a surrounding mode, and the rear cabin is formed by a longitudinal screen and the pressure surface in a surrounding mode.

The width of the front cabin is larger than that of the rear cabin.

The cross section of the air film hole is in a water drop shape and forms an included angle with the tangential direction of the blade molded line.

The transverse screen is made of porous metal foam.

The gas-liquid two-phase cooling medium is air-water two-phase flow or steam-water two-phase flow.

When the blades rotate, gas-liquid two-phase cooling media in the rear cabin flow from the blade roots to the blade tops, and a thin liquid film and a large velocity gradient are formed on the pressure surface under the action of Coriolis force; gas-liquid two-phase cooling medium in the front cabin flows from the blade top to the blade root, and a thin liquid film and a large velocity gradient are formed on the suction surface; the thick liquid film that adsorbs on horizontal screen gets into the efflux chamber through the efflux hole, and the jet impingement cooling is followed the discharge through the air film hole to blade leading edge, carries out the air film cooling to suction surface and pressure surface.

The invention has the beneficial effects that: the transverse screen is creatively arranged to divide the flow channel into a front chamber and a rear chamber, the Coriolis force generated in the rotation process of the turbine blade is utilized to change the liquid film distribution of gas-liquid two-phase flow in the channel, the flow directions of gas-liquid two-phase cooling media in the front chamber and the rear chamber are opposite, a thin liquid film and a large velocity gradient are simultaneously generated on two heating surfaces of a pressure surface and a suction surface, the heat transfer is strengthened, a thick liquid film is adsorbed on the transverse screen of the divided channel and becomes the source of jet impact cooling and gas film cooling, and the gas-liquid two-phase cooling media are fully utilized. Therefore, the invention innovatively designs the flow channel, strengthens the cooling of the blade and ensures the safe operation of the turbine, is an innovative technology in the field of gas turbines and aero-engines and solves the temperature-resistant neck problem of the turbine blade.

Drawings

FIG. 1 is a schematic view of a divided nacelle turbine blade structure and operating principle of the present invention

Reference numbers in the figures: the turbine blade comprises a turbine blade 1, a turbine blade longitudinal screen 2, a turbine blade 3, a turbine blade transverse screen 4, a turbine blade front chamber 5, a turbine blade rear chamber 6, a turbine blade suction surface 7, a turbine blade pressure surface 8, a turbine blade top 9, a turbine blade communicating hole 10, a turbine blade root 11, a turbine blade jet hole 12, a turbine blade front edge 13, a turbine blade jet cavity 14, a turbine blade film hole 15, high-temperature gas 16, a turbine blade gas-liquid two-phase flow cooling medium 17, a turbine blade cooling medium inlet 18, a thin liquid film 19 and a.

Detailed Description

The invention is described in further detail below with reference to the attached drawing figures, but the invention is not limited in any way by the claims.

FIG. 1 is a schematic view of a divided nacelle turbine blade configuration and operating principles of the present invention. The turbine blade 1 is divided into a plurality of stages of flow passages 3 by the longitudinal screen 2, and each stage of flow passage 3 is divided into a front chamber 5 and a rear chamber 6 by the transverse screen 4. The front and back heating surfaces of the turbine blade 1 along the rotation direction are respectively a suction surface 7 and a pressure surface 8. The longitudinal screen 2, the transverse screen 4 and the suction surface 7 enclose a front cabin 5, and the longitudinal screen 2, the transverse screen 4 and the pressure surface 8 enclose a rear cabin 6. Preferably, the transverse screen 4 is made of porous metal foam. The transverse screen 4 divides the flow channel 3 non-uniformly, the front compartment 5 being wider than the rear compartment 6. The front cabin 5 and the rear cabin 6 of the same-stage flow channel 3 are communicated at the blade top 9, and the front cabin 5 of the previous-stage flow channel 3 and the rear cabin 6 of the next-stage flow channel 3 are connected at the blade root 11 through a communication hole 10. The cross screen 4 in the first stage flow passage 3 is connected with a jet cavity 14 of a turbine blade leading edge 13 through a jet hole 12. The jet cavity 14 is provided with air film holes 15 on the suction surface 7 and the pressure surface 8.

When the divided cabin type turbine blade 1 rotates in the high-temperature gas 16, the gas-liquid two-phase flow cooling medium 17 is introduced from the cooling medium inlet 18 of the blade root 11. The gas-liquid two-phase flow cooling medium 17 firstly enters the rear cabin 6 of the primary flow channel 3, flows from the blade root 11 to the blade top 9 in the flow direction, and forms a thin liquid film 19 and a large velocity gradient on the pressure surface 8 under the action of Coriolis force, and forms a thick liquid film 20 on the cross screen 4. The gas-liquid two-phase flow cooling medium 17 enters the front cabin 5 after reaching the blade top 9, the flow direction is reversed, the cooling medium flows from the blade top 9 to the blade root 11, a thin liquid film 19 and a large velocity gradient are formed on the suction surface 7 under the action of Coriolis force, and a thick liquid film 20 is also formed on the cross screen 4. After reaching the blade root 11, the gas-liquid two-phase flow cooling medium 17 enters the front cavity 5 of the next-stage flow channel 3 through the communication hole 10. The formation of a thin liquid film 19 and a large velocity gradient on the pressure side 8 and on the suction side 7 contributes to an increased cooling of the turbine blade 1. And the thick liquid film 20 absorbed on the transverse screen 4 enters the incident flow cavity 14 through the jet hole 12, performs jet impact cooling on the blade front edge 13, and then is discharged through the air film hole 15 to perform air film cooling on the suction surface 7 and the pressure surface 8.

Example 1:

the turbine blade with the height of 200 mm, the length of 200 mm and the width of 20 mm is divided into three stages of flow channels by the longitudinal screen, and each stage of flow channel is 40 mm in length, 15 mm in width and 5 mm in thickness of the longitudinal screen. Each stage of flow channel is divided by a transverse screen into a front chamber adjacent to the suction surface and a back chamber adjacent to the pressure surface. The front cavity is 8 mm wide and the back cavity is 5 mm wide. The transverse screen adopts metal copper foam, the mesh number is 400PPI, and the thickness is 2 mm. The front cabin and the rear cabin of the same stage are communicated at the blade top, and the front cabin of the previous stage runner and the rear cabin of the next stage runner are connected at the blade root through a communicating hole. The communicating hole is U-shaped, the section is rectangular, the length is 20 mm, and the width is 5 mm. The transverse screen in the first-stage flow channel is connected with the jet cavity at the front edge of the turbine blade through the jet hole. The jet holes are arranged at equal intervals along the height of the turbine, the diameter of each jet hole is 1.5 mm, and the interval is 15 mm. The cross-sectional area of the fluidic chamber is 400 square millimeters. The jet cavity is provided with 10 air film holes respectively at equal intervals on a suction surface and a pressure surface. Each air film hole has a diameter of 1 mm and a distance of 15 mm. The section of the air film hole is in a water drop shape, and an included angle of 10 degrees is formed between the air film hole and the tangent direction of the molded line of the turbine blade.

When the turbine blade runs, a steam-water two-phase flow cooling medium is introduced from a cooling medium inlet with the diameter of 5 mm at the blade root, sequentially passes through the rear cavity and the front cavity of each stage of channel, part of the two-phase flow cooling medium is discharged from the air film hole, and most of the two-phase flow cooling medium is recovered from the blade top. The pressure surface and the suction surface in the flow passage of the turbine blade generate a thin liquid film with the magnitude of 10 microns. When the turbine blade is made of the ceramic matrix composite material, the gas temperature of the divided cabin type turbine blade can reach 1800 ℃ and above.

The embodiments are only preferred embodiments of the present invention, but the scope of the present invention is not limited thereto, and any changes or substitutions that can be easily conceived by those skilled in the art within the technical scope of the present invention are included in the scope of the present invention. Therefore, the protection scope of the present invention shall be subject to the protection scope of the claims.

Claims

1. A kind of divided cabin type turbine blade, the jet cavity of the blade leading edge opens the air film hole on suction surface and pressure surface of both sides of blade, characterized by that, set up multiple longitudinal screens in the blade along the direction from leading edge to trailing edge so as to cut apart and form the multistage flow path, each stage of flow path is cut apart into the top and communicates the front deck and rear deck separated at the bottom by the horizontal screen again, the front deck of the previous stage of flow path communicates with rear deck of the next stage of flow path at the root; the transverse screen is provided with a plurality of jet holes which are connected with a plurality of jet cavities on the front edge of the blade, the bottom of the rear cabin of the first-stage flow channel is connected with a cooling medium inlet at the root part of the blade, the front cabin is surrounded by the longitudinal screen and the suction surface, and the rear cabin is surrounded by the longitudinal screen and the pressure surface; the gas-liquid two-phase cooling medium in the rear cabin flows to the blade top from the blade root, and the gas-liquid two-phase cooling medium in the front cabin flows to the blade root from the blade top; the thick liquid film that adsorbs on horizontal screen gets into the efflux chamber through the efflux hole, and the jet impingement cooling is followed the discharge through the air film hole to blade leading edge, carries out the air film cooling to suction surface and pressure surface.

2. The divided nacelle turbine blade as claimed in claim 1, wherein the suction side and the pressure side are the front and rear sides of the blade in the direction of rotation, respectively.

3. The divided nacelle turbine blade of claim 1, wherein the forward nacelle width is greater than the aft nacelle width.

4. The divided nacelle turbine blade of claim 1, wherein the film holes are drop-shaped in cross-section and form an angle with the tangential direction of the blade profile.

5. The divided nacelle turbine blade as claimed in claim 1, wherein the cross screen is made of porous metal foam.

6. The divided nacelle turbine blade of claim 1, wherein the cooling medium is an air-water two-phase flow or a steam-water two-phase flow.

7. The divided cabin turbine blade according to any one of claims 1 to 6, wherein when the blade rotates, the gas-liquid two-phase cooling medium in the aft cabin flows from the blade root to the blade tip, and forms a thin liquid film and a large velocity gradient on the pressure surface under the action of Coriolis force; gas-liquid two-phase cooling medium in the front cabin flows from the blade top to the blade root, and a thin liquid film and a large velocity gradient are formed on the suction surface; the thick liquid film that adsorbs on horizontal screen gets into the efflux chamber through the efflux hole, and the jet impingement cooling is followed the discharge through the air film hole to blade leading edge, carries out the air film cooling to suction surface and pressure surface.