CN112922681A

CN112922681A - Aeroengine rim sealing structure

Info

Publication number: CN112922681A
Application number: CN202110307748.6A
Authority: CN
Inventors: 王家友; 段玉发; 范宇; 邓明春; 周建军
Original assignee: AECC Shenyang Engine Research Institute
Current assignee: AECC Shenyang Engine Research Institute
Priority date: 2021-03-23
Filing date: 2021-03-23
Publication date: 2021-06-08

Abstract

The application provides an aeroengine rim structure of obturating includes: a guide blade; a rotor blade; the front flange sealing structure is provided with a first backward protrusion and a second backward protrusion which extend along the direction of a main airflow flow channel, the first backward protrusion and the second backward protrusion form a first groove, a flow deflector is arranged on the upper wall of the first groove, the rear flange sealing structure is provided with a first forward protrusion which extends in the direction opposite to the direction of the main airflow flow channel, and at least part of the first forward protrusion extends into the first groove. This application on the one hand can increase the flow resistance of hot gas flow direction gap in engine operation process to promote the effect of obturating, on the other hand can reduce the mixing loss of gas and air conditioning, and make air conditioning do work to the turbine, promote the efficiency of doing work of gas, improve turbine efficiency, reduce the engine oil consumption rate.

Description

Aeroengine rim sealing structure

Technical Field

The application belongs to the technical field of aeroengines, in particular to an aeroengine rim sealing structure.

Background

During the operation of the aircraft engine, high-temperature combustion gas in the main flow passage of the turbine enters the turbine disk cavity from a gap between the rotor and the stator of the turbine, and the phenomenon is called gas invasion. The invasion of the fuel gas can increase the temperature of the disk cavity, so that the temperature of the turbine disk is overheated, the temperature of the bearing cavity is overhigh, potential safety hazards are brought to an engine, and the service life of the engine is influenced.

To prevent the ingress of combustion gases, cold air is typically supplied to the turbine disk cavity and discharged into the main flowpath through the gap between the turbine rotor and stator. On the basis, a rim sealing structure is usually arranged on the outer diameter of the turbine disc to increase the flow resistance of the invading gas. These two measures are usually used in combination in an aircraft engine.

However, the existing turbine rim sealing structure still has small sealing resistance, which results in large flow of sealing cold air, and further reduces the turbine efficiency. In addition, the cold air for preventing the gas from invading directly blows into the main runner of the turbine in the radial direction and is mixed with the gas flowing in the axial direction, obvious mixing loss is generated, the working efficiency of the gas is reduced, and the efficiency of the turbine is reduced. This results in reduced engine efficiency and increased fuel consumption.

Disclosure of Invention

It is an object of the present application to provide an aircraft engine rim seal structure to address or mitigate at least one of the problems of the background art.

In one aspect, the application provides an aeroengine rim structure of obturating, includes:

a guide blade;

a rotor blade; and

the structure comprises a front flange sealing structure arranged on the guide blade and a rear flange sealing structure arranged on the rotor blade, wherein the front flange sealing structure is provided with a first backward protrusion and a second backward protrusion extending along the direction of a main airflow flow channel, the first backward protrusion and the second backward protrusion form a first groove, a flow deflector is arranged on the upper wall of the first groove, the rear flange sealing structure is provided with a first forward protrusion extending opposite to the direction of the main airflow flow channel, and at least part of the first forward protrusion extends into the first groove.

Further, the first aft projection is further from the aircraft engine axis than the second aft projection.

Further, the rim pre-sealing structure is also provided with a third backward protrusion close to the axis of the aircraft engine compared with the second backward protrusion, and the third backward protrusion and the second backward protrusion form a second groove.

Further, the first groove and the second groove are both rectangular in shape.

Furthermore, the guide vanes are uniformly distributed along the axis of the aircraft engine.

Further, the guide vane and the axial direction of the aircraft engine form a preset angle.

Further, the predetermined angle is 15 to 35 degrees.

Further, the first forward projection extends into the first recess no more than 1/2 times the depth of the first recess.

In another aspect, the application further provides an aircraft engine comprising an aircraft engine rim seal structure as described in any one of the above.

The application provides an aeroengine rim structure of obturating is through setting up the water conservancy diversion piece in first recess, can increase the flow resistance of hot gas flow direction gap in the engine operation process on the one hand to promote the effect of obturating, on the other hand can change direction and the speed that air conditioning got into the turbine main entrance, reduce the mixing loss of gas and air conditioning, and make air conditioning do work to the turbine, promote the work efficiency of gas, improve turbine efficiency, reduce the engine oil consumption rate.

Drawings

In order to more clearly illustrate the technical solutions provided by the present application, the following briefly introduces the accompanying drawings. It is to be expressly understood that the drawings described below are only illustrative of some embodiments of the invention.

Fig. 1 is a schematic view of a rim sealing structure of the present application.

Fig. 2 is a schematic view of a structure and an installation form of a baffle according to the present application.

Fig. 3 is a three-dimensional schematic view of the structure of the present application before rim sealing.

FIG. 4 is a schematic diagram of a three-dimensional matching state of a front rim sealing structure and a rear rim sealing structure of the present application.

Reference numerals:

1-a turbine case;

2-guide vanes;

3-rotor blades;

4-a turbine disc, wherein the turbine disc is provided with a plurality of blades,

5-structure before sealing of rim, 51-first backward projection, 52-second backward projection, 53-third backward projection, and 54-deflector;

6-sealing the rear structure of the wheel rim, 61-protruding forwards.

Detailed Description

In order to make the implementation objects, technical solutions and advantages of the present application clearer, the technical solutions in the embodiments of the present application will be described in more detail below with reference to the drawings in the embodiments of the present application.

In order to effectively promote the effect of sealing of rim, reduce the mixing loss of gas and air conditioning, make the cooling gas do work to the turbine, promote turbine efficiency, reduce the engine oil consumption rate, provide a rim structure of sealing with circumference water conservancy diversion function in this application.

As shown in fig. 1 to 4, the present application specifically provides an aeroengine rim sealing structure, including: guide vane 2, rotor blade 3 and install preceding structure 5 of rim seal on guide vane 2 and install the rim structure 6 of rim seal on rotor blade 3, preceding structure 5 of rim seal has the first backward arch 51 and the second backward arch 52 that extend along the air current direction in the mainstream way (extend to the right side in fig. 1 promptly), first backward arch 51 and the second backward arch 52 constitute first recess, be equipped with water conservancy diversion piece 54 at the upper wall that first recess is close to the mainstream way position, the rim structure 6 of rim seal has the first preceding arch 61 that extends opposite to main air current runner direction, first preceding arch 61 at least partially stretches into in the first recess.

Through increase the water conservancy diversion piece 54 on rim seal preceding structure 5, can increase the high temperature gas resistance of obturating at turbine wheel rim department, reduce the air conditioning quantity of obturating to and can change the flow direction that the air conditioning flowed into the main entrance, reduce the mixing loss of gas and air conditioning, simultaneously because the effect of water conservancy diversion piece 54, after air conditioning got into the main entrance, flow along turbine rotor rotation direction, do work to the turbine, promote turbine efficiency. The flow directions of the high-temperature fuel gas and the sealed cold air are shown in fig. 2, wherein, the dotted line arrow represents the flow direction of the fuel gas in the main flow channel, the solid line arrow represents the flow direction of the sealed cold air of the disc cavity,

in the above-described embodiments of the present application, the first rearward projection 51 is closer to the primary flowpath than the second rearward projection 52, and therefore it is further away from the aircraft engine axis.

In the further preferred embodiment described above, the rim pre-sealing structure 5 also has a third rearward projection 53, the third rearward projection 53 being closer to the engine axis than the second rearward projection 52, the third rearward projection 53 and the second rearward projection 52 thereby constituting a second groove.

In the present application, the first groove and the second groove are both rectangular or approximately rectangular in shape.

In the present application, the deflector 54 provided on the upper wall of the first recess is generally in plurality, which is equispaced about the axis of the aircraft engine. In the preferred embodiment described above, the guide vanes 54 are at a predetermined angle to the axial direction of the aircraft engine. The predetermined angle is usually determined according to the sealing effect, and according to the test result, the sealing effect is better when the predetermined angle is 15-35 degrees.

In addition, the first distance that stretches into in the first recess of protruding 61 also can influence the effect of obturating to protruding, and it is usually not more than 1/2 of first recess degree of depth, stretches into too deeply, takes place the rotor easily and bumps the mill, stretches into too shallowly, and the gas invasion easily takes place for the hot gas of sprue.

In addition, the application claims an aeroengine, the aeroengine comprises the aeroengine rim sealing structure. Specifically, the aero-engine comprises a casing 1, a guide blade 2 is fixedly installed on the casing 1, a rotor blade 3 is installed on a rotating shaft (not shown), a certain radially extending gap is formed between the guide blade 2 and the rotor blade 3, and the rim sealing front structure 5 and the rim sealing rear structure 6 are respectively installed on the guide blade 2 and the rotor blade 3 to seal the radial gap.

According to the rim sealing structure, the flow resistance of a channel between the rotor and the stator is increased through the flow deflector positioned on the cold air side of the inner ring of the turbine stator, and the sealing effect on gas is improved; the flow deflector changes the flow direction of the sealed cold air from the disc cavity, so that the sealed cold air enters the turbine main channel along the disc edge, and the mixing loss of the turbine main channel is reduced; due to the flow guide effect of the flow guide sheet, the sealed cold air enters the main turbine channel along the rotation direction of the turbine rotor, the work can be done on the turbine, the work efficiency of the gas can be effectively improved, the turbine efficiency is improved, and the oil consumption rate of the engine is reduced.

The above description is only for the specific embodiments of the present application, but the scope of the present application 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 application should be covered within the scope of the present application. Therefore, the protection scope of the present application shall be subject to the protection scope of the claims.

Claims

1. An aeroengine rim structure of obturating, comprising:

a guide blade (2);

a rotor blade (3); and

the rotor blade sealing structure comprises a rim sealing front structure (5) installed on a guide blade (2) and a rim sealing rear structure (6) installed on a rotor blade (3), wherein the rim sealing front structure (5) is provided with a first rear protrusion (51) and a second rear protrusion (52) extending along the direction of a main airflow flow channel, the first rear protrusion (51) and the second rear protrusion (52) form a first groove, a flow deflector (54) is arranged on the upper wall of the first groove, the rim sealing rear structure (6) is provided with a first front protrusion (61) extending opposite to the direction of the main airflow flow channel, and the first front protrusion (61) at least partially extends into the first groove.

2. An aeroengine rim seal structure according to claim 1, wherein the first rearward projection (51) is further from the aeroengine axis than the second rearward projection (52).

3. An aeroengine rim seal according to claim 2, wherein the rim seal front structure (5) further has a third rearward projection (53) closer to the engine axis than the second rearward projection (52), the third rearward projection (53) and the second rearward projection (52) forming a second groove.

4. An aircraft engine rim seal according to any one of claims 1 to 3, wherein the first and second grooves are both rectangular in shape.

5. An aeroengine rim seal structure according to claim 1, wherein the deflectors (54) are equispaced about the axis of the aeroengine.

6. An aircraft engine rim seal according to claim 5, characterised in that the guide vanes (54) are at a predetermined angle to the aircraft engine axial direction.

7. An aircraft engine rim seal according to claim 6, wherein the predetermined angle is 15 to 35 degrees.

8. An aircraft engine rim seal according to claim 1, characterised in that the first forward projection (61) projects into the first groove a distance not exceeding 1/2 of the first groove depth.

9. An aircraft engine, characterized in that it comprises an aircraft engine rim seal structure according to any one of claims 1 to 8.