CN112283156B

CN112283156B - Gas compressor bleed structure and aeroengine

Info

Publication number: CN112283156B
Application number: CN202011588641.5A
Authority: CN
Inventors: 王进春; 郭玲; 言赞; 王家广; 曹传军; 李游; 陆晓锋; 吴志青
Original assignee: AECC Commercial Aircraft Engine Co Ltd
Current assignee: AECC Commercial Aircraft Engine Co Ltd
Priority date: 2020-12-29
Filing date: 2020-12-29
Publication date: 2021-03-19
Anticipated expiration: 2040-12-29
Also published as: CN112283156A

Abstract

The disclosure relates to a compressor bleed air structure and an aircraft engine. Wherein, compressor bleed structure includes: a two-stage rotor disk; two stages of rotor blades; the drum is provided with a plurality of air-entraining ports distributed along the circumferential direction; and the drainage pipes are arranged on the outer wall of the drum barrel in a circumferential distribution manner and are communicated with the air-entraining port in a one-to-one manner, the drainage pipes are arc pipes, and the offset direction of the stacking axis of the drainage pipes is configured to be consistent with the rotating direction of the rotor disc. The drainage tube can rectify the turbulent flow field near the air entraining port and smoothly guide the air flow to the air entraining port, so that the mechanical strength of the drainage structure is improved, the flow loss of the entrained air is effectively reduced, and the air entraining efficiency and the stable reliability of the entrained air are effectively improved.

Description

Gas compressor bleed structure and aeroengine

Technical Field

The disclosure relates to the technical field of gas turbines, in particular to a gas compressor air-entraining structure and an aircraft engine.

Background

In an aircraft engine compressor, a turbine stator component with a relatively high working temperature generally needs to be cooled by introducing cold air from a compressor end with a relatively low working temperature, due to the limitations of structural dimensions of an engine and the like, the introduced air from a main flow passage of the compressor mostly flows to the turbine end through a rotor cavity, and a hole is usually directly formed in a wall of a rotor drum to introduce the cooled air with certain pressure in the main flow passage of the compressor to the rotor cavity, but the following problems exist in the design technology:

1. the air-entraining ports on the drum wall are usually arranged in a through hole form and are distributed at intervals along the circumferential direction, and the air-entraining ports rotate along with the rotor, so that the air flow at the air-entraining ports has higher circumferential involving speed;

2. the drum walls are usually thin and it is difficult to arrange flow guiding devices on the drum walls.

Disclosure of Invention

The inventor researches and discovers that the air compressor air entraining structure in the related technology has the technical problem of low air entraining efficiency.

In view of this, the embodiment of the disclosure provides a compressor bleed air structure and an aircraft engine, which can effectively reduce bleed air loss and improve bleed air efficiency.

Some embodiments of the present disclosure provide a compressor bleed air structure, comprising:

two stages of rotor discs which are oppositely arranged;

the two-stage rotor blades are correspondingly arranged on the outer edge of the first-stage rotor wheel disc;

the drum barrel is arranged between the two stages of rotor discs, each end of the drum barrel is correspondingly connected with the first stage of rotor disc, and a plurality of air-entraining ports distributed along the circumferential direction are arranged on the drum barrel; and

the drainage tubes are circumferentially distributed on the outer wall of the drum, are communicated with the air-entraining ports one by one, and are used for guiding air flow and guiding the air flow to the air-entraining ports;

wherein, the drainage tube is an arc tube and the offset direction of the stacking axis of the drainage tube is configured to be consistent with the rotating direction of the rotor disc.

In some embodiments, the opening direction of the air inlet of the draft tube is at an angle of less than 90 ° to the direction of rotation of the rotor disk.

In some embodiments, the draft tube is in a diverging configuration along the direction of the bleed air.

In some embodiments, the cross-sectional shape of the draft tube is elliptical or circular.

In some embodiments, the bleed ports are of a diverging configuration along the bleed direction.

In some embodiments, the draft tube and bleed port are both located on a side of the rotor disk adjacent the subsequent stage.

In some embodiments, the rotor blade further comprises a stator blade located between the two stages of rotor blades, and the draft tube and the bleed port are both located between the stator blade and the subsequent stage of rotor blades at an axial position.

Some embodiments of the present disclosure provide an aircraft engine including the compressor bleed air structure described above.

Therefore, according to the embodiment of the disclosure, the drainage tubes distributed in the circumferential direction are arranged on the outer wall of the drum barrel, the drainage tubes are arc-shaped tubes, the offset direction of the stacking axis of the drainage tubes is configured to be consistent with the rotating direction of the rotor wheel disc, the turbulent flow field near the air-entraining port can be rectified, air flow can be smoothly guided to the air-entraining port, the mechanical strength of the drainage structure is improved, meanwhile, the air-entraining flow loss is effectively reduced, and the air-entraining efficiency and the air-entraining stability and reliability are effectively improved.

Drawings

The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the disclosure and together with the description, serve to explain the principles of the disclosure.

The present disclosure may be more clearly understood from the following detailed description, taken with reference to the accompanying drawings, in which:

FIG. 1 is a schematic structural view of some embodiments of a compressor bleed air arrangement of the present disclosure;

FIG. 2 is a cross-sectional view of the drum taken through the full ring at location A-A in FIG. 1;

FIG. 3 is a schematic velocity triangle;

FIG. 4 is a cross-sectional view taken at location B-B of FIG. 2;

fig. 5 is a cross-sectional view at the position C-C in fig. 2.

Description of the reference numerals

1. A rotor disk; 2. a rotor blade; 3. a drum; 4. a gas-introducing port; 5. a bleed air line; 6. a drainage tube; 7. stator blades.

Detailed Description

Various exemplary embodiments of the present disclosure will now be described in detail with reference to the accompanying drawings. The description of the exemplary embodiments is merely illustrative and is in no way intended to limit the disclosure, its application, or uses. The present disclosure may be embodied in many different forms and is not limited to the embodiments herein. These embodiments are provided so that this disclosure will be thorough and complete, and will fully convey the scope of the disclosure to those skilled in the art. It should be noted that: the relative arrangement of parts and steps, the composition of materials, numerical expressions and numerical values set forth in these embodiments are to be construed as merely illustrative, and not as limitative, unless specifically stated otherwise.

The use of "first," "second," and similar terms in this disclosure is not intended to indicate any order, quantity, or importance, but rather are used to distinguish one element from another. The word "comprising" or "comprises", and the like, means that the element preceding the word covers the element listed after the word, and does not exclude the possibility that other elements are also covered. "upper", "lower", "left", "right", and the like are used merely to indicate relative positional relationships, and when the absolute position of the object being described is changed, the relative positional relationships may also be changed accordingly.

In the present disclosure, when a specific device is described as being located between a first device and a second device, there may or may not be intervening devices between the specific device and the first device or the second device. When a particular device is described as being coupled to other devices, the particular device may be directly coupled to the other devices without intervening devices or may be directly coupled to the other devices with intervening devices.

All terms used in the present disclosure have the same meaning as understood by one of ordinary skill in the art to which the present disclosure belongs, unless otherwise specifically defined. It will be further understood that terms, such as those defined in commonly used dictionaries, should be interpreted as having a meaning that is consistent with their meaning in the context of the relevant art and will not be interpreted in an idealized or overly formal sense unless expressly so defined herein.

Techniques, methods, and apparatus known to those of ordinary skill in the relevant art may not be discussed in detail, but are intended to be part of the specification where appropriate.

As shown in fig. 1, some embodiments of the present disclosure provide for a compressor bleed air arrangement, comprising: the device comprises a two-stage rotor disk 1, two-stage rotor blades 2, a drum barrel 3 and a plurality of drainage tubes 6, wherein the two-stage rotor disk 1 is arranged oppositely; each stage of rotor blade 2 is correspondingly arranged on the outer edge of the first stage of rotor disk 1; the drum barrel 3 is arranged between the two stages of rotor discs 1, each end of the drum barrel is correspondingly connected with the one stage of rotor disc 1, and a plurality of air-entraining ports 4 distributed along the circumferential direction are arranged on the drum barrel; a bleed air line 5 of the turbine is installed in the drum 3, so that the high-pressure air of the compressor is led from a bleed air port of the compressor to the bleed air line 5 to cool the turbine. As shown in fig. 2, a plurality of draft tubes 6 are circumferentially distributed on the outer wall of the drum 3 and communicate with the bleed air ports 4 one-to-one for guiding the air flow to the bleed air ports 4. Wherein the draft tube 6 is an arc tube and the offset direction of the stacking axis thereof is configured to coincide with the rotation direction of the rotor disk 1. The curved arrows in figure 2 show the direction of rotation of the rotor disc 1. The stacking axis of the draft tube means a connecting line of the draft tube 6 from the drum 3 to the center of each section of the tail end of the draft tube 6, and the tail end of the draft tube 6 is provided with an air inlet.

In the exemplary embodiment, the drainage tubes 6 are circumferentially arranged on the outer wall of the drum 3, the drainage tubes 6 are arc-shaped tubes, and the offset direction of the stacking axis of the drainage tubes 6 is configured to be consistent with the rotation direction of the rotor disk 1, that is, the flow path of the drainage tubes 6 is offset in the stacking process from the drum 3 to the tail end of the drainage tube 6 toward the rotation direction of the rotor disk 1, so that the turbulent flow field near the air entraining port 4 can be rectified, and the air flow can be smoothly guided to the air entraining port 4, which is beneficial to improving the air entraining quality. Moreover, the draft tube 6 is easily installed on the outer wall of the drum 3, and has high feasibility of implementation.

In some embodiments, as shown in fig. 1, the draft tube 6 and the bleed port 4 are both located near one side of the subsequent rotor disk 1 to facilitate bleeding air.

In particular, in some embodiments, as shown in fig. 1, the compressor bleed air structure further includes a stator blade 7 located between the two stages of rotor blades 2, the draft tube 6 and the bleed air port 4 are both located between the stator blade 7 and the subsequent stage of rotor blade 2 in the axial position, and the draft tube 6 can effectively reduce the bleed air loss at the rotor-stator interface, improve the bleed air efficiency, and have high implementability.

In some embodiments, as shown in fig. 1, the opening direction of the air inlet of the draft tube 6 makes an angle of less than 90 ° with the rotation direction of the rotor disc 1. As shown in fig. 2 and 3, w is the relative velocity of the air flow, u is the bulk velocity of the air flow, and V is the radial velocity of the air flow, and based on the velocity triangle principle, the draft tube 6 can relatively smoothly rectify the turbulent flow field near the bleed port 4 and lead to the bleed port 4.

As shown in FIGS. 4 and 5, in some embodiments, the draft tube 6 is of a diverging configuration in the direction of the bleed air. The cross section of the tail end of the drainage tube 6 is vertical to the relative speed w of the air flow, r is the radial position of the center of the cross section of the tail end of the drainage tube 6, and omega is the angular speed of the drum barrel 3. Here, the approximation assumes that the absolute velocity of the gas flow is in the radial direction, and the bulk velocity is calculated as: u = ω r.

As shown in FIGS. 4 and 5, the flow area of draft tube 6 at section C-C is larger than that at section B-B, and the flow of air through the expanded draft tube 6 is decelerated and pressurized, thereby facilitating the increase in the pressure of the bleed air. Similarly, in some embodiments, as shown in fig. 2, the bleed ports 4 are of a diverging configuration along the bleed direction.

To ensure smooth drainage of the airflow, in some embodiments, as shown in fig. 4 and 5, the cross-sectional shape of the draft tube 6 is elliptical or circular in some embodiments.

To further boost the bleed air pressure, in some embodiments the bleed air ports 4 are of a diverging configuration in the bleed air direction. The air current can slow down the pressure boost through expanded bleed port 4 to be favorable to improving bleed pressure, further promoted bleed efficiency.

The drainage tube 6 can be processed into a whole with the drum barrel 3 through machining, and can also be made into a split structure.

Some embodiments of the disclosure provide that an aircraft engine comprises the compressor bleed air structure. The aircraft engine accordingly has the beneficial technical effects.

Thus, various embodiments of the present disclosure have been described in detail. Some details that are well known in the art have not been described in order to avoid obscuring the concepts of the present disclosure. It will be fully apparent to those skilled in the art from the foregoing description how to practice the presently disclosed embodiments.

Although some specific embodiments of the present disclosure have been described in detail by way of example, it should be understood by those skilled in the art that the foregoing examples are for purposes of illustration only and are not intended to limit the scope of the present disclosure. It will be understood by those skilled in the art that various changes may be made in the above embodiments or equivalents may be substituted for elements thereof without departing from the scope and spirit of the present disclosure. The scope of the present disclosure is defined by the appended claims.

Claims

1. A compressor bleed air structure, comprising:

the two-stage rotor wheel disc (1) is oppositely arranged;

the two-stage rotor blade (2), each stage of rotor blade (2) is correspondingly arranged on the outer edge of the rotor wheel disc (1);

the drum barrel (3) is arranged between the two stages of the rotor wheel discs (1), each end of the drum barrel is correspondingly connected with the one stage of the rotor wheel discs (1), and a plurality of air-entraining ports (4) distributed along the circumferential direction are arranged on the drum barrel;

stator blades (7) located between the two stages of rotor blades (2); and

the drainage tubes (6) are circumferentially distributed on the outer wall of the drum (3), are communicated with the air-introducing ports (4) one by one, and are used for guiding airflow and guiding the airflow to the air-introducing ports (4);

wherein the drainage tube (6) is an arc-shaped tube and the offset direction of the stacking axis of the drainage tube is configured to be consistent with the rotating direction of the rotor disc (1); the drainage tube (6) and the air-entraining port (4) are both positioned between the stator blade (7) and the rotor blade (2) at the next stage in the axial position.

2. Compressor bleed air structure according to claim 1, characterised in that the opening direction of the air inlet of the draft tube (6) forms an angle of less than 90 ° with the direction of rotation of the rotor disc (1).

3. The compressor air-entraining structure according to claim 1, characterised in that the draft tube (6) is of a diverging structure in the air-entraining direction.

4. The compressor air guide structure according to claim 1, characterised in that the cross-sectional shape of the draft tube (6) is oval or circular.

5. The compressor bleed air arrangement according to claim 1, characterised in that the bleed air openings (4) are of a divergent configuration in the bleed air direction.

6. The compressor air-entraining structure according to claim 1, characterized in that the draft tube (6) and the air-entraining port (4) are both located at a side close to the rotor disk (1) of the subsequent stage.

7. An aircraft engine, characterized by comprising the air entraining structure of the compressor in any one of claims 1 to 6.