CN112283155B

CN112283155B - Gas compressor bleed structure and aeroengine

Info

Publication number: CN112283155B
Application number: CN202011588628.XA
Authority: CN
Inventors: 王家广; 王进春; 严冬青; 曹传军; 李游; 况成玉; 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: CN112283155A

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; the diversion containing shells are circumferentially distributed on the outer wall of the drum barrel and are positioned at the same axial position with the air bleed port; the diversion containing shells are arranged on the upstream sides of the air guide ports in the rotating direction of the rotor wheel disc in a one-to-one mode, are of one-way opening structures, and are configured to be consistent with the rotating direction of the rotor wheel disc in the opening direction. The diversion containing shell can rectify a turbulent flow field near the air entraining port and smoothly introduce air flow into the air entraining port, so that the mechanical strength of the drainage structure is improved, the flow loss of air entraining is effectively reduced, and the air entraining efficiency and the air entraining stability and reliability 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 positioned at the interstage concave cavity, and the flow field at the position is greatly disordered and has low air-entraining efficiency under the influence of the position and the shape of the air-entraining ports;

2. usually, the bleed ports on the drum wall are circumferentially spaced and rotate with the rotor, and although the air flow at the bleed ports has a certain circumferential speed, the circumferential speed is lower than the circumferential rotation speed of the bleed ports, so that the bleed air flow loss is further aggravated when the bleed air flow flows to the bleed ports with higher relative movement speed, and the bleed efficiency is reduced.

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 flow guide containing shells are circumferentially distributed on the outer wall of the drum barrel, are positioned at the same axial position with the air introducing port and are used for guiding air flow and guiding the air flow to the air introducing port;

the diversion containing shells are arranged on the upstream sides of the air guide ports in the rotating direction of the rotor wheel disc in a one-to-one mode, are of one-way opening structures, and are configured to be consistent with the rotating direction of the rotor wheel disc in the opening direction.

In some embodiments, the flow guide housing is an arc-shaped housing.

In some embodiments, the pod is an 1/4 spherical shell.

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

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

In some embodiments, the air guide receiving shell and the air guide port are located between the stator blade and the rotor blade of the next stage in the 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 flow guide accommodating shell which is circumferentially distributed is arranged on the outer wall of the drum barrel, the flow guide accommodating shell is of a one-way opening structure, the opening direction of the flow guide accommodating shell is configured to be consistent with the rotating direction of the rotor wheel disc, a turbulent flow field near the air entraining port can be rectified, air flow is smoothly introduced into the air entraining port, the mechanical strength of the flow guide 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 schematic perspective view of a drum in some embodiments of the compressor bleed air configuration of the present disclosure;

fig. 3 is a cross-sectional view of the drum taken through the full circle at a-a in fig. 1.

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 diversion housing shell; 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 a compressor bleed air structure, including: the device comprises a two-stage rotor disk 1, two-stage rotor blades 2, a drum barrel 3 and a plurality of flow guide containing shells 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; an air bleed pipeline 5 of a turbine is arranged in the drum 3, so that high-pressure air of the air compressor is led to the air bleed pipeline 5 from an air bleed port of the air compressor to cool the turbine, and a plurality of diversion containing shells 6 are circumferentially distributed on the outer wall of the drum 3, are positioned at the same axial position with the air bleed port 4 and are used for guiding air flow and leading the air flow to the air bleed port 4; as shown in fig. 2 and 3, the storage cases 6 are provided one for one on the upstream side of the bleed ports 4 in the rotation direction of the rotor disk 1, and the arc-shaped arrows in fig. 3 show the rotation direction of the rotor disk 1. The surge tank 6 has a one-way opening structure and an opening direction thereof is arranged to coincide with a rotation direction of the rotor disk 1.

In the exemplary embodiment, as shown in fig. 2 and fig. 3, the circumferentially distributed diversion housing shells 6 are disposed on the outer wall of the drum 3, the diversion housing shells 6 are of a one-way opening structure, and the opening direction of the diversion housing shells 6 is configured to be consistent with the rotation direction of the rotor disk 1, so that dynamic pressure of air flow having a certain circumferential speed relative to the diversion housing shells 6 can be converted into static pressure, a turbulent flow field near the air introducing port 4 can be rectified, and the air flow can be smoothly introduced to the air introducing port 4, which is beneficial to improving air introducing quality.

As shown in fig. 3, the upstream side in the rotation direction of the rotor disc 1 means a side located on the rear side in the direction along the rotation of the rotor disc 1, and correspondingly, the downstream side is a side located on the front side.

In some embodiments, as shown in fig. 1, the diversion housing 6 and the bleed port 4 are both located at a side close to the rear-stage rotor disk 1, so as to bleed 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, and the flow guide housing 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 flow guide housing 6 can effectively reduce the bleed air loss at the rotor-stator interface, improve the bleed air efficiency, and has high practicability.

In some embodiments, as shown in fig. 2 and 3, the diversion housing 6 is an arc-shaped housing, which effectively reduces airflow loss. In some embodiments, as shown in fig. 2 and 3, the diversion housing 6 is an 1/4 spherical shell, which is convenient for processing and effectively increases the diversion area.

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 diversion containing shell 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 present disclosure provide an aircraft engine including the compressor bleed air structure described above. 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; and

the flow guide containing shells (6) are circumferentially distributed on the outer wall of the drum barrel (3), are positioned at the same axial position with the air introducing port (4), and are used for guiding air flow and guiding the air flow to the air introducing port (4);

the air guide containing shell (6) is arranged on the upstream side of the air guide ports (4) in the rotating direction of the rotor disc (1) in a one-to-one mode, the air guide containing shell (6) is of a one-way opening structure, the opening direction of the air guide containing shell is configured to be consistent with the rotating direction of the rotor disc (1), and the air guide containing shell (6) is an 1/4 spherical shell.

2. 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.

3. The compressor air-entraining structure according to claim 1, characterized in that the air-entraining housing (6) and the air-entraining opening (4) are both located on one side close to the rotor disk (1) of the subsequent stage.

4. The compressor bleed air arrangement according to claim 1, characterised by further comprising a stator blade (7) located between the two stages of rotor blades (2), the guide housing (6) and the bleed air opening (4) being located in the axial position between the stator blade (7) and the rotor blade (2) of the subsequent stage.

5. An aircraft engine, characterized by comprising the air entraining structure of the compressor of any one of claims 1 to 4.