CN113389755B - Compressor of gas turbine, gas turbine and aircraft - Google Patents
Compressor of gas turbine, gas turbine and aircraft Download PDFInfo
- Publication number
- CN113389755B CN113389755B CN202110941140.9A CN202110941140A CN113389755B CN 113389755 B CN113389755 B CN 113389755B CN 202110941140 A CN202110941140 A CN 202110941140A CN 113389755 B CN113389755 B CN 113389755B
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- Prior art keywords
- compressor
- gas turbine
- compression
- compression flow
- rotating shaft
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- 230000006835 compression Effects 0.000 claims abstract description 72
- 238000007906 compression Methods 0.000 claims abstract description 72
- 238000002955 isolation Methods 0.000 claims description 13
- 238000005192 partition Methods 0.000 claims description 10
- 238000007789 sealing Methods 0.000 claims description 6
- 125000006850 spacer group Chemical group 0.000 claims description 6
- 230000004323 axial length Effects 0.000 abstract description 9
- 238000009413 insulation Methods 0.000 description 3
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
Images
Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/40—Casings; Connections of working fluid
- F04D29/52—Casings; Connections of working fluid for axial pumps
- F04D29/522—Casings; Connections of working fluid for axial pumps especially adapted for elastic fluid pumps
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/08—Sealings
- F04D29/083—Sealings especially adapted for elastic fluid pumps
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/26—Rotors specially for elastic fluids
- F04D29/32—Rotors specially for elastic fluids for axial flow pumps
- F04D29/38—Blades
- F04D29/388—Blades characterised by construction
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/40—Casings; Connections of working fluid
- F04D29/52—Casings; Connections of working fluid for axial pumps
- F04D29/54—Fluid-guiding means, e.g. diffusers
- F04D29/541—Specially adapted for elastic fluid pumps
- F04D29/542—Bladed diffusers
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
The invention relates to a compressor of a gas turbine, the gas turbine and an airplane, the compressor of the gas turbine comprises: a housing (1); the rotating shaft (6) is rotatably arranged in the shell (1); and the compression flow passages are arranged between the rotating shaft (6) and the shell (1) in a side-by-side mode in the radial direction and are sequentially communicated in the gas flow direction, a plurality of stator blades (2) and a plurality of rotor blades (3) are arranged in each compression flow passage in a side-by-side mode in the gas flow direction, and the stator blades (2) and the rotor blades (3) are arranged alternately. A plurality of compression flow passages which are arranged side by side along the radial direction and are sequentially communicated along the airflow flowing direction are formed in the space between the shell (1) and the rotating shaft (6), so that the path of the compression flow passages is effectively increased, the axial length of the compressor is favorably reduced, and the ratio of the axial length to the radial size of the compressor is reduced.
Description
Technical Field
The invention relates to the technical field of aviation, in particular to a compressor of a gas turbine, the gas turbine and an airplane.
Background
In order to realize a high compression ratio of a compressor of a gas turbine (such as an aircraft engine), a plurality of stages of compressors are generally required to be designed for air flow compression, and further, the axial length of the compressor is long, so that the space utilization rate of the compressor is low due to the arrangement. Meanwhile, the ratio of the axial length to the radial size of the compressor is large, and the rigidity of parts is influenced. Therefore, the design of the compressor with good space utilization rate and good rigidity is considered.
Disclosure of Invention
The invention aims to provide a compressor of a gas turbine, the gas turbine and an airplane so as to solve the problem that the ratio of the axial length to the radial dimension of the compressor is large in the related art.
According to an aspect of an embodiment of the present invention, there is provided a compressor of a gas turbine, the compressor of the gas turbine including: a housing; the rotating shaft is rotatably arranged in the shell; and a plurality of compression flow passages which are arranged between the rotating shaft and the shell side by side along the radial direction and are sequentially communicated along the gas flowing direction, a plurality of stator blades and a plurality of rotor blades are arranged in each compression flow passage side by side along the gas flowing direction, and the stator blades and the rotor blades are alternately arranged.
In some embodiments, the compression flow path extends in an axial direction.
In some embodiments, the plurality of compression runners form a runner having an S-shaped longitudinal cross-section.
In some embodiments, the plurality of compression runners includes: a first compression channel including an inlet at a first axial end of the shaft and an outlet at a second axial end; the second compression flow passage comprises an outlet positioned at the first axial end and an inlet positioned at the second axial end, and the inlet of the second compression flow passage is communicated with the outlet of the first compression flow passage; and a third compression flow passage including an inlet at the first axial end and an outlet at the second axial end, the inlet of the third compression flow passage being in communication with the outlet of the second compression flow passage.
In some embodiments, the compression flow passage is an annular flow passage coaxial with the axis of rotation.
In some embodiments, the stator blades are connected to the housing and extend radially toward the shaft; and the rotor blades are connected with the rotating shaft and extend towards the shell along the radial direction.
In some embodiments, the compressor of the gas turbine further comprises an isolation component comprising: a first partition member connected to the stator vane and extending toward the rotor vane adjacent to the stator vane to divide a space between the rotary shaft and the housing into different compression flow passages arranged side by side in a radial direction; and/or a second partition member connected to the rotor blade and extending toward the stator blade adjacent to the rotor blade to divide a space between the rotary shaft and the housing into different compression flow passages arranged side by side in a radial direction.
In some embodiments, the isolation member further comprises a sealing member coupled between the first isolation member and the second isolation member.
In some embodiments, the first isolation member is annular and coaxial with the axis of rotation; the second isolation component is annular and coaxial with the rotating shaft.
In some embodiments, the compressor of the gas turbine comprises a compressor of an aircraft engine.
According to another aspect of the invention, a gas turbine is also provided, and the gas turbine comprises the compressor of the gas turbine.
According to another aspect of the invention, there is also provided an aircraft comprising a compressor of a gas turbine as described above.
By applying the technical scheme of the invention, a plurality of compression flow channels which are arranged side by side along the radial direction and are sequentially communicated along the airflow flowing direction are formed in the space between the shell and the rotating shaft, so that the path of the compression flow channels is effectively increased, the axial length of the compressor is favorably reduced, and the ratio of the axial length to the radial size of the compressor is reduced.
Other features of the present invention and advantages thereof will become apparent from the following detailed description of exemplary embodiments thereof, which proceeds with reference to the accompanying drawings.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the related art, the drawings needed to be used in the description of the embodiments or the related art will be briefly introduced below, it is obvious that the drawings in the following description are only some embodiments of the present invention, and for those skilled in the art, other drawings can be obtained according to these drawings without creative efforts.
Fig. 1 shows a schematic structural view of a longitudinal section of a compressor of a gas turbine according to an embodiment of the present invention.
Fig. 2 is a schematic structural view showing a longitudinal section of a compressor casing and stator blades of a gas turbine according to an embodiment of the present invention.
Fig. 3 is a schematic perspective view of a compressor casing and stator blades of a gas turbine according to an embodiment of the present invention.
Fig. 4 is a side view schematically showing a compressor casing and stator blades of a gas turbine according to an embodiment of the present invention.
Fig. 5 shows a schematic structural view of a longitudinal section of a rotor shaft and a rotor blade of a compressor of a gas turbine according to an embodiment of the present invention.
Fig. 6 is a perspective view showing a structure of a rotor shaft and a rotor blade of a compressor of a gas turbine according to an embodiment of the present invention.
In the figure:
1. a housing; 2. a stator blade; 3. a rotor blade; 4. a bearing; 5. a sealing member; 6. a rotating shaft; 7. a first isolation member; 8. a second isolation member; 10. A first compression flow path; 20. a second compression flow path; 30. and a third compression flow path.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. The following description of at least one exemplary embodiment is merely illustrative in nature and is in no way intended to limit the invention, its application, or uses. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
Referring to fig. 1 to 6, a compressor of a gas turbine according to the present embodiment includes a casing 1, a rotating shaft 6, and a plurality of compression flow passages. The rotating shaft 6 is rotatably arranged in the shell 1; the plurality of compression flow passages are arranged between the rotating shaft 6 and the shell 1 in parallel along the radial direction and are sequentially communicated along the gas flow direction, a plurality of stator blades 2 and a plurality of rotor blades 3 are arranged in each compression flow passage in parallel along the gas flow direction, and the stator blades 2 and the rotor blades 3 are alternately arranged.
In the embodiment, a plurality of compression flow passages which are arranged side by side in the radial direction and are sequentially communicated in the airflow flowing direction are formed in the space between the shell 1 and the rotating shaft 6, so that the path of the compression flow passages is effectively increased, the axial length of the compressor is favorably reduced, and the ratio of the axial length to the radial size of the compressor is reduced.
In this embodiment, the compression flow passages all extend along the axial direction, the plurality of compression flow passages are arranged side by side along the radial direction, and the plurality of compression flow passages are sequentially communicated end to form a reciprocating circuitous compression flow passage. The plurality of compression flow passages form a flow passage with an S-shaped longitudinal section.
As shown in fig. 1, the plurality of compression flow passages includes a first compression flow passage 10, a second compression flow passage 10, and a third compression flow passage 10.
The first compression flow passage 10 includes an inlet at a first axial end of the rotary shaft 6 and an outlet at a second axial end; the second compression flow passage 20 includes an outlet at a first end in the axial direction and an inlet at a second end in the axial direction, the inlet of the second compression flow passage 20 being communicated with the outlet of the first compression flow passage 10; the third compression flow passage 30 includes an inlet at a first end in the axial direction and an outlet at a second end in the axial direction, and the inlet of the third compression flow passage 30 communicates with the outlet of the second compression flow passage 20.
In the present embodiment, the compression flow passage is an annular flow passage coaxial with the rotary shaft 6. The compressor also comprises a bearing 4 for bearing a rotating shaft 6, and two ends of the rotating shaft 6 are respectively provided with the bearing 4.
The stator blades 2 are connected with the shell 1 and extend towards the rotating shaft 6 along the radial direction; the rotor blades 3 are connected to the rotor shaft 6 and extend radially towards the housing 1.
The compressor of the gas turbine also comprises an insulation part comprising a first insulation part 7 and a second insulation part 8.
The first partition member 7 is connected to the stator blade 2 and extends toward the rotor blade 3 adjacent to the stator blade 2 to divide a space between the rotating shaft 6 and the casing 1 into different compression flow passages arranged side by side in a radial direction.
The second partition member 8 is connected to the rotor blade 3 and extends toward the stator blade 2 adjacent to the rotor blade 3 to divide a space between the rotary shaft 6 and the casing 1 into different compression flow passages arranged side by side in a radial direction.
In this embodiment, the compressor includes two layers of partition members arranged in a radial direction to divide a space between the rotary shaft 6 and the casing 1 into first, second, and third compression flow passages.
The spacer member further comprises a sealing member 5 connected between the first spacer member 7 and the second spacer member 8. The seal member 5 is connected to one of the first partition member 7 and the second partition member 8 and is in sliding engagement with the other to prevent communication between radially adjacent compression flow passages when the rotor blade 3 is rotating.
The first isolation component 7 is annular and coaxial with the rotating shaft 6; the second partition member 8 has an annular shape coaxial with the rotary shaft 6, thereby dividing an annular space between the rotary shaft 6 and the casing 1 into a first compression flow passage 10, a second compression flow passage 20, and a third compression flow passage 30, each of which has an annular shape.
In some embodiments, the compressor of the gas turbine comprises a compressor of an aircraft engine.
In the present embodiment, the rotor blades 3 rotate to apply work to the air flow to compress the air; and flowing the gas; the stator blade 2 changes the incoming flow of the rotor blade 3, and adjusts the incoming flow angle to better pass through the next-stage rotor; the housing 1 guides the air flow in the direction of the arrow.
In order to prevent gas leakage between the rotor blade 3 and the stator blade 2, a sealing member 5 is provided between the rotor blade 3 and the stator blade 2, and the sealing member 5 may also prevent radial play of gas flow between adjacent passages.
According to another aspect of the invention, a gas turbine is also provided, and the gas turbine comprises the compressor of the gas turbine.
According to another aspect of the invention, there is also provided an aircraft comprising a compressor of a gas turbine as described above.
The present invention is not limited to the above exemplary embodiments, and any modifications, equivalent replacements, improvements, etc. within the spirit and principle of the present invention should be included in the protection scope of the present invention.
Claims (13)
1. A compressor for a gas turbine engine, comprising:
a housing (1);
the rotating shaft (6) is rotatably arranged in the shell (1); and
the compressor comprises a plurality of compression flow passages, wherein the compression flow passages are arranged between a rotating shaft (6) and a shell (1) side by side in the radial direction and are sequentially communicated in the gas flowing direction, a plurality of stator blades (2) and a plurality of rotor blades (3) are arranged in each compression flow passage side by side in the gas flowing direction, and the stator blades (2) and the rotor blades (3) are alternately arranged.
2. The compressor of a gas turbine according to claim 1, wherein the compression flow path extends in an axial direction.
3. The compressor of a gas turbine according to claim 1, wherein a plurality of the compression flow passages collectively form a flow passage having an S-shaped longitudinal section.
4. The compressor of a gas turbine engine as set forth in claim 1, wherein said plurality of compression runners includes:
a first compression channel (10) comprising an inlet at a first axial end of the shaft (6) and an outlet at a second axial end;
a second compression channel (20) comprising an outlet at said first axial end and an inlet at said second axial end, said inlet of said second compression channel (20) being in communication with said outlet of said first compression channel (10); and
a third compression flowpath (30) including an inlet at the first axial end and an outlet at the second axial end, the inlet of the third compression flowpath (30) communicating with the outlet of the second compression flowpath (20).
5. Compressor for a gas turbine according to claim 1, characterised in that the compression flow path is an annular flow path coaxial with the shaft (6).
6. The compressor of a gas turbine according to claim 1,
the stator blades (2) are connected with the shell (1) and extend towards the rotating shaft (6) along the radial direction; and
the rotor blades (3) are connected to the rotor shaft (6) and extend radially towards the housing (1).
7. The compressor of the gas turbine according to any one of claims 1 to 6, further comprising an isolation member, the isolation member including:
a first partition member (7) connected to the stator blade (2) and extending toward the rotor blade (3) adjacent to the stator blade (2) to divide a space between the rotary shaft (6) and the casing (1) into different compression flow passages arranged side by side in a radial direction; and
and a second partition member (8) connected to the rotor blade (3) and extending toward the stator blade (2) adjacent to the rotor blade (3) to divide a space between the rotary shaft (6) and the housing (1) into different compression flow passages arranged side by side in a radial direction.
8. Compressor for a gas turbine according to claim 7, characterised in that said spacer element further comprises a sealing element (5) connected between said first spacer element (7) and said second spacer element (8).
9. The compressor of a gas turbine according to claim 7,
the first isolation part (7) is in a ring shape coaxial with the rotating shaft (6);
the second isolation component (8) is annular and coaxial with the rotating shaft (6).
10. A compressor for a gas turbine engine according to claim 1, comprising a compressor for an aircraft engine.
11. The compressor of a gas turbine according to claim 2, wherein a plurality of the compression flow passages form a flow passage having an S-shaped longitudinal section.
12. A gas turbine engine, characterized by comprising a compressor of a gas turbine engine as claimed in any one of claims 1 to 11.
13. An aircraft, characterized in that it comprises a compressor of a gas turbine as claimed in any one of claims 1 to 11.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110941140.9A CN113389755B (en) | 2021-08-17 | 2021-08-17 | Compressor of gas turbine, gas turbine and aircraft |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110941140.9A CN113389755B (en) | 2021-08-17 | 2021-08-17 | Compressor of gas turbine, gas turbine and aircraft |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN113389755A CN113389755A (en) | 2021-09-14 |
| CN113389755B true CN113389755B (en) | 2021-12-28 |
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| Application Number | Title | Priority Date | Filing Date |
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| CN202110941140.9A Active CN113389755B (en) | 2021-08-17 | 2021-08-17 | Compressor of gas turbine, gas turbine and aircraft |
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Citations (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN202165129U (en) * | 2011-07-13 | 2012-03-14 | 哈尔滨工程大学 | Radial flow type turbine with splitter vanes |
Family Cites Families (10)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH07247996A (en) * | 1994-03-11 | 1995-09-26 | Ishikawajima Harima Heavy Ind Co Ltd | Compressor passage shape |
| US8192141B1 (en) * | 2007-04-05 | 2012-06-05 | The United States Of America As Represented By The Secretary Of The Air Force | Dual compression rotor |
| JP5502695B2 (en) * | 2010-10-14 | 2014-05-28 | 株式会社日立製作所 | Axial flow compressor |
| US9284851B2 (en) * | 2012-02-21 | 2016-03-15 | Mitsubishi Heavy Industries, Ltd. | Axial-flow fluid machine, and variable vane drive device thereof |
| CN107454922B (en) * | 2015-04-10 | 2020-11-03 | 开利公司 | Integrated fan heat exchanger |
| US10711702B2 (en) * | 2015-08-18 | 2020-07-14 | General Electric Company | Mixed flow turbocore |
| US10260523B2 (en) * | 2016-04-06 | 2019-04-16 | Rolls-Royce North American Technologies Inc. | Fluid cooling system integrated with outlet guide vane |
| CN109209995B (en) * | 2017-06-30 | 2020-04-28 | 中国航发商用航空发动机有限责任公司 | Axial flow compressor |
| DE102019006484B3 (en) * | 2019-09-11 | 2020-08-06 | Friedrich Grimm | MANIFOLDED POWER PLANT WITH AT LEAST ONE TORQUE LEVEL |
| CN112160919A (en) * | 2020-09-28 | 2021-01-01 | 东北大学 | Turbomolecular pump and composite molecular pump including the same |
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2021
- 2021-08-17 CN CN202110941140.9A patent/CN113389755B/en active Active
Patent Citations (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN202165129U (en) * | 2011-07-13 | 2012-03-14 | 哈尔滨工程大学 | Radial flow type turbine with splitter vanes |
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| CN113389755A (en) | 2021-09-14 |
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