CN113883090A - Aero-engine stator blade rotation angle adjusting mechanism - Google Patents
Aero-engine stator blade rotation angle adjusting mechanism Download PDFInfo
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- CN113883090A CN113883090A CN202111275254.0A CN202111275254A CN113883090A CN 113883090 A CN113883090 A CN 113883090A CN 202111275254 A CN202111275254 A CN 202111275254A CN 113883090 A CN113883090 A CN 113883090A
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- stator
- stator blade
- angle adjustment
- positioning column
- adjustment mechanism
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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
- F04D29/26—Rotors specially for elastic fluids
- F04D29/32—Rotors specially for elastic fluids for axial flow pumps
- F04D29/321—Rotors specially for elastic fluids for axial flow pumps for axial flow compressors
- F04D29/322—Blade mountings
- F04D29/323—Blade mountings adjustable
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
The application belongs to the technical field of aero-engine stator blade rotation angle adjustment design, concretely relates to aero-engine stator blade rotation angle adjustment mechanism, include: the positioning columns are connected to the stator casing and distributed at intervals with upper journals of the stator blades extending out of the stator casing mounting holes; and the linkage ring partially surrounds the upper shaft necks on one side, is in toothed meshing connection with the upper shaft necks on the other side, partially surrounds the positioning columns, can slide along the circumferential direction of the stator casing relative to the positioning columns, and accordingly enables the stator blades to rotate synchronously.
Description
Technical Field
The application belongs to the technical field of adjusting and designing the rotating angle of stator blades of aero-engines, and particularly relates to a rotating angle adjusting mechanism of stator blades of aero-engines.
Background
In order to stably work in a compressor of an engine, the gas flow flowing through the compressor needs to be adjusted according to actual conditions, the angle of each stator blade in the compressor is adjustable, and the angle of each stator blade is adjusted to synchronously rotate through an angle adjusting mechanism so as to synchronously change the angle of each stator blade, thereby adjusting the gas flow flowing through the compressor.
Each stator blade in the gas compressor is arranged between a stator casing and an inner ring of the gas compressor and distributed along the circumferential direction, a lower shaft neck of each stator blade is inserted into a mounting hole in the stator inner ring, and an upper shaft neck extends out of the mounting hole in the stator casing. The existing aero-engine stator blade rotation angle adjusting mechanism mainly comprises a plurality of rocker arms, a linkage ring and an actuating cylinder, wherein one end of each rocker arm is correspondingly connected with an upper journal of a stator blade extending out of a stator casing mounting hole; the linkage ring is sleeved on the stator casing and is hinged with the other end of each rocker arm; the actuating cylinder is arranged on the stator casing, and a piston rod of the actuating cylinder is connected with the linkage ring so as to drive the linkage ring to move, so that each rocker arm is driven to synchronously swing in the circumferential direction of the stator casing, each stator blade synchronously rotates, and the synchronous adjustment of the rotating angle of each stator blade is realized, as shown in fig. 1, the angle adjusting mechanism has the following defects:
1) in the process of adjusting the angle of the stator blade by the linkage ring driven by the actuating cylinder, the linkage ring rotates along the circumferential direction of the stator casing and moves axially along the stator casing in a large range, namely, the linkage ring bears axial force and axial force, is easy to deform greatly, is difficult to realize accurate adjustment of the rotation angle of the stator blade, seriously restricts the improvement of the performance of an engine, and needs to occupy a large space outside the stator casing;
2) the link ring needs to move in the axial direction of the stator casing greatly, if an interference structure exists in the axial movement range of the link ring, such as a pipeline, a connecting edge and various bolts outside the stator casing, the link ring needs to be designed to have larger radial size, so that the rigidity of the link ring is weakened, the link ring is easy to deform greatly, the accurate adjustment of the rotation angle of the stator blade is further influenced, the rocker arms need to move outwards in the direction away from the stator casing integrally and adaptively, the height of the journal on each stator blade, which extends out of the mounting hole, is correspondingly increased, the strength of the journal on each stator blade is higher, the overall size of the aircraft engine is increased, similarly, the upper journal of each stator blade is connected with the link ring through the rocker arms, if the interference structure exists outside the stator casing in the swing range of each rocker arm, the above problems also exist;
3) in order to ensure that the angle adjusting mechanism acts smoothly, each rocker arm is connected with the pin through a bearing, and the joint bearing is easy to be blocked due to errors in machining and assembly;
4) the structure is complicated, a plurality of parts and connecting pieces thereof exist, and the assembly is difficult.
The present application has been made in view of the above-mentioned technical drawbacks.
It should be noted that the above background disclosure is only for the purpose of assisting understanding of the inventive concept and technical solutions of the present invention, and does not necessarily belong to the prior art of the present patent application, and the above background disclosure should not be used for evaluating the novelty and inventive step of the present application without explicit evidence to suggest that the above content is already disclosed at the filing date of the present application.
Disclosure of Invention
It is an object of the present application to provide an aircraft engine stator vane turning angle adjustment mechanism to overcome or alleviate at least one of the technical disadvantages of the known prior art.
The technical scheme of the application is as follows:
an aircraft engine stator blade turning angle adjustment mechanism, comprising:
the positioning columns are connected to the stator casing and distributed at intervals with upper journals of the stator blades extending out of the stator casing mounting holes;
and the linkage ring partially surrounds the upper shaft necks on one side, is in toothed meshing connection with the upper shaft necks on the other side, partially surrounds the positioning columns, can slide along the circumferential direction of the stator casing relative to the positioning columns, and accordingly enables the stator blades to rotate synchronously.
According to at least one embodiment of the application, in the aero-engine stator blade rotation angle adjustment mechanism, each positioning column is provided with an annular positioning groove;
one side of the linkage ring corresponding to each positioning column is clamped into the annular positioning groove.
According to at least one embodiment of the present application, the above-mentioned mechanism for adjusting the rotation angle of the stator blade of the aircraft engine further includes:
and each bearing is correspondingly sleeved on one positioning column, and the outer ring of the bearing is contacted with one side of the linkage ring, which corresponds to each positioning column.
According to at least one embodiment of the present application, the above-mentioned mechanism for adjusting the rotation angle of the stator blade of the aircraft engine further includes:
and each lantern ring is correspondingly sleeved on one upper shaft neck and is in meshed connection with one side of the linkage ring corresponding to each upper shaft neck through insections.
According to at least one embodiment of the application, in the aircraft engine stator blade rotation angle adjusting mechanism, each sleeve ring is provided with a special-shaped hole; each special-shaped hole is correspondingly sleeved on one upper journal.
According to at least one embodiment of the present application, the above-mentioned mechanism for adjusting the rotation angle of the stator blade of the aircraft engine further includes:
and the actuating cylinder is connected between the stator casing and the linkage ring and can push the linkage ring to slide along the circumferential direction of the stator casing relative to each positioning column.
Drawings
FIG. 1 is a schematic view of a prior art aero-engine stator blade turning angle adjustment mechanism;
FIG. 2 is a schematic view of an aircraft engine stator blade turning angle adjustment mechanism provided by an embodiment of the present application;
FIG. 3 is a partial view from the A-direction of FIG. 2;
wherein:
1-a positioning column; 2-a stator case; 3-stator blades; 4-a linkage ring; 5-a lantern ring.
For the purpose of better illustrating the embodiments, certain features of the drawings may be omitted, enlarged or reduced, and do not represent the size of an actual product; further, the drawings are for illustrative purposes, and terms describing positional relationships are limited to illustrative illustrations only and are not to be construed as limiting the patent.
Detailed Description
In order to make the technical solutions and advantages of the present application clearer, the technical solutions of the present application will be further clearly and completely described in the following detailed description with reference to the accompanying drawings, and it should be understood that the specific embodiments described herein are only some of the embodiments of the present application, and are only used for explaining the present application, but not limiting the present application. It should be noted that, for convenience of description, only the parts related to the present application are shown in the drawings, other related parts may refer to general designs, and the embodiments and technical features in the embodiments in the present application may be combined with each other to obtain a new embodiment without conflict.
In addition, unless otherwise defined, technical or scientific terms used in the description of the present application shall have the ordinary meaning as understood by one of ordinary skill in the art to which the present application belongs. The terms "upper", "lower", "left", "right", "center", "vertical", "horizontal", "inner", "outer", and the like used in the description of the present application, which indicate orientations, are used only to indicate relative directions or positional relationships, and do not imply that the devices or elements must have a specific orientation, be constructed and operated in a specific orientation, and when the absolute position of the object to be described is changed, the relative positional relationships may be changed accordingly, and thus, should not be construed as limiting the present application. The use of "first," "second," "third," and the like in the description of the present application is for descriptive purposes only to distinguish between different components and is not to be construed as indicating or implying relative importance. The use of the terms "a," "an," or "the" and similar referents in the context of describing the application is not to be construed as an absolute limitation on the number, but rather as the presence of at least one. The word "comprising" or "comprises", and the like, when used in this description, is intended to specify the presence of stated elements or items, but not the exclusion of other elements or items.
Further, it is noted that, unless expressly stated or limited otherwise, the terms "mounted," "connected," and the like are used in the description of the invention in a generic sense, e.g., connected as either a fixed connection or a removable connection or integrally connected; can be mechanically or electrically connected; they may be directly connected or indirectly connected through an intermediate medium, or they may be connected through the inside of two elements, and those skilled in the art can understand their specific meaning in this application according to the specific situation.
The present application is described in further detail below with reference to fig. 1 to 3.
An aircraft engine stator blade turning angle adjustment mechanism, comprising:
the positioning columns 1 are connected to the stator casing 2 and distributed at intervals with upper journals of the stator blades 3 extending out of the mounting holes of the stator casing 2;
and a linkage ring 4, one side of which partially surrounds each upper shaft neck and is in meshing connection with insections between the upper shaft necks, and the other side of which partially surrounds each positioning column 1 and can slide along the circumferential direction of the stator casing 2 relative to each positioning column 1, so that each stator blade 2 can rotate synchronously.
To the aero-engine stator blade rotation angle adjustment mechanism disclosed in the above embodiment, those skilled in the art can understand that, in the process of adjusting the rotation angle of each stator blade 3, the link ring 4 mainly rotates along the circumferential direction of the stator casing 2, and all parts are stressed in a balanced manner and are not prone to large irregular deformation, so that accurate adjustment of the rotation angle of the stator blade 3 can be ensured.
For the aero-engine stator blade rotation angle adjustment mechanism disclosed in the above embodiment, it can be further understood by those skilled in the art that the positioning posts 1 and the upper journals of the stator blades 3 are arranged in the circumferential direction of the stator casing 2, one side of the link ring 4 partially surrounds the upper journals of the stator blades 3, the other side and the other side partially surround the positioning posts 1, and the link ring is substantially in a continuous S shape, so that the space between the upper journals is fully utilized, the rest space outside the stator casing 2 is not additionally occupied, the structure interference with other parts outside the stator casing 2 does not exist, the link ring 4 does not need to be designed to have a larger radial dimension, so as to ensure the rigidity of the link ring, ensure the accurate adjustment of the rotation angle of the stator blades 3, and not additionally increase the overall dimension of the aero-engine, in addition, the link ring 4 realizes the positioning along the axial direction of the stator casing 2 by means of the positioning posts 1 and the upper journals of the stator blades 3, no extra parts are needed, the structure is simple, and the assembly is easy.
For the aero-engine stator blade rotation angle adjusting mechanism disclosed in the above embodiment, it can be further understood by those skilled in the art that, in the process of adjusting the rotation angle of each stator blade 3, the link ring 5 can directly contact with the journal on each stator blade 3 to transfer force, a rocker arm with a large size is eliminated, the situation that the rocker arm interferes with the external structure of the stator casing 3 does not exist, the overall dimension of the aero-engine is not additionally increased, a corresponding bearing is not required to be arranged to transfer force, clamping stagnation is not likely to occur, the structure is simplified as a whole, and the assembly is easy.
With regard to the aero-engine stator vane rotation angle adjustment mechanism disclosed in the above embodiment, it will be further understood by those skilled in the art that the unique configuration of the link ring 4 is limited, and the mechanism can maintain a relatively large rigidity in a small size, and is particularly suitable for an aero-engine with a relatively small radial size.
In some alternative embodiments, in the above-mentioned aircraft engine stator blade rotation angle adjusting mechanism, each positioning column 1 has an annular positioning groove;
one side card that link ring 4 corresponds each reference column 1 is gone into in the annular positioning groove to realize the location of link ring 4 in radial ascending, restraint, avoid taking place great deformation, influence the accurate regulation of 3 turned angles of stator blade.
In some optional embodiments, the above-mentioned mechanism for adjusting the rotation angle of the stator blade of the aircraft engine further comprises:
each bearing is correspondingly sleeved on one positioning column 1, and the outer ring of each bearing is in contact with one side, corresponding to each positioning column 1, of the linkage ring 4, so that the friction force of each positioning column 1 to the linkage ring 4 is reduced.
In some optional embodiments, the above-mentioned mechanism for adjusting the rotation angle of the stator blade of the aircraft engine further comprises:
and a plurality of lantern rings 5, wherein each lantern ring 5 is correspondingly sleeved on one upper shaft neck and is in meshed connection with the side of the linkage ring 4 corresponding to each upper shaft neck through insections.
In some alternative embodiments, in the above-mentioned mechanism for adjusting the rotation angle of the stator blade of an aircraft engine, each collar 5 has a shaped hole; each special-shaped hole is correspondingly sleeved on one upper journal.
In some optional embodiments, the above-mentioned mechanism for adjusting the rotation angle of the stator blade of the aircraft engine further comprises:
and the actuating cylinder is connected between the stator casing 2 and the linkage ring 4 and can push the linkage ring 4 to slide along the circumferential direction of the stator casing 2 relative to each positioning column 1.
The embodiments are described in a progressive manner in the specification, each embodiment focuses on differences from other embodiments, and the same and similar parts among the embodiments are referred to each other.
Having thus described the present application in connection with the preferred embodiments illustrated in the accompanying drawings, it will be understood by those skilled in the art that the scope of the present application is not limited to those specific embodiments, and that equivalent modifications or substitutions of related technical features may be made by those skilled in the art without departing from the principle of the present application, and those modifications or substitutions will fall within the scope of the present application.
Claims (6)
1. The utility model provides an aeroengine stator blade turned angle adjustment mechanism which characterized in that includes:
the positioning columns (1) are connected to the stator casing (2) and distributed at intervals with upper journals, extending out of the stator casing (2) mounting holes, of the stator blades (3);
and a linkage ring (4) which partially surrounds each upper shaft neck on one side, is in toothed meshing connection with the upper shaft necks on the other side, partially surrounds each positioning column (1), can slide along the circumferential direction of the stator casing (2) relative to each positioning column (1), and accordingly enables each stator blade (2) to synchronously rotate.
2. The aircraft engine stator blade turning angle adjustment mechanism according to claim 1,
each positioning column (1) is provided with an annular positioning groove;
one side of the linkage ring (4) corresponding to each positioning column (1) is clamped into the annular positioning groove.
3. The aircraft engine stator blade turning angle adjustment mechanism according to claim 2,
further comprising:
each bearing is correspondingly sleeved on one positioning column (1), and the outer ring of each bearing is in contact with one side, corresponding to each positioning column (1), of the linkage ring (4).
4. The aircraft engine stator blade turning angle adjustment mechanism according to claim 1,
further comprising:
and each lantern ring (5) is correspondingly sleeved on one upper shaft neck and is in meshed connection with one side of the linkage ring (4) corresponding to each upper shaft neck through insections.
5. The aircraft engine stator blade turning angle adjustment mechanism according to claim 4,
each collar (5) has a profiled hole therein; each special-shaped hole is correspondingly sleeved on one upper journal.
6. The aircraft engine stator blade turning angle adjustment mechanism according to claim 1,
further comprising:
and the actuating cylinder is connected between the stator casing (2) and the linkage ring (4) and can push the linkage ring (4) to slide along the circumferential direction of the stator casing (2) relative to each positioning column (1).
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202111275254.0A CN113883090B (en) | 2021-10-29 | 2021-10-29 | Aero-engine stator blade rotation angle adjusting mechanism |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202111275254.0A CN113883090B (en) | 2021-10-29 | 2021-10-29 | Aero-engine stator blade rotation angle adjusting mechanism |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN113883090A true CN113883090A (en) | 2022-01-04 |
| CN113883090B CN113883090B (en) | 2023-04-07 |
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| Application Number | Title | Priority Date | Filing Date |
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| CN202111275254.0A Active CN113883090B (en) | 2021-10-29 | 2021-10-29 | Aero-engine stator blade rotation angle adjusting mechanism |
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Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB819032A (en) * | 1955-09-30 | 1959-08-26 | Power Jets Res & Dev Ltd | Improvements in or relating to jet propulsion engines |
| CN102834316A (en) * | 2010-04-09 | 2012-12-19 | 斯奈克玛 | Unshrouded fan for turbomachine |
| CN103644144A (en) * | 2013-11-25 | 2014-03-19 | 乐金空调(山东)有限公司 | Inlet guide vane adjusting device of compressor |
| CN104295330A (en) * | 2014-09-16 | 2015-01-21 | 萍乡市德博科技发展有限公司 | Section-variable nozzle ring for turbocharger |
| CN209704644U (en) * | 2019-02-25 | 2019-11-29 | 中国航发商用航空发动机有限责任公司 | Aero-engine |
| CN111288020A (en) * | 2020-02-24 | 2020-06-16 | 中国航发沈阳发动机研究所 | Compressor stator blade linkage structure |
-
2021
- 2021-10-29 CN CN202111275254.0A patent/CN113883090B/en active Active
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB819032A (en) * | 1955-09-30 | 1959-08-26 | Power Jets Res & Dev Ltd | Improvements in or relating to jet propulsion engines |
| CN102834316A (en) * | 2010-04-09 | 2012-12-19 | 斯奈克玛 | Unshrouded fan for turbomachine |
| CN103644144A (en) * | 2013-11-25 | 2014-03-19 | 乐金空调(山东)有限公司 | Inlet guide vane adjusting device of compressor |
| CN104295330A (en) * | 2014-09-16 | 2015-01-21 | 萍乡市德博科技发展有限公司 | Section-variable nozzle ring for turbocharger |
| CN209704644U (en) * | 2019-02-25 | 2019-11-29 | 中国航发商用航空发动机有限责任公司 | Aero-engine |
| CN111288020A (en) * | 2020-02-24 | 2020-06-16 | 中国航发沈阳发动机研究所 | Compressor stator blade linkage structure |
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| Publication number | Publication date |
|---|---|
| CN113883090B (en) | 2023-04-07 |
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