CN113863993A - Stator blade angle adjusting mechanism in aircraft engine - Google Patents
Stator blade angle adjusting mechanism in aircraft engine Download PDFInfo
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- CN113863993A CN113863993A CN202111250280.8A CN202111250280A CN113863993A CN 113863993 A CN113863993 A CN 113863993A CN 202111250280 A CN202111250280 A CN 202111250280A CN 113863993 A CN113863993 A CN 113863993A
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- stator
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- aircraft engine
- stator blade
- stator casing
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- 230000001681 protective effect Effects 0.000 claims description 5
- 230000001360 synchronised effect Effects 0.000 claims description 3
- 238000006073 displacement reaction Methods 0.000 claims 1
- 230000010355 oscillation Effects 0.000 claims 1
- 238000012986 modification Methods 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 238000006467 substitution reaction Methods 0.000 description 2
- 230000004323 axial length Effects 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 230000007717 exclusion Effects 0.000 description 1
- 230000000750 progressive effect Effects 0.000 description 1
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Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D7/00—Rotors with blades adjustable in operation; Control thereof
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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 angle adjustment design, concretely relates to stator blade angle adjustment mechanism among aero-engine, include: the plurality of rocker arms are distributed along the circumferential direction of the stator casing; 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 stator casing and the adjacent rotor casing are butted through an annular connecting edge protruding out of the periphery to form an annular protruding part; the linkage ring is sleeved on the protruding part and hinged with the other end of each rocker arm, can move axially along the stator casing and can rotate circumferentially along the stator casing, so that each rocker arm can be driven to synchronously swing circumferentially on the stator casing, and each stator blade can synchronously rotate.
Description
Technical Field
The application belongs to the technical field of aero-engine stator blade angle adjusting design, and particularly relates to an aero-engine stator blade angle adjusting mechanism.
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, each stator blade is adjusted to rotate synchronously through an angle adjusting mechanism, so that the angle of each stator blade is changed synchronously, and the gas flow flowing through the compressor is adjusted.
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 stator blade angle adjusting mechanism in the existing aircraft engine 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 connected between the stator casing and the linkage ring to drive the linkage ring to move and drive each rocker arm to synchronously swing in the circumferential direction of the stator casing, so that each stator blade synchronously rotates, and the synchronous adjustment of the rotating angle of each stator blade is realized, and the technical scheme has the following defects:
rotor blades and stator blades of each stage in the gas compressor are distributed at intervals, a stator casing is arranged on the periphery of each stage of stator blade, a rotor casing is arranged on the periphery of each stage of rotor blade, adjacent stator casings and rotor casings are butted through annular connecting edges protruding out of the periphery, the angle of each stator blade is adjusted through the stator blade angle adjusting mechanism, the linkage ring can move along the axial direction, in order to avoid interference between the linkage ring and the annular connecting edges, the length of the corresponding part of the stator casing is usually increased, so that an enough space range can be provided for the linkage ring to move along the axial direction, the axial length of the whole structure of the gas compressor is larger and not compact enough, and the improvement of the performance of the gas compressor is limited.
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 a stator vane angle adjustment mechanism in an aircraft engine that overcomes or mitigates at least one aspect of the technical disadvantages known to exist.
The technical scheme of the application is as follows:
a stator blade angle adjustment mechanism in an aircraft engine, comprising:
the plurality of rocker arms are distributed along the circumferential direction of the stator casing; 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 stator casing and the adjacent rotor casing are butted through an annular connecting edge protruding out of the periphery to form an annular protruding part;
the linkage ring is sleeved on the protruding part and hinged with the other end of each rocker arm, can move axially along the stator casing and can rotate circumferentially along the stator casing, so that each rocker arm can be driven to synchronously swing circumferentially on the stator casing, and each stator blade can synchronously rotate.
According to at least one embodiment of the application, in the stator blade angle adjusting mechanism in the aircraft engine, the outer edge of the annular connecting edge of the stator casing is folded over to cover the annular connecting edge of the rotor casing.
According to at least one embodiment of the present application, the stator blade angle adjusting mechanism in an aircraft engine further includes:
and each pin correspondingly hinges one end of one rocker arm back to the corresponding upper journal on the periphery of the linkage ring.
According to at least one embodiment of the present application, the stator blade angle adjusting mechanism in an aircraft engine further includes:
and each joint bearing is correspondingly connected with one end of one pin and the corresponding rocker arm, which is opposite to the corresponding upper shaft neck.
According to at least one embodiment of the present application, the stator blade angle adjusting mechanism in an aircraft engine further includes:
and the protective cover is connected to the outer walls of the stator casing and the rotor casing and covers each upper shaft neck, each rocker arm, the annular convex part and the linkage ring.
According to at least one embodiment of the present application, the stator blade angle adjusting mechanism in an aircraft engine further includes:
and the actuating cylinder is connected between the linkage ring and the stator casing or the rotor casing so as to drive the linkage ring to move along the axial direction of the stator casing and rotate along the circumferential direction of the stator casing.
Drawings
FIG. 1 is a schematic view of a stator vane angle adjustment mechanism in an aircraft engine provided by an embodiment of the present application;
FIG. 2 is a schematic view of the A-direction partial structure of FIG. 1;
wherein:
1-a rocker arm; 2-a stator case; 3-stator blades; 4-rotor case; 5-a linkage ring; 6-pin; 7-knuckle bearing; 8-a protective cover; 9-actuating cylinder.
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. 1-2.
A stator blade angle adjustment mechanism in an aircraft engine, comprising:
a plurality of rocker arms 1 distributed along the circumferential direction of the stator case 2; one end of each rocker arm 1 is correspondingly connected with an upper journal of a stator blade 3 extending out of a mounting hole of the stator casing 2; the stator casing 2 is butted with the adjacent rotor casing 4 through an annular connecting edge protruding out of the periphery to form an annular protruding part;
the link ring 5 is sleeved on the protruding portion and hinged to the other end of each rocker arm 1, can axially move along the stator casing 2 and can circumferentially rotate along the stator casing 2, so that each rocker arm 1 can be driven to synchronously swing in the circumferential direction of the stator casing 2, and each stator blade 3 can synchronously rotate.
For the angle adjusting mechanism for the stator blade in the aircraft engine disclosed in the above embodiment, it can be understood by those skilled in the art that the design is to sleeve the link ring 5 on the annular protruding portion formed by the annular connecting edge of the outer periphery of the stator casing 2 and the adjacent rotor casing 4, so as to avoid the interference between the link ring 5 and the annular connecting edge when the link ring 5 moves axially along the stator casing 2, and the annular protruding portion formed by the annular connecting edge of the outer periphery of the stator casing 2 and the adjacent rotor casing 4 abuts against the inner side of the link ring 5, so as to effectively ensure the roundness of the link ring 5, avoid the great deformation of the link ring 5 along the radial direction, and ensure the accuracy of adjusting the rotation angle of each stator blade 3.
In some optional embodiments, in the above-mentioned stator blade angle adjusting mechanism in an aircraft engine, the outer edge of the annular connecting edge of the stator casing 2 is folded over to cover the annular connecting edge of the rotor casing 4, and the linkage ring 5 is sleeved on the folded portion of the outer edge of the annular connecting edge of the stator casing 2, so as to ensure the stability and reliability of the structure.
In some optional embodiments, in the above stator blade angle adjusting mechanism in an aircraft engine, further comprising:
a plurality of pins 6, each pin 6 hinging one end of one rocker arm 1, facing away from the corresponding upper journal, to the periphery of the link ring 5.
In some optional embodiments, in the above stator blade angle adjusting mechanism in an aircraft engine, further comprising:
and each joint bearing 7 is correspondingly connected with one end of one pin 6, which is opposite to the corresponding upper shaft neck, of the corresponding rocker arm 1.
In some optional embodiments, in the above stator blade angle adjusting mechanism in an aircraft engine, further comprising:
and the protective cover 8 is connected to the outer walls of the stator casing 2 and the rotor casing 4 and covers each upper shaft neck, each rocker arm 1, the annular convex part and the linkage ring.
In some optional embodiments, in the above stator blade angle adjusting mechanism in an aircraft engine, further comprising:
and an actuating cylinder 9 connected between the link ring 5 and the stator casing 2 or the rotor casing 4 to drive the link ring 5 to move axially along the stator casing 2 and to rotate circumferentially along the stator casing 2.
With regard to the stator blade angle adjustment mechanism in the aircraft engine disclosed in the above embodiment, it can be understood by those skilled in the art that the protective cover 8 may be provided with corresponding through holes for the actuator cylinder 9 to pass through, and the actuator cylinder 9 and the corresponding through holes may be provided in plural.
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. A stator blade angle adjustment mechanism among aeroengine, its characterized in that includes:
a plurality of rocker arms (1) distributed along the circumferential direction of the stator casing (2); one end of each rocker arm (1) is correspondingly connected with an upper journal of a stator blade (3) extending out of a mounting hole of the stator casing (2); the stator casing (2) is butted with the adjacent rotor casing (4) through an annular connecting edge protruding out of the periphery to form an annular protruding part;
the cover of link ring (5) is established on the protrusion position, with each the other end of rocking arm (1) is articulated, can follow stator machine casket (2) axial displacement, and can follow stator machine casket (2) circumferential direction, thereby can drive each rocking arm (1) is in stator machine casket (2) synchronous oscillation makes each stator blade (3) synchronous rotation.
2. The mechanism of adjusting the angle of a stator vane in an aircraft engine according to claim 1,
the outer edge of the annular connecting edge of the stator casing (2) is turned over to cover the annular connecting edge of the rotor casing (4).
3. The mechanism of adjusting the angle of a stator vane in an aircraft engine according to claim 1,
further comprising:
a plurality of pins (6), wherein one end of one rocker arm (1) which is back to the corresponding upper journal is hinged to the periphery of the linkage ring (5) through each pin (6).
4. The mechanism of adjusting the angle of a stator vane in an aircraft engine according to claim 3,
further comprising:
and each joint bearing (7) is correspondingly connected with one end of one pin (6) and one end of the corresponding rocker arm (1) which are back to the corresponding upper shaft neck.
5. The mechanism of adjusting the angle of a stator vane in an aircraft engine according to claim 1,
further comprising:
and the protective cover (8) is connected to the outer walls of the stator casing (2) and the rotor casing (4) and covers the upper shaft neck, the rocker arms (1), the annular protruding part and the linkage ring.
6. The mechanism of adjusting the angle of a stator vane in an aircraft engine according to claim 1,
further comprising:
and the actuating cylinder (9) is connected between the linkage ring (5) and the stator casing (2) or the rotor casing (4) so as to drive the linkage ring (5) to axially move along the stator casing (2) and to circumferentially rotate along the stator casing (2).
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202111250280.8A CN113863993A (en) | 2021-10-26 | 2021-10-26 | Stator blade angle adjusting mechanism in aircraft engine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202111250280.8A CN113863993A (en) | 2021-10-26 | 2021-10-26 | Stator blade angle adjusting mechanism in aircraft engine |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN113863993A true CN113863993A (en) | 2021-12-31 |
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ID=78998177
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202111250280.8A Pending CN113863993A (en) | 2021-10-26 | 2021-10-26 | Stator blade angle adjusting mechanism in aircraft engine |
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| Country | Link |
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| CN (1) | CN113863993A (en) |
Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| DE10352099A1 (en) * | 2003-11-08 | 2005-06-09 | Mtu Aero Engines Gmbh | Setting device for guide blades of gas turbine has recess in second end of setting lever for shaft of guide blade |
| CN108350906A (en) * | 2015-11-04 | 2018-07-31 | 川崎重工业株式会社 | variable stator blade operating device |
| CN109162829A (en) * | 2018-09-04 | 2019-01-08 | 中国航发沈阳发动机研究所 | The compressibility of variable cycle engine |
| CN110030111A (en) * | 2019-04-04 | 2019-07-19 | 中国航发沈阳发动机研究所 | A kind of variable cycle engine core engine driving fan level structure |
| CN111312058A (en) * | 2019-11-29 | 2020-06-19 | 中国科学院工程热物理研究所 | Test piece structure of gas compressor |
| CN111561480A (en) * | 2020-05-14 | 2020-08-21 | 中国航发沈阳发动机研究所 | Stator structure |
| CN111911461A (en) * | 2020-08-28 | 2020-11-10 | 中国航发沈阳发动机研究所 | Stator blade angle adjusting mechanism and stator casing structure thereof |
| CN112746991A (en) * | 2019-10-30 | 2021-05-04 | 韩华压缩机株式会社 | Rotating device |
-
2021
- 2021-10-26 CN CN202111250280.8A patent/CN113863993A/en active Pending
Patent Citations (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| DE10352099A1 (en) * | 2003-11-08 | 2005-06-09 | Mtu Aero Engines Gmbh | Setting device for guide blades of gas turbine has recess in second end of setting lever for shaft of guide blade |
| US20050271502A1 (en) * | 2003-11-08 | 2005-12-08 | Mtu Aero Engines Gmbh | Apparatus for adjusting stator vanes |
| CN108350906A (en) * | 2015-11-04 | 2018-07-31 | 川崎重工业株式会社 | variable stator blade operating device |
| CN109162829A (en) * | 2018-09-04 | 2019-01-08 | 中国航发沈阳发动机研究所 | The compressibility of variable cycle engine |
| CN110030111A (en) * | 2019-04-04 | 2019-07-19 | 中国航发沈阳发动机研究所 | A kind of variable cycle engine core engine driving fan level structure |
| CN112746991A (en) * | 2019-10-30 | 2021-05-04 | 韩华压缩机株式会社 | Rotating device |
| CN111312058A (en) * | 2019-11-29 | 2020-06-19 | 中国科学院工程热物理研究所 | Test piece structure of gas compressor |
| CN111561480A (en) * | 2020-05-14 | 2020-08-21 | 中国航发沈阳发动机研究所 | Stator structure |
| CN111911461A (en) * | 2020-08-28 | 2020-11-10 | 中国航发沈阳发动机研究所 | Stator blade angle adjusting mechanism and stator casing structure thereof |
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