CN112160798A - Aeroengine turbine baffle location structure and aeroengine thereof - Google Patents
Aeroengine turbine baffle location structure and aeroengine thereof Download PDFInfo
- Publication number
- CN112160798A CN112160798A CN202010990095.1A CN202010990095A CN112160798A CN 112160798 A CN112160798 A CN 112160798A CN 202010990095 A CN202010990095 A CN 202010990095A CN 112160798 A CN112160798 A CN 112160798A
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- Prior art keywords
- turbine
- baffle
- groove
- aircraft engine
- disc
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- 230000000712 assembly Effects 0.000 abstract 1
- 238000000429 assembly Methods 0.000 abstract 1
- 238000009434 installation Methods 0.000 description 8
- 238000000034 method Methods 0.000 description 4
- 230000004048 modification Effects 0.000 description 3
- 238000012986 modification Methods 0.000 description 3
- 230000009286 beneficial effect Effects 0.000 description 2
- 238000001816 cooling Methods 0.000 description 1
- 238000006073 displacement reaction Methods 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 238000011900 installation process Methods 0.000 description 1
- 230000014759 maintenance of location Effects 0.000 description 1
- 230000002093 peripheral effect Effects 0.000 description 1
- 230000002035 prolonged effect Effects 0.000 description 1
Images
Classifications
-
- 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
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/30—Fixing blades to rotors; Blade roots ; Blade spacers
- F01D5/32—Locking, e.g. by final locking blades or keys
- F01D5/326—Locking of axial insertion type blades by other means
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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
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/30—Fixing blades to rotors; Blade roots ; Blade spacers
- F01D5/3007—Fixing blades to rotors; Blade roots ; Blade spacers of axial insertion type
- F01D5/3015—Fixing blades to rotors; Blade roots ; Blade spacers of axial insertion type with side plates
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention discloses a positioning structure of a turbine baffle of an aircraft engine, belongs to the technical field of turbine disk blade fixing assemblies, and solves the technical problems of stress concentration and low reliability of a turbine disk in the prior art. This location structure is including installing baffle (2) and the locating component on turbine dish (1), and the blade is installed to turbine dish (1) circumference, wherein: the baffle (2) is arranged on the side surface of the turbine disc (1) and can axially limit the blades; the positioning assembly can clamp the baffle (2) and the turbine disc (1) at the inner circle position of the turbine disc (1). The invention is used for improving the performance of the turbine disc, avoiding the stress concentration of the turbine disc while fixing the blades, improving the reliability of the turbine rotor and meeting the requirement of people on the long service life of the turbine disc.
Description
Technical Field
The invention belongs to the technical field of fixing pieces of blades of turbine disks of aero-engines, and particularly relates to a positioning structure of a turbine baffle of an aero-engine and the aero-engine.
Background
In modern aviation gas turbine engines, fir-tree tenon connection is often used for connecting a blade and a turbine disc, and when the blade of the fir-tree tenon is installed on the turbine disc, an independent structure is needed for axially positioning the blade. In the traditional mode, the structure can be a locking plate with a single function, the structure is simple, the blade is convenient to replace, the method is complex to assemble, and the locking plate is not suitable for air-cooled hollow blades and is less adopted on modern aero-engines.
In the prior art, a bolt is adopted to connect baffle plates, and as shown in figure 1, the baffle plates are connected by bolts. However, the bolt head of the bolted baffle protrudes out of the disc body. When the rotor rotates, the peripheral air flow is disturbed, the cooling efficiency of the turbine is reduced, and holes are formed in the disk, so that stress concentration is easily generated, the fatigue life of the disk is reduced, and the thickness and the weight of the turbine disk are increased.
In view of the above, the present invention is particularly proposed.
Disclosure of Invention
The invention aims to provide a positioning assembly, which solves the technical problems of stress concentration and low reliability of a turbine disc in the prior art. The technical scheme of the scheme has a plurality of technical beneficial effects, which are described as follows:
the present case provides an aeroengine turbine baffle location structure on the one hand, including baffle and the locating component of installing on the turbine dish, the blade is installed to turbine dish circumference, wherein:
the baffle is arranged on the side surface of the turbine disc and can axially limit the blades; the positioning assembly is capable of engaging the baffle and the turbine disk at an inner circular location of the turbine disk.
In a preferred or alternative embodiment, the baffle is mounted on the side of the turbine disk adjacent the exhaust side of the blades.
In a preferred or alternative embodiment, the inner circle part of the turbine disc is provided with a first groove, and a flanging is arranged at a position close to the inner circle part of the turbine disc, and the turbine disc is characterized in that a first protrusion is arranged on the side surface of the baffle plate corresponding to the position of the first groove, and the shape of the first protrusion is matched with that of the first groove;
the positioning assembly can be used for clamping the flanging or the boss and the baffle and can limit the baffle in the axial direction of the turbine disc.
In a preferred or alternative embodiment, the locating component comprises a limiting ring, wherein a second groove is arranged on the outer circle of the limiting ring, and the second groove can be used for clamping the baffle and the turbine disc.
In a preferred or alternative embodiment, the stop collar is provided in a non-closed loop configuration, and the positioning assembly further comprises a locking block, wherein:
the side surface of the limiting ring is connected with the baffle in a buckling mode at a non-closed position to limit the circumferential rotation of the limiting ring.
In a preferred or alternative embodiment, the side surface of the limit ring is provided with a third groove, and a second protrusion is arranged on the baffle plate at a position corresponding to the third groove, and the second protrusion is clamped in the third groove to limit the circumferential rotation of the limit ring.
In a preferred or alternative embodiment, a fourth groove is formed in the non-closed position of the inner circle of the stop collar, and the external force acts on the locking piece to enable the locking piece to be partially embedded into the fourth groove for moving two ends of the stop collar.
In a preferred or alternative embodiment, the non-closed line of the inner circle of the stop collar is the centre line of the fourth groove.
In a preferred or alternative embodiment, the locking piece has a thickness of 0.5-0.7 mm.
In another aspect, the present disclosure provides an aircraft engine equipped with a turbine baffle positioning structure of an aircraft engine as partially or completely described above.
Compared with the prior art, the technical scheme provided by the invention has the following beneficial effects:
the technical scheme of present case adopts the baffle structure of no bolt hookup, need not set up the bolt hole on the turbine dish, just can be in the same place baffle and rim plate assembly reliably. Moreover, the function required by the turbine disc can be realized by a small number of parts, the weight of the engine rotor is effectively reduced, the condition that the local stress concentration coefficient is overlarge due to the opening on the rotor can be reduced, and compared with a baffle plate connected by a bolt, the baffle plate has no protruding structure, the flow loss in the disc cavity is reduced, and the engine performance is favorably improved.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly described below, and 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 is a schematic structural view of a disk with bolt holes according to the prior art;
FIG. 2 is a perspective view of an aircraft engine turbine shroud positioning structure of the present invention;
FIG. 3 is a perspective view of a third groove of the aircraft engine turbine shroud positioning structure of the present invention;
FIG. 4 is a perspective view of the third groove and the second protrusion of the positioning structure of the turbine shroud of the aircraft engine according to the present invention;
FIG. 5 is a perspective view of the lock block assembly of the aircraft engine turbine shroud positioning structure of the present invention;
FIG. 6 is a perspective view of a fourth groove of the aircraft engine turbine shroud positioning structure of the present invention;
FIG. 7 is a perspective view of a locking block of the aircraft engine turbine shroud positioning structure of the present invention.
Wherein:
1. a turbine disk; 11. a blade mounting groove; 12. flanging; 13. a first groove; 2. a baffle plate; 21. a first protrusion; 22. a second protrusion; 31. a limiting ring; 32. a locking block; 311. a second groove; 312. a third groove; 313. and a fourth groove.
Detailed Description
The embodiments of the present invention are described below with reference to specific embodiments, and other advantages and effects of the present invention will be easily understood by those skilled in the art from the disclosure of the present specification. It is to be understood that the described embodiments are merely exemplary of the invention, and not restrictive of the full scope of the invention. The invention is capable of other and different embodiments and of being practiced or of being carried out in various ways, and its several details are capable of modification in various respects, all without departing from the spirit and scope of the present invention. It is to be noted that the features in the following embodiments and examples may be combined with each other without conflict. 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.
It is noted that various aspects of the embodiments are described below within the scope of the appended claims. It should be apparent that the aspects described herein may be embodied in a wide variety of forms and that any specific structure and/or function described herein is merely illustrative. Based on the disclosure, one skilled in the art should appreciate that one aspect described herein may be implemented independently of any other aspects and that two or more of these aspects may be combined in various ways. For example, an apparatus may be implemented and/or a method practiced using any number of the aspects set forth herein. Additionally, such an apparatus may be implemented and/or such a method may be practiced using other structure and/or functionality in addition to one or more of the aspects set forth herein.
It should be noted that the drawings provided in the following embodiments are only for illustrating the basic idea of the present invention, and the drawings only show the components related to the present invention rather than the number, shape and size of the components in practical implementation, and the type, quantity and proportion of the components in practical implementation can be changed freely, and the layout of the components can be more complicated.
In addition, in the following description, specific details are provided to facilitate a thorough understanding of the examples. However, it will be understood by those skilled in the art that aspects may be practiced without these specific details. In order that those skilled in the art will better understand the disclosure, the invention will be described in further detail with reference to the accompanying drawings and specific embodiments. The terms "first", "second" and "first" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include one or more of that feature. In the description of the present invention, "a plurality" means two or more unless otherwise specified.
As shown in FIGS. 2 to 7, the aeroengine turbine baffle positioning structure comprises a baffle 2 and a positioning assembly which are installed on a turbine disc 1, blades are installed on the circumferential direction of the turbine disc 1, the turbine disc 1 comprises an outer circle part and an inner circle part, a blade installation groove 11 is axially arranged on the turbine disc 1, and tenons of the blades are installed in the blade installation groove 11. When the engine works, the acting force of the airflow F can be received, so that the blades have small displacement in the axial direction of the turbine disc 1 along the airflow direction, therefore, the baffle 2 is arranged on the side surface of the turbine disc 1 in the opposite direction, the movement of the blades is blocked, and the blades are limited in the axial direction. The positioning component, for example, a groove clamping structure, can clamp the baffle 2 and the turbine disk 1 at the inner circle position of the turbine disk 1 to limit the baffle 2 in the circumferential direction and the axial direction.
This structural arrangement, no bolted connection's baffle structure need not set up the bolt hole on turbine dish 1, can be in the same place baffle 2 and rim plate assembly reliably. Moreover, the function required by the turbine disc 1 can be realized by a small number of parts, the weight of the engine rotor is effectively reduced, the condition that the local stress concentration coefficient is overlarge due to the opening on the rotor can be reduced, and compared with the baffle 2 connected by the bolt, the baffle has no convex structure, the flow loss in the disc cavity is reduced, and the performance of the engine is favorably improved.
It should be noted that the baffle 2 is mounted on the side of the blade which is remote from the outward force of the turbine shaft, or the baffle 2 is mounted on the side opposite to the incoming flow, i.e. at the position a in the figure. The baffle 2 is provided in a hollow circular ring-shaped configuration, which also includes an outer circular portion and an inner circular portion.
As a specific embodiment provided in the present disclosure, as shown in fig. 2, a first groove 13 is disposed at an inner circumferential portion of the turbine disc 1, and a flange 12 or a boss is disposed at a position adjacent to the inner circumferential portion of the turbine disc 1, where the flange 12 is turned downward with reference to the placement direction of fig. 1. The side surface of the baffle plate 2 is provided with a first protrusion 21 corresponding to the position of the first groove 13, and the shape of the first protrusion 21 is matched with that of the first groove 13. Generally, in order to meet the requirement of convenience in installation of the turbine disc 1, the first groove 13 and the flanging 12 or the boss are arranged at the inner circle part of the turbine disc 1, the baffle 2 is clamped with the turbine disc 1 through the first protrusion 21 arranged on the baffle 2, so that circumferential rotation of the baffle 2 on the turbine disc 1 is limited, the positioning assembly is buckled and overturned to the inner circle part of the baffle 2, axial limiting of the baffle 2 is further limited, and the turbine disc 1 and the disc are ensured to rotate together.
As a specific embodiment provided in the present disclosure, as shown in fig. 3, the positioning assembly includes a limiting ring 31, a second groove 311 is disposed on an outer circumference of the limiting ring 31, the second groove 311 can engage with an inner circumference of the baffle 2 and the flange 12 of the turbine disk 1, so as to fix the baffle 2 and the turbine disk 1, that is, fix the baffle 2 and the turbine disk 1 in the axial direction.
As a specific embodiment provided in the present disclosure, as shown in fig. 4, the limiting ring 31 is disposed in a non-closed ring structure, and the positioning assembly further includes a locking block 32, wherein:
the stop ring 31 is connected with the baffle plate 2 in a snap-fit manner at the side of the stop ring 31 and at a non-closed position so as to limit the stop ring 31 to rotate in the circumferential direction.
In order to meet the requirement of convenient installation, the scheme is further improved on the basis of the technology of the scheme, and a cutting line of 1-2mm is adopted to cut the closed limiting ring on the circumference (the traditional process) of the limiting ring to form an unclosed limiting ring or a limiting ring with one open end.
During installation, one end of the non-closed limiting ring 31 extends into the other end of the non-closed limiting ring, and the second groove 311 is clamped with the flange 12 of the turbine disc 1 and the inner circle part of the baffle 2. After the installation is finished, the limiting ring 31 is deformed, and the elastic force of the material for restoring the deformation is utilized for integral installation, so that the installation process is very convenient.
As an embodiment provided by the present disclosure, as shown in fig. 3, a third groove 312 is provided on a side surface of the retainer ring 31, a second protrusion 22 is provided on the baffle 2 at a position corresponding to the third groove 312, and the second protrusion 22 is engaged in the third groove 312 to limit circumferential rotation of the retainer ring 31.
Considering that the above-mentioned retainer ring 31 is of a non-closed structure, for further circumferential retention, the second protrusion 22 on the baffle 2 is engaged in the third groove 312, thereby limiting the circumferential movement of the retainer ring 31.
Further, as shown in fig. 5 and 6, a fourth groove 313 is disposed at a non-closed position of the inner circle of the limiting ring 31, and an external force acts on the locking piece 32 to partially embed the locking piece into the fourth groove 313 for moving two ends of the limiting ring 31.
As a specific embodiment provided in the present application, the locking piece is as shown in fig. 7, and the thickness of the locking piece is 0.5-0.7 mm. During installation, the area of the corresponding position of the locking block 32 and the fourth groove 313 is embedded into the fourth groove 313 by using a tool, namely, the locking block moves along the circumferential direction of the two ends of the limiting ring 31. The lock 32 is made of, for example, GH 4169. The locking piece 32 is L-shaped, one end of the locking piece is provided with an arc-shaped flanging which is connected with the limiting ring 31, and the other end of the locking piece is embedded into the fourth groove 313 by adopting a tool.
In another aspect, the present disclosure provides an aircraft engine equipped with a turbine baffle positioning structure of an aircraft engine as partially or completely described above. The service life of the turbine disk can be prolonged, bolt holes do not need to be formed, and the situation that stress is concentrated when the turbine disk works actually is avoided.
The positioning structure of the turbine baffle plate of the aircraft engine and the aircraft engine thereof provided by the invention are described in detail above. The principles and embodiments of the present invention are explained herein using specific examples, which are presented only to assist in understanding the core concepts of the present invention. It should be noted that, for those skilled in the art, it is possible to make various improvements and modifications to the invention without departing from the inventive concept, and those improvements and modifications also fall within the scope of the claims of the invention.
Claims (10)
1. The utility model provides an aeroengine turbine baffle location structure which characterized in that, including baffle (2) and the locating component of installing on turbine dish (1), the blade is installed to turbine dish (1) circumference, wherein:
the baffle (2) is arranged on the side surface of the turbine disc (1) and can axially limit the blades; the positioning assembly can clamp the baffle (2) and the turbine disc (1) at the inner circle position of the turbine disc (1).
2. The aeroengine turbine baffle positioning structure of claim 1, wherein the baffle (2) is mounted on the side of the turbine disk and adjacent to the exhaust side of the blades.
3. The aeroengine turbine baffle positioning structure according to claim 1 or 2, wherein a first groove (13) is formed in the inner circle of the turbine disc (1), a flanging (12) or a boss is arranged at a position close to the inner circle of the turbine disc (1), a first protrusion (21) is arranged on the side surface of the baffle (2) corresponding to the position of the first groove (13), and the shape of the first protrusion (21) is matched with that of the first groove (13);
the positioning assembly can be used for clamping the flanging (12) and the baffle (2) and can be used for limiting the baffle (2) in the axial direction of the turbine disc (1).
4. The aircraft engine turbine baffle positioning structure of claim 3, wherein the positioning assembly comprises a limiting ring, a second groove is arranged on the outer circle of the limiting ring, and the second groove can be used for clamping the baffle (2) and the turbine disc (1).
5. The aircraft engine turbine shroud positioning structure of claim 4, wherein the stop collar is disposed in a non-closed annular configuration, the positioning assembly further comprising a lock, wherein:
the side surface of the limiting ring is connected with the baffle in a buckling mode at a non-closed position to limit the circumferential rotation of the limiting ring.
6. The aircraft engine turbine baffle positioning structure of claim 5, wherein a third groove is formed in the side surface of the limit ring, a second protrusion is arranged on the baffle (2) at a position corresponding to the third groove, and the second protrusion is clamped in the third groove to limit circumferential rotation of the limit ring.
7. The aircraft engine turbine baffle positioning structure of claim 6, wherein a fourth groove is formed at a non-closed position of the inner circle of the limit ring, and external force acts on the locking piece to enable the locking piece to be partially embedded into the fourth groove for movement of two ends of the limit ring.
8. The aircraft engine turbine baffle positioning structure of claim 7, wherein the non-closed line of the inner circle of the stop collar is a center line of the fourth groove.
9. The aircraft engine turbine shroud positioning structure of claim 6, wherein the thickness of the locking piece is 0.5-0.7 mm.
10. An aircraft engine, characterized in that an aircraft engine turbine shroud positioning structure as claimed in any one of claims 1 to 9 is installed.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202010990095.1A CN112160798A (en) | 2020-09-18 | 2020-09-18 | Aeroengine turbine baffle location structure and aeroengine thereof |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202010990095.1A CN112160798A (en) | 2020-09-18 | 2020-09-18 | Aeroengine turbine baffle location structure and aeroengine thereof |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN112160798A true CN112160798A (en) | 2021-01-01 |
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Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202010990095.1A Pending CN112160798A (en) | 2020-09-18 | 2020-09-18 | Aeroengine turbine baffle location structure and aeroengine thereof |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN112160798A (en) |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114483202A (en) * | 2021-12-17 | 2022-05-13 | 中国航发湖南动力机械研究所 | A spacing subassembly and turbine rotor that is used for root, take hat interlocking blade to extend |
| CN114718658A (en) * | 2021-01-05 | 2022-07-08 | 中国航发商用航空发动机有限责任公司 | Tool and method for controlling width of baffle ring installation groove |
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| US4171930A (en) * | 1977-12-28 | 1979-10-23 | General Electric Company | U-clip for boltless blade retainer |
| RU2315868C1 (en) * | 2005-12-12 | 2008-01-27 | Государственное предприятие "Запорожское машиностроительное конструкторское бюро "Прогресс" имени академика А.Г. Ивченко | Device for fastening blades of rotor wheels of axial-flow turbomachine |
| CN102713161A (en) * | 2010-01-29 | 2012-10-03 | 斯奈克玛 | Means for locking a sealing ring on a turbine wheel |
| CN202756026U (en) * | 2012-06-11 | 2013-02-27 | 湖南航翔燃气轮机有限公司 | Locking mechanism of turbine blades |
| CN204783124U (en) * | 2015-07-09 | 2015-11-18 | 中国航空工业集团公司沈阳发动机设计研究所 | Heavy gas turbine turbine rotor blade axial fixity structure |
| US20180023401A1 (en) * | 2015-02-24 | 2018-01-25 | Siemens Aktiengesellschaft | Wheel disk assembly having simplified sealing-plate mounting |
| CN110439628A (en) * | 2019-09-02 | 2019-11-12 | 重庆天骄航空动力有限公司 | A kind of connection structure of gas turbine blades and the turbine disk |
| CN111022128A (en) * | 2019-12-05 | 2020-04-17 | 中国航发四川燃气涡轮研究院 | Integral blade ring structure and manufacturing method thereof |
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2020
- 2020-09-18 CN CN202010990095.1A patent/CN112160798A/en active Pending
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4171930A (en) * | 1977-12-28 | 1979-10-23 | General Electric Company | U-clip for boltless blade retainer |
| RU2315868C1 (en) * | 2005-12-12 | 2008-01-27 | Государственное предприятие "Запорожское машиностроительное конструкторское бюро "Прогресс" имени академика А.Г. Ивченко | Device for fastening blades of rotor wheels of axial-flow turbomachine |
| CN102713161A (en) * | 2010-01-29 | 2012-10-03 | 斯奈克玛 | Means for locking a sealing ring on a turbine wheel |
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| CN204783124U (en) * | 2015-07-09 | 2015-11-18 | 中国航空工业集团公司沈阳发动机设计研究所 | Heavy gas turbine turbine rotor blade axial fixity structure |
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Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114718658A (en) * | 2021-01-05 | 2022-07-08 | 中国航发商用航空发动机有限责任公司 | Tool and method for controlling width of baffle ring installation groove |
| CN114483202A (en) * | 2021-12-17 | 2022-05-13 | 中国航发湖南动力机械研究所 | A spacing subassembly and turbine rotor that is used for root, take hat interlocking blade to extend |
| CN114483202B (en) * | 2021-12-17 | 2023-11-17 | 中国航发湖南动力机械研究所 | Limiting assembly for long-extension root and crown interlocking blade and turbine rotor |
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