CN212837968U - Turbine disc and blade locking mechanism for turboprop engine - Google Patents
Turbine disc and blade locking mechanism for turboprop engine Download PDFInfo
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- CN212837968U CN212837968U CN202020922849.5U CN202020922849U CN212837968U CN 212837968 U CN212837968 U CN 212837968U CN 202020922849 U CN202020922849 U CN 202020922849U CN 212837968 U CN212837968 U CN 212837968U
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Abstract
The utility model discloses a turbine disc and blade locking mechanism for a turboprop engine, which comprises a turbine disc and a plurality of blades uniformly distributed on the periphery of the turbine disc, wherein the blades are arranged on the periphery of the turbine disc; the periphery of the turbine disc is uniformly provided with a mortise structure, the mortise structure consists of a first mortise positioned on the outer side and a second mortise positioned on the inner side, and the first mortise is communicated with the second mortise; the blade root is provided with a tenon mechanism, and the tenon mechanism consists of a first tenon and a second tenon positioned at the end part of the first tenon; wherein, the first tenon is embedded in the first mortise; the second tenon is embedded in the second mortise and is locked and fixed by a locking plate embedded between the second tenon and the second mortise. The turbine disc and blade locking mechanism provided by the utility model has the advantages that after the blades are matched with the turbine disc through the tenons and the mortises, only the locking plates need to be inserted and bent, and the operation is simple; the turbine has the advantages of novel design, simple and convenient assembly, strong practicability and excellent stability, and can be widely applied to turbines with various functions.
Description
Technical Field
The utility model relates to a turboprop engine technical field especially relates to a turboprop is turbine disc and blade locking mechanism for turboprop engine.
Background
The rotor blade of the turboprop engine works in a severe environment with high temperature and high rotating speed, and bears the centrifugal force, aerodynamic force, thermal stress and vibration load of the blade when the rotor rotates at high speed, the blade is connected with a turbine disc through a tenon, the blade is a part with larger load in the engine, the centrifugal force borne by the root of the rotor blade of a general turbine is higher, the tenon works at high temperature, the mechanical performance of materials is greatly reduced, and the failure is easy to occur. Therefore, in the turbine rotor connection structure, the connection of the turbine blades to the turbine disk is important, and the operating environment of the turbine causes the blades to bear a great load.
In the traditional fixing mode of the turbine disk and the blades, the turbine blades are matched with a tongue-and-groove structure on the turbine disk through tenons, so that the turbine is embedded into the periphery of the turbine disk. In addition, in a turbine in a modern aircraft engine, the most widely applied mode is that the root of a blade is designed into a longitudinal tree-shaped tenon, the tenon is wedge-shaped, trapezoidal semicircular tooth forms are symmetrically distributed on two sides of the tenon, and the tenon tooth form is generally 4-6 pairs.
In order to prevent the turbine blades from moving axially, the front end and the rear end of a wheel disc of some engine turbine rotors adopt blade baffle structures. The baffle plates and the disc are connected through bolts in two connection modes, one mode is that through holes are formed between mortises at the edge of the disc, and the front baffle plate and the rear baffle plate are connected to the disc through bolts. One is that the edge of the disc at one side of the disc where the baffle needs to be installed is flanged inwards to form an installation edge, and the baffle is fixed on the installation edge by screws. The analysis of the longitudinal tree-shaped tenon tooth connection structure is described in detail in the related literature, and is not described in detail here. However, the existing fixing mode of adopting the longitudinal tree-shaped tenon blade and the turbine disc has the common defect that the bolt head protrudes out of the disc body, and when the rotor rotates, the air flow around is disturbed, so that the temperature is increased, and the efficiency of the turbine is influenced.
The existing bolt-free baffle plate structure changes a bolt into a clamping ring for connection, and comprises a baffle plate and the clamping ring, wherein the clamping ring is open. During assembly, the clamping ring is pressed tightly by a tool, the baffle plate is arranged in the clamping ring, the tool pressing plate is used for elastically deforming the clamping ring until the leftmost side leans against the plate, at the moment, the clamping ring is loosened to rebound, then the baffle plate is loosened to return to a free state, and the baffle plate and the clamping ring are locked. However, the structure is complex, and the processing and assembling difficulty is high (see the tenth volume P227 of the aviation engine design manual for an illustration).
SUMMERY OF THE UTILITY MODEL
The utility model aims to solve the technical problem that: aiming at the defects in the prior art, a turbine disc and blade locking mechanism for a turboprop engine is provided.
In order to achieve the above purpose, the utility model adopts the following technical scheme:
the utility model provides a turbine disc and blade locking mechanism for a turboprop engine, which comprises a turbine disc and a plurality of blades uniformly distributed on the periphery of the turbine disc, wherein the blades are arranged on the periphery of the turbine disc;
the periphery of the turbine disc is uniformly provided with a mortise structure, the mortise structure consists of a first mortise positioned on the outer side and a second mortise positioned on the inner side, and the first mortise is communicated with the second mortise;
the blade root is provided with a tenon mechanism, and the tenon mechanism consists of a first tenon and a second tenon positioned at the end part of the first tenon;
the first tenon is embedded in the first mortise; the second tenon is embedded in the second mortise and locked and fixed through a locking plate embedded between the second tenon and the second mortise.
Further, turbine disc and blade locking mechanism for turboprop on, first tongue-and-groove is for indulging tree-shaped tongue-and-groove, first tenon be with first tongue-and-groove matched with indulge tree-shaped tenon.
Further, on the turbine disc and the blade locking mechanism for the turboprop engine, the cross sections of the second mortise and the second tenon are in a convex structure matched with each other.
Further preferably, on the turbine disc and the blade locking mechanism for the turboprop engine, groove platforms are respectively arranged at the convex waist positions at the upper end and the lower end of the second mortise.
Further preferably, on the turbine disc for the turboprop engine and the blade locking mechanism, mortises flush with the groove platforms are respectively formed at the upper end and the lower end of the second tenon.
Further comparatively preferably, turboprop for the turboprop engine and blade locking mechanism on, the locking plate is the cuboid structure, its by the inlay and by the spacing section that both ends were buckled and are formed about the inlay constitutes.
Still more preferably, on the turbine disk for the turboprop and the blade locking mechanism, the inlay is embedded between the second tenon and the second mortise in an interference fit manner.
Furthermore, on the turbine disc and the blade locking mechanism for the turboprop engine, the limiting sections at the upper end and the lower end are horizontally attached to a mortise formed by the groove platform and the tenon platform.
Still further preferably, in the turbine disc and blade locking mechanism for a turboprop engine, an outer end surface of the limiting section is not higher than an outer end surface of the turbine disc, and a thickness of the limiting section is at least 0.3 mm.
More preferably, in the turbine disk and blade locking mechanism for a turboprop engine, the inlay has a thickness of 0.8 to 1.2mm, and a total length at least 4mm greater than the thickness of the turbine disk.
The above technical scheme is adopted in the utility model, compared with the prior art, following technological effect has:
(1) the locking plate adopted by the locking structure is simple in structure, and after the blade is matched with the turbine disc through the tenon and the mortise, the locking plate only needs to be inserted and bent, so that the operation is simple; the working efficiency of fixing the blades and the turbine disc is improved, and the fixing difficulty and cost are reduced;
(2) the tenon structure at the root of the blade has simpler processing difficulty, and the head of the second tenon is a plane instead of a special-shaped shape, so that the processing time of the blade is shortened, and the manufacturing efficiency of the turbine is further improved;
(3) the locking plate inserts between second tenon and the second tongue-and-groove with the interference mode, provides sufficient pretightning force for blade and turbine dish, and the pretightning force direction is the same with during operation blade tenon and turbine dish tongue-and-groove effort, and the reliability that blade and turbine dish are connected can be guaranteed to abundant frictional force:
(4) the bent sections at the upper end and the lower end of the locking plate also play a role in preventing the turbine blade from moving, and meanwhile, the bent parts fill the turbine disc grooves, so that disturbance to surrounding airflow during rotation is greatly reduced, and the stability of the turbine structure of the engine is further ensured;
(5) this turbine dish and blade locking mechanism, the modern design, the assembly is simple, convenient, and the practicality is strong, and the stability can be superior, but the wide application in the turbine of all kinds of functions.
Drawings
Fig. 1 is a schematic view of an overall structure of a turbine disk and blade locking mechanism for a turboprop engine according to the present invention;
fig. 2 is a schematic view of a partial structure of a turbine disk and blade locking mechanism for a turboprop engine according to the present invention;
fig. 3 is a schematic cross-sectional structural view of a turbine disk and blade locking mechanism for a turboprop engine according to the present invention;
FIG. 4 is a schematic structural view of a locking plate in a turbine disk and blade locking mechanism for a turboprop engine according to the present invention;
wherein the reference symbols are:
10-turbine disk, 11-first mortise, 12-second mortise and 13-groove platform; 20-blades; 21-a first tenon, 22-a second tenon, 23-a tenon table; 30-locking plate, 31-inlay and 32-spacing segment.
Detailed Description
The present invention will be described in detail and specifically with reference to specific embodiments so as to provide a better understanding of the present invention, but the following embodiments do not limit the scope of the present invention.
Referring to fig. 1, the present embodiment provides a turbine disk and blade locking mechanism for a turboprop engine, including a turbine disk 10 and a plurality of blades 20 uniformly distributed around the turbine disk 10, wherein the plurality of blades 20 are embedded on an outer periphery of the turbine disk 10 by a tongue-and-groove and tenon connection manner, and are locked and fixed by a locking plate 30. Specifically, the blade 20 is embedded in the tongue-and-groove structure of the turbine disk 10 by using the tongue structure of the root thereof, and the blade 20 is fixed on the turbine disk 10 by the locking plate 30.
In this embodiment, please refer to fig. 2, a mortise structure is uniformly distributed on the periphery of the turbine disk 10, the mortise structure is composed of a first mortise 11 located on the outer side and a second mortise 12 located on the inner side, and the first mortise 11 is communicated with the second mortise 12; the root of the blade 20 is provided with a tenon mechanism, and the tenon mechanism consists of a first tenon 21 and a second tenon 22 positioned at the end part of the first tenon 21; wherein, the first tenon 21 is embedded in the first mortise 11; the second tenon 22 is embedded in the second mortise 12 and is locked and fixed by a locking plate 30 embedded between the second tenon 22 and the second mortise 12.
In this embodiment, please continue to refer to fig. 2, the first mortise 11 is a longitudinal tree-shaped mortise, the first tenon 21 is a longitudinal tree-shaped tenon matched with the first mortise 11, and the stability of the matching between the first tenon 21 and the first mortise 11 is ensured by the mutual matching between the longitudinal tree-shaped tenon and the longitudinal tree-shaped mortise.
In this embodiment, please continue to refer to fig. 2, the processing difficulty of the tenon structure at the root of the blade 20 is simpler, specifically, the cross sections of the second tenon groove 12 and the second tenon 22 are in a mutually matched convex structure, the side end surfaces of the convex structure are planar, and the head of the second tenon 22 and the inner side surface of the second tenon groove 12 are both planar instead of in an opposite shape, so that the processing time of the second tenon 22 at the root of the blade 10 and the second tenon groove 12 on the turbine disk 10 is shortened, and the manufacturing efficiency of the turbine is further improved.
In this embodiment, please refer to fig. 3, in order to further ensure the stability of the turbine structure of the engine, a groove 13 is respectively disposed at the convex waist positions at the upper and lower ends of the second tongue 12, and a tenon 23 flush with the groove 13 is respectively disposed at the upper and lower ends of the second tongue 22. And the groove platform 13 and the tenon platform 23 are arranged in parallel and level, and the upper space of the groove platform forms a mortise for accommodating the bent end part of the locking plate 30. The limiting sections 32 at the upper end and the lower end of the locking plate 30 are horizontally attached to the mortise formed by the groove platform 13 and the tenon platform 14, and the limiting sections 32 at the two ends are tightly attached to the upper end and the lower end of the groove platform 13 and simultaneously play a limiting role in the tenon platform 23, namely the limiting sections 32 formed by bending the upper end and the lower end of the locking plate 10 play a role in preventing the turbine blade 20 from moving up and down.
In the present embodiment, please refer to fig. 4, the locking plate 30 is a rectangular parallelepiped structure, and is composed of an inlay 31 and a limiting section 32 formed by bending upper and lower ends of the inlay 31. The adopted locking plate 30 is simple in structure, after the blade 20 and the turbine disc 10 are matched through the tenon structure and the mortise structure, the locking plate 30 only needs to be inserted, and then the upper end and the lower end of the locking plate 30 are bent, so that the assembling operation is simple and convenient, the working efficiency of fixing the blade 20 and the turbine disc 10 is greatly improved, and the fixing difficulty and the cost of the blade 20 are reduced.
In the present embodiment, please refer to fig. 4, the inlay 31 is embedded between the second tenon 22 and the second mortise 12 in an interference fit manner, and the limiting segments 32 at the upper and lower ends are horizontally attached to the mortise formed by the slot platform 13 and the mortise 14. The locking plate 30 is inserted between the second tenon 22 and the second mortise 12 in an interference manner, so that sufficient pre-tightening force is provided for the blade 20 and the turbine disk 10, the pre-tightening force direction is the same as the acting force of the tenon structure of the blade 20 and the mortise structure of the turbine disk 10 in operation, and the reliability and the stability of connection between the blade 20 and the turbine disk 10 can be ensured by sufficient friction force.
In this embodiment, the depth of the mortise formed by the groove platform 13 and the tenon platform 14 on the turbine disk 10 needs to ensure that the outer end surface of the limiting section 32 is not higher than the outer end surface of the turbine disk 10 after the fixing locking plate 32 is bent, and the thickness of the limiting section 32 at the upper end and the lower end is at least 0.3 mm.
In the present embodiment, the crescent-shaped turbine blades 20 allow the turbine to obtain maximum power after the turbine hot gas acts on the turbine blades. Turbine blades 20 are typically angled to direct the hot gases into and out of the direction. Locking tabs 30 are arranged non-axially and the angle of inclination of the arrangement in the mortise slot is related to the configuration of turbine blade 20. The thickness of the inlay 31 is 0.8-1.2mm, preferably 1 mm; and the total length of the inlay 31 (including the length of the position-limiting section 32 before bending) is at least 4mm greater than the thickness of the turbine disk 10, i.e. the length of the position-limiting section 32 is at least 2 mm.
In addition, in the embodiment, since the turbine disk and the blade generally need to bear a large working stress and a high working temperature, and the stress and the temperature change frequently and severely, there is a certain requirement on the material of the locking plate 30. The locking plate 30 is usually made of materials with good high temperature resistance, corrosion resistance and heat strength, such as INCONEL718 and 347H.
According to the turbine disc and blade locking mechanism for the turboprop engine, after the blades are matched with the turbine disc through the tenons and the mortises, only the locking plates need to be inserted and bent, and the operation is simple; the working efficiency of fixing the blades and the turbine disc is improved, and the fixing difficulty and cost are reduced; the turbine has the advantages of novel structural design, simple and convenient assembly, strong practicability and excellent stability, and can be widely applied to turbines with various functions.
The above detailed description of the embodiments of the present invention is only for exemplary purposes, and the present invention is not limited to the above described embodiments. Any equivalent modifications and substitutions to those skilled in the art are also within the scope of the present invention. Accordingly, variations and modifications in equivalents may be made without departing from the spirit and scope of the invention, which is intended to be covered by the following claims.
Claims (10)
1. A turbine disk and blade locking mechanism for a turboprop engine is characterized by comprising a turbine disk and a plurality of blades uniformly distributed on the periphery of the turbine disk, wherein the blades are arranged on the periphery of the turbine disk;
the periphery of the turbine disc is uniformly provided with a mortise structure, the mortise structure consists of a first mortise positioned on the outer side and a second mortise positioned on the inner side, and the first mortise is communicated with the second mortise;
the blade root is provided with a tenon structure, and the tenon structure consists of a first tenon and a second tenon positioned at the end part of the first tenon;
the first tenon is embedded in the first mortise; the second tenon is embedded in the second mortise and locked and fixed through a locking plate embedded between the second tenon and the second mortise.
2. The turbine disk and blade locking mechanism for a turboprop according to claim 1, wherein the first mortise is a longitudinal tree-shaped mortise, and the first tenon is a longitudinal tree-shaped tenon engaged with the first mortise.
3. The turbine disk and blade locking mechanism for a turboprop according to claim 1, wherein the cross section of the second tongue-and-groove and the cross section of the second tongue-and-groove are in a mutually-matched convex structure.
4. The turbine disk and blade locking mechanism for the turboprop engine according to claim 3, wherein the raised waist portions at the upper and lower ends of the second mortise are respectively provided with a groove.
5. The turbine disk and blade locking mechanism of claim 4, wherein the second tenon has a tenon formed at its upper and lower ends, respectively, to be flush with the groove.
6. The turbine disk and blade locking mechanism for the turboprop engine according to claim 5, wherein the locking plate is of a rectangular parallelepiped structure and is composed of an inlay and a limiting section formed by bending the upper end and the lower end of the inlay.
7. The turbine disk and blade locking mechanism for a turboprop according to claim 6, wherein the inlay is inserted between the second tenon and the second mortise in an interference fit manner.
8. The turbine disk and blade locking mechanism for a turboprop according to claim 6,
the upper end and the lower end of the limiting section are horizontally attached to the mortise formed by the groove platform and the tenon platform.
9. The turbine disk and blade locking mechanism for a turboprop according to claim 8, wherein an outer end surface of the stopper section is not higher than an outer end surface of the turbine disk, and a thickness of the stopper section is at least 0.3 mm.
10. The turbine disk and blade locking mechanism for a turboprop according to claim 6, wherein the inlay has a thickness of 0.8-1.2mm and an overall length at least 4mm greater than the thickness of the turbine disk.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202020922849.5U CN212837968U (en) | 2020-05-27 | 2020-05-27 | Turbine disc and blade locking mechanism for turboprop engine |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202020922849.5U CN212837968U (en) | 2020-05-27 | 2020-05-27 | Turbine disc and blade locking mechanism for turboprop engine |
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| CN212837968U true CN212837968U (en) | 2021-03-30 |
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Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN111472845A (en) * | 2020-05-27 | 2020-07-31 | 上海尚实能源科技有限公司 | Turbine disc and blade locking mechanism for turboprop engine |
| CN115791142A (en) * | 2023-02-09 | 2023-03-14 | 中国航发四川燃气涡轮研究院 | Axial limiting blade structure and configuration method |
-
2020
- 2020-05-27 CN CN202020922849.5U patent/CN212837968U/en active Active
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN111472845A (en) * | 2020-05-27 | 2020-07-31 | 上海尚实能源科技有限公司 | Turbine disc and blade locking mechanism for turboprop engine |
| CN115791142A (en) * | 2023-02-09 | 2023-03-14 | 中国航发四川燃气涡轮研究院 | Axial limiting blade structure and configuration method |
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Address after: 201611 3 3 Ting Ting Road, Che Dun Town, Songjiang District, Shanghai Patentee after: Shanghai Shangshi aeroengine Co.,Ltd. Address before: 201611 3 3 Ting Ting Road, Che Dun Town, Songjiang District, Shanghai Patentee before: SHANGHAI SHANGSHI ENERGY TECHNOLOGY CO.,LTD. |
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