CN114810219A - Aircraft engine - Google Patents
Aircraft engine Download PDFInfo
- Publication number
- CN114810219A CN114810219A CN202110126398.3A CN202110126398A CN114810219A CN 114810219 A CN114810219 A CN 114810219A CN 202110126398 A CN202110126398 A CN 202110126398A CN 114810219 A CN114810219 A CN 114810219A
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- Prior art keywords
- mortise
- mounting hole
- rod
- locking block
- aircraft engine
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- 230000007704 transition Effects 0.000 claims description 6
- 238000004891 communication Methods 0.000 claims description 2
- 238000005516 engineering process Methods 0.000 abstract description 2
- 238000003780 insertion Methods 0.000 description 6
- 230000037431 insertion Effects 0.000 description 6
- 238000000034 method Methods 0.000 description 5
- 230000008569 process Effects 0.000 description 5
- 230000008878 coupling Effects 0.000 description 3
- 238000010168 coupling process Methods 0.000 description 3
- 238000005859 coupling reaction Methods 0.000 description 3
- 238000012423 maintenance Methods 0.000 description 2
- 230000008439 repair process Effects 0.000 description 2
- 230000009471 action Effects 0.000 description 1
- 230000000712 assembly Effects 0.000 description 1
- 238000000429 assembly Methods 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 230000006835 compression Effects 0.000 description 1
- 238000007906 compression Methods 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 238000013461 design 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
- 230000013011 mating Effects 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification 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
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention relates to an aircraft engine, comprising: a drum provided with a tongue-and-groove extending in a circumferential direction thereof; the blade comprises a tenon matched with the mortise; the locking block assembly is fixedly installed in the mortise and abutted to the blade to limit the blade to move along the mortise, the locking block assembly comprises a locking block (10) arranged in the mortise and an elastic component (40) configured to enable the locking block (10) to abut against the inner surface of the mortise, and the elastic component (40) pushes and presses the locking block (10) to abut against the inner surface of the mortise, so that the problem that a locking structure in the related technology is prone to failure is solved.
Description
Technical Field
The invention relates to the field of aviation equipment, in particular to an aero-engine.
Background
Referring to fig. 1 to 4, an aircraft engine comprises a drum 2 and a plurality of blades 1 arranged along the circumference of the drum 2, and fig. 1 shows a schematic view of the structure of a meridian projection plane of a typical drum. The common drum barrel 2 is connected with the blade 1 in a tenon mode, a dovetail-shaped mortise is arranged on the circumferential surface of the drum barrel 2 and extends along the circumferential direction of the drum barrel 2, a tenon matched with the mortise is arranged at the root of the blade 1, and the width of an opening at the radial outer end of the mortise is smaller than the width of the inside of the mortise so as to limit the blade 1 to move towards the outer side along the radial direction. The mortises in the drum 2 are provided with insertion openings, in which the tenons of the rotor blades 1 are inserted and moved in the circumferential direction to a predetermined mounting position. The aircraft engine further comprises a locking block assembly 3 which limits circumferential movement of the blade in the mortise in which the locking block assembly is fixedly mounted to define the rotor blade 1 in a mounted position.
Fig. 5 and 6 are schematic structural views of a locking block assembly 3, wherein the locking block assembly 3 comprises a locking block 4, a locking screw 5 and a self-locking screw sleeve 6.
FIGS. 7 and 8 illustrate a schematic view of an alternative lock block assembly; fig. 9 to 12 show a schematic view of the lockblock structure of the lockblock assembly 3. The locking block assembly 3 includes a locking block 4 and a locking screw 5. The locking block 4 comprises a mating face 41 and a working face 42. The locking screw 5 has an external thread 51 and a wrench engaging structure 52 at the end of the locking screw 5, and the wrench engaging structure 52 may be a polygonal hole. The locking block 4 is provided with a threaded hole matched with the locking screw 5, the threaded hole extends along the radial direction of the aircraft engine, and in the process of rotating the locking screw 5 through a wrench, the locking block 4 moves outwards along the radial direction under the driving of the locking screw, so that the working face and/or the matching face of the locking block 4 are abutted against the inner surface of the mortise, and the locking block 4 is fixed in the mortise.
The working principle of the locking block component is as follows:
1. the working surface 42 of the locking block cooperates with the tongue and groove of the drum 2 to bear the centrifugal force load;
2. the locking screws 5 are assembled through the mounting groove structure, the external threads of the locking screws 5 are matched with the internal threads of the locking blocks 4 to radially lift the locking blocks 4, so that the matching surfaces at two ends of the locking blocks 4 are matched with the working grooves of the drum barrel 2, and circumferential positioning is realized;
3. under the action of the two locking blocks 4, the rotor blades are circumferentially positioned and fixed and cannot fall off from the drum barrel, so that the safe operation is ensured.
In order to prevent the locking screw 5 from loosening or falling off in the working process, which causes the rotor blade to fly out, damages other blades, reduces the efficiency of the engine, and can not work normally, the connection of the locking screw 5 must be reliable. In the first structure form, the locking screw 5 is prevented from loosening or falling off by the self-locking torque of the self-locking screw sleeve 6; the second structure form does not contain a self-locking screw sleeve, and the locking block 4 is deformed and punched to provide self-locking torque to meet the requirement in the installation process.
The above-mentioned structural style for preventing the locking screw from falling off mainly has the following defects:
1. the self-locking torque of the first structural form self-locking screw sleeve is easy to lose efficacy and low in reliability;
2. in the second structure form, the locking block is deformed and punched, so that the assembly is complex, and the locking block is easy to damage;
3. in the assembly process of the two structural forms, the locking screw mounting groove is easy to damage when being screwed down due to larger self-locking torque, the assembly time is prolonged, and the cost is increased;
4. the assembly and replacement difficulty is high, and the maintenance and repair cost is increased.
Disclosure of Invention
The invention aims to provide an aircraft engine to improve the problem that a locking structure for limiting the circumferential movement of a blade along a mortise in the related art is prone to failure.
According to an aspect of an embodiment of the present invention, there is provided an aircraft engine, comprising:
a drum provided with a tongue-and-groove extending in a circumferential direction thereof;
the blade comprises a tenon matched with the mortise;
the locking block assembly is fixedly installed in the mortise and abutted against the blade to limit the blade to move along the mortise, and the locking block assembly comprises a locking block arranged in the mortise and an elastic component configured to enable the locking block to abut against the inner surface of the mortise.
In some embodiments, the aircraft engine further comprises a rod-like member, the locking block is provided with a mounting hole adapted to the rod-like member, a first end of the rod-like member is inserted into the mounting hole, the elastic member is configured to push the rod-like member toward an outer side of the locking block, and a second end of the rod-like member abuts against a bottom wall of the mortise.
In some embodiments of the present invention, the,
a flange is arranged at the second end of the rod-shaped component, which is positioned outside the mounting hole;
the mounting holes comprise a first mounting hole and a second mounting hole with the inner diameter smaller than that of the first mounting hole, and a transition step surface is formed at the junction of the first mounting hole and the second mounting hole;
one end of the elastic component is abutted with the transition step surface, and the other end of the elastic component is abutted with the flange.
In some embodiments, the aircraft engine further comprises a limiting member for limiting the distance that the rod-like member moves towards the inside of the mounting hole.
In some embodiments, the stop feature comprises a spring collar.
In some embodiments, the mounting hole is a through hole, and the first of the rod-shaped members is provided with a tool coupling portion for moving the rod-shaped member into the mounting hole when the locking block is mounted in the mortise.
In some embodiments, the tool attachment portion includes a threaded bore provided at the first end of the rod member.
In some embodiments, the tongue and groove includes a tongue and groove body, an outer end of the tongue and groove body in a radial direction of the drum is open, and a width of the opening of the tongue and groove is smaller than a width of at least a portion of the inner cavity of the tongue and groove to restrict the blade from radially disengaging the tongue and groove.
In some embodiments, the mortise further comprises an insertion opening in communication with the mortise body, the insertion opening having a width not less than a maximum width of the tenon.
By applying the technical scheme of the invention, the elastic part pushes the locking block to tightly abut against the inner surface of the mortise, so that the problem that the locking structure is easy to lose efficacy in the related technology is solved.
Other features of the present invention and advantages thereof will become apparent from the following detailed description of exemplary embodiments of the invention, which proceeds with reference to the accompanying drawings.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the related art, the drawings needed to be used in the description of the embodiments or the related art will be briefly introduced below, it is obvious that the drawings in the following description are only some embodiments of the present invention, and for those skilled in the art, other drawings can be obtained according to these drawings without creative efforts.
FIG. 1 illustrates a schematic structural view of a drum and a blade of a related art aircraft engine; and
FIG. 2 shows a side view schematic of the structure of FIG. 1;
FIG. 3 shows an enlarged view of a portion E of FIG. 1;
FIG. 4 shows a schematic cross-sectional view at C-C in FIG. 2;
FIG. 5 is a schematic view of a first type of lock block assembly of a related art aircraft engine;
FIG. 6 shows a schematic cross-sectional view A-A of FIG. 5;
FIG. 7 is a schematic view of a second lock block assembly of a related art aircraft engine;
FIG. 8 shows a schematic cross-sectional view B-B of FIG. 7;
FIG. 9 is a schematic view of a lock block of a second lock block assembly of a related art aircraft engine;
FIG. 10 shows a schematic cross-sectional view at D-D in FIG. 9;
FIG. 11 is a schematic view of a lock block of a second lock block assembly of a related art aircraft engine;
FIG. 12 shows a schematic cross-sectional view at E-E in FIG. 11;
FIG. 13 shows a schematic structural view of a lock block assembly of an aircraft engine of an embodiment of the invention;
FIG. 14 shows a schematic cross-sectional view at F-F in FIG. 11;
FIG. 15 shows a schematic structural view of a locking screw of a locking block assembly of an aircraft engine of an embodiment of the invention;
FIG. 16 is a schematic cross-sectional view of the locking screw of FIG. 15;
FIG. 17 shows a schematic structural view of a lock block assembly (not assembled) of an aircraft engine of an embodiment of the invention; and
fig. 18 shows a schematic cross-sectional structure at H-H in fig. 17.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. The following description of at least one exemplary embodiment is merely illustrative in nature and is in no way intended to limit the invention, its application, or uses. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
The aero-engine of the embodiment comprises a drum and blades, wherein the drum is provided with a tongue-and-groove extending along the circumferential direction of the drum; the blade comprises a tenon matched with the mortise;
the mortise comprises a mortise body, the outer side end of the mortise body along the radial direction of the drum barrel is open, and the width of the open end of the mortise is smaller than the width of at least one part of the inner cavity of the mortise so as to limit the blade to be separated from the mortise along the radial direction.
In some embodiments, the tongue and groove are dovetail grooves. In other embodiments, the tongue groove is generally triangular in cross-section, with the apex of the triangle being further from the axis of the drum than the base opposite the apex.
The mortise comprises a mortise body and an insertion opening communicated with the mortise body, and the width of the insertion opening is not less than the maximum width of the tenon. When the blade is mounted on the drum, the tenon of the blade is first inserted into the mortise from the insertion opening, and then the tenon of the blade is moved in the mortise in the circumferential direction of the drum, thereby moving the blade to a predetermined mounting position.
The aircraft engine also includes a lock block assembly configured to limit movement of the blade in the mortise along a circumferential direction of the drum, the lock block assembly being fixedly mounted in the mortise and abutting the blade to limit movement of the blade along the mortise.
As shown in fig. 13 and 14, the locking block assembly includes a locking block 10 disposed in the mortise and an elastic member 40 configured to make the locking block 10 abut against an inner surface of the mortise.
The aircraft engine further comprises a rod-shaped member 20, the locking block 10 is provided with a mounting hole adapted to the rod-shaped member 20, a first end of the rod-shaped member 20 is inserted into the mounting hole, the elastic member 40 is configured to push the rod-shaped member 20 towards the outside of the locking block 10, and a second end of the rod-shaped member 20 abuts against the bottom wall of the mortise.
As shown in fig. 14 to 18, the second end of the rod member 20 located outside the mounting hole is provided with a flange 60; the mounting holes comprise a first mounting hole and a second mounting hole with the inner diameter smaller than that of the first mounting hole, and a transition step surface is formed at the junction of the first mounting hole and the second mounting hole; the elastic member 40 has one end abutting the stepped transition surface and the other end abutting the flange 60. In this embodiment, the elastic member 40 is a spring fitted over the rod member 20.
The aircraft engine further includes a stopper member 30 for limiting the distance that the rod member 20 moves toward the inside of the mounting hole.
In some embodiments, the stop 30 comprises a spring collar.
The mounting hole is a through hole, and the first end of the rod member 20 is provided with a tool coupling portion 50 to move the rod member 20 into the mounting hole when the locking block 10 is mounted in the mortise.
In some embodiments, the tool attachment portion 50 includes a threaded bore provided at a first end of the rod member 20.
In the present embodiment, the locking block assembly includes a locking block 10, a rod member 20, an elastic member 40, and a stopper member 30. As shown in fig. 15 and 16, the first end of the rod member 20 is internally threaded to form the tool coupling portion 50, and the first end of the rod member 20 is free of an external threaded configuration.
The assembly process of this patent structure:
1. before assembly, the spring collar (limiting part 30) is removed, the rod-shaped part is lifted upwards by matching a tool with the internal thread of the rod-shaped part 20, and the spring is in a compressed state, as shown in fig. 18;
2. and after the locking block assembly and the drum barrel are assembled, removing the tool, assembling the spring retainer ring, and completing assembly.
The working principle of the patent is as follows:
1. in the assembling process, the radial distance between the rod-shaped component and the locking block is reduced by compressing the spring, and the locking block component moves in the drum barrel to realize assembling;
2. under the working condition, the spring is in a compression state, radial force is provided to enable the locking block to be tightly matched with the drum barrel, the spring retainer ring further limits the radial displacement of the rod-shaped component, normal work is guaranteed, and the circumferential positioning requirement of the rotor blade is met.
The blade locking block assembly is simple in structure and device, easy to install and high in reliability, and can well meet the requirements of structure assembly and positioning. It is mainly characterized by the design of spring and spring retainer ring. The embodiment of the patent refers to the field of 'constructional details of locking-block assemblies, spring arrangements and quantities, spring collars, locking-block arrangements and the like'.
The locking block assembly of the embodiment has the following technical effects:
the embodiment provides a novel rotor blade locking block subassembly, and this locking block subassembly has following beneficial effect:
1. the structure is simple, the failure rate of the locking block assembly is reduced, and the reliability of the engine is improved;
2. the assembly is easy, the assembly time is saved, and the assembly cost is reduced;
3. the locking block and the rod-shaped component are not easy to damage, and the maintenance and repair cost is low.
The present invention is not limited to the above exemplary embodiments, and any modifications, equivalent replacements, improvements, etc. within the spirit and principle of the present invention should be included in the protection scope of the present invention.
Claims (9)
1. An aircraft engine, comprising:
a drum provided with a tongue-and-groove extending in a circumferential direction thereof;
the blade comprises a tenon matched with the mortise;
the locking block assembly is fixedly installed in the mortise and abutted to the blade to limit the blade to move along the mortise, and comprises a locking block (10) arranged in the mortise and an elastic component (40) configured to enable the locking block (10) to abut against the inner surface of the mortise.
2. An aircraft engine according to claim 1, characterised in that it further comprises a rod-shaped part (20), the locking block (10) being provided with a mounting hole adapted to the rod-shaped part (20), into which mounting hole a first end of the rod-shaped part (20) is inserted, the elastic part (40) being configured to push the rod-shaped part (20) towards the outside of the locking block (10), a second end of the rod-shaped part (20) abutting against the bottom wall of the mortise.
3. The aircraft engine of claim 2,
a second end of the rod-shaped member (20) located outside the mounting hole is provided with a flange (60);
the mounting holes comprise a first mounting hole and a second mounting hole with the inner diameter smaller than that of the first mounting hole, and a transition step surface is formed at the junction of the first mounting hole and the second mounting hole;
one end of the elastic component (40) is abutted against the transition step surface, and the other end of the elastic component is abutted against the flange (60).
4. An aircraft engine according to claim 2, further comprising a limiting member (30) for limiting the distance of movement of the rod-like member (20) towards the inside of the mounting hole.
5. An aircraft engine according to claim 4, characterised in that the stop means (30) comprise a spring collar.
6. An aircraft engine according to claim 4, characterised in that the mounting hole is a through hole and the first of the rod-shaped members (20) is provided with a tool attachment portion (50) to move the rod-shaped member (20) into the mounting hole when the locking block (10) is mounted in the mortise slot.
7. An aircraft engine according to claim 6, characterised in that the tool attachment portion (50) comprises a threaded hole provided at a first end of the rod-like member (20).
8. The aircraft engine of claim 1, wherein said tongue and groove includes a tongue and groove body, an outer end of said tongue and groove body in a radial direction of said drum being open, a width of said tongue and groove opening being less than a width of at least a portion of said tongue and groove internal cavity to limit radial escape of said blade from said tongue and groove.
9. The aircraft engine of claim 8 wherein said mortise slot further comprises an insert opening in communication with said mortise slot body, said insert opening having a width no less than a maximum width of said tenon.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110126398.3A CN114810219A (en) | 2021-01-29 | 2021-01-29 | Aircraft engine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110126398.3A CN114810219A (en) | 2021-01-29 | 2021-01-29 | Aircraft engine |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN114810219A true CN114810219A (en) | 2022-07-29 |
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Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202110126398.3A Pending CN114810219A (en) | 2021-01-29 | 2021-01-29 | Aircraft engine |
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| CN (1) | CN114810219A (en) |
Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN1959066A (en) * | 2005-11-03 | 2007-05-09 | 通用电气公司 | Damper seal system and damping inner cover ring vibration method |
| CN101892866A (en) * | 2009-05-20 | 2010-11-24 | 通用电气公司 | The low stress circumferential dovetail attachment that is used for rotor blade |
| 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 |
| CN106014490A (en) * | 2016-06-22 | 2016-10-12 | 中国航空工业集团公司沈阳发动机设计研究所 | Rotor vane locking structure |
| 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 |
-
2021
- 2021-01-29 CN CN202110126398.3A patent/CN114810219A/en active Pending
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN1959066A (en) * | 2005-11-03 | 2007-05-09 | 通用电气公司 | Damper seal system and damping inner cover ring vibration method |
| CN101892866A (en) * | 2009-05-20 | 2010-11-24 | 通用电气公司 | The low stress circumferential dovetail attachment that is used for rotor blade |
| 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 |
| CN106014490A (en) * | 2016-06-22 | 2016-10-12 | 中国航空工业集团公司沈阳发动机设计研究所 | Rotor vane locking structure |
| 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 |
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