CN116877213A - A dish edge cone shell clamping stagnation structure for preventing aeroengine turbine is changeed - Google Patents
A dish edge cone shell clamping stagnation structure for preventing aeroengine turbine is changeed Download PDFInfo
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- CN116877213A CN116877213A CN202311074068.XA CN202311074068A CN116877213A CN 116877213 A CN116877213 A CN 116877213A CN 202311074068 A CN202311074068 A CN 202311074068A CN 116877213 A CN116877213 A CN 116877213A
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- baffle
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- edge
- ring surface
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- 241000237942 Conidae Species 0.000 title claims abstract description 48
- 238000009434 installation Methods 0.000 claims description 3
- 238000006073 displacement reaction Methods 0.000 claims description 2
- 239000003921 oil Substances 0.000 description 5
- 238000002485 combustion reaction Methods 0.000 description 3
- 239000002737 fuel gas Substances 0.000 description 3
- 230000009977 dual effect Effects 0.000 description 2
- 239000007789 gas Substances 0.000 description 2
- 238000004519 manufacturing process Methods 0.000 description 2
- 238000000034 method Methods 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 1
- 230000015556 catabolic process Effects 0.000 description 1
- 238000010276 construction Methods 0.000 description 1
- 238000005520 cutting process Methods 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 230000002708 enhancing effect Effects 0.000 description 1
- 239000012634 fragment Substances 0.000 description 1
- 239000000446 fuel Substances 0.000 description 1
- 239000000295 fuel oil Substances 0.000 description 1
- 239000000463 material Substances 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 238000003825 pressing Methods 0.000 description 1
- 230000002265 prevention Effects 0.000 description 1
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
- F01D21/00—Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for
- F01D21/04—Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for responsive to undesired position of rotor relative to stator or to breaking-off of a part of the rotor, e.g. indicating such position
- F01D21/045—Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for responsive to undesired position of rotor relative to stator or to breaking-off of a part of the rotor, e.g. indicating such position special arrangements in stators or in rotors dealing with breaking-off of part of rotor
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention provides a disc edge cone shell clamping structure for preventing an aeroengine turbine from over-rotating, which comprises a disc edge baffle and a cone shell baffle, wherein the disc edge baffle is arranged on the disc edge cone shell; the disk edge baffle is arranged at the rear side of the turbine disk, and the head of the disk edge baffle is provided with a petal-shaped outer ring surface; the cone shell baffle is arranged on the oil cavity bushing, and the shoulder part of the cone shell baffle is provided with a petal-shaped inner ring surface. In the mounting state and the normal left-side state of the engine, the outer ring surface of the head part of the disc edge baffle plate is opposite to the front and back of the inner ring surface of the shoulder part of the cone shell baffle plate, and a certain distance is reserved between the outer ring surface and the inner ring surface in the axial direction; once the rotating shaft breaks down, the turbine rotor moves backwards, the petal-shaped outer ring surface of the disc edge baffle head and the petal-shaped inner ring surface of the cone shell baffle shoulder are contacted and clamped, friction and severe deformation are generated, the kinetic energy of the rotor is absorbed, the rotating speed of the turbine rotor is limited, rapid braking is realized, and the integrity and safety of the turbine rotor are ensured.
Description
Technical Field
The invention belongs to the field of aeroengine design, and particularly relates to a disk edge cone shell clamping stagnation structure for preventing an aeroengine turbine from overturning.
Background
Due to manufacturing and assembly errors, or extreme loads, material defects, etc., the aeroengine may experience shaft breakage failure during operation, which may cause over-rotation of the turbine rotor under high energy gas drive or even rupture of the turbine disc, resulting high energy fragments and possible breakdown of the nacelle and even the aircraft body, with catastrophic consequences. Therefore, it is necessary to perform turbine over-rotation prevention design, and to perform down-rotation on the turbine rotor after the shaft breakage occurs.
At present, the turbine of an aeroengine is researched for preventing over-rotation, and most of the turbine is characterized in that the fuel oil supply of a combustion chamber is cut off after a rotating shaft is broken, so that a turbine rotor is lack of high-energy fuel gas driving to decelerate. However, the engine control system still requires a certain amount of time to monitor and determine the shaft breakage event and shut off the fuel supply, during which the rotational speed of the turbine rotor increases rapidly. Therefore, the method of actively cutting off oil supply is independently relied on to prevent the rotation speed of the turbine rotor, which is still a very demanding requirement, and it is necessary to design and install an over-rotation protection structure of the turbine rotor, and to rapidly reduce the rotation speed of the turbine rotor after the breakage of the rotating shaft occurs.
Disclosure of Invention
In order to solve the technical problems, the invention aims to provide a disk edge cone shell clamping structure for preventing an aeroengine turbine from over-rotating, which comprises a disk edge baffle and a cone shell baffle; the disk edge baffle is arranged at the rear side of the turbine disk, and the head of the disk edge baffle is provided with a petal-shaped outer ring surface; the cone shell baffle is arranged on the oil cavity bushing, and the shoulder part of the cone shell baffle is provided with a petal-shaped inner ring surface. In the mounting state and the normal left-side state of the engine, the outer ring surface of the head part of the disc edge baffle plate is opposite to the front and back of the inner ring surface of the shoulder part of the cone shell baffle plate, and a certain distance is reserved between the outer ring surface and the inner ring surface in the axial direction; once the rotating shaft breaks down, the turbine rotor moves backwards, the petal-shaped outer ring surface of the disc edge baffle head and the petal-shaped inner ring surface of the cone shell baffle shoulder are contacted and clamped, friction and severe deformation are generated, the kinetic energy of the rotor is absorbed, the rotating speed of the turbine rotor is limited, rapid braking is realized, and the integrity and safety of the turbine rotor are ensured. The invention can limit the rotation speed of the turbine after the rotation shaft of the engine is broken, so as to avoid the catastrophic accident of turbine disc breakage; meanwhile, the disc edge conical shell clamping structure has the advantages of being simple in configuration, convenient to assemble and disassemble and capable of achieving rapid braking.
In order to achieve the above purpose, the present invention adopts the following technical scheme:
a disc edge cone shell clamping structure for preventing an aeroengine turbine from over-rotating comprises a disc edge baffle and a cone shell baffle; the disc edge conical shell clamping structure is arranged at the position of the low-pressure turbine component, and limits the rotating speed of the low-pressure turbine rotor component after the low-pressure rotating shaft is broken, so that the over-rotation of the low-pressure turbine rotor component and the rupture of the low-pressure turbine disc are avoided;
the disc edge baffle is provided with a bottom mounting edge, a head axial end surface and a top outer annular surface, wherein the outer annular surface is provided with a certain taper and is wavy, and first pits and first bosses are uniformly distributed on the outer annular surface at intervals in the circumferential direction;
the cone shell baffle is provided with a bottom mounting edge, a head joint edge and a shoulder inner ring surface, wherein the inner ring surface is provided with a certain taper, but smaller than the outer ring surface, the inner ring surface is wavy, and second pits and second bosses are uniformly distributed on the inner ring surface at intervals in the circumferential direction;
the disc edge baffle is connected with the turbine disc mounting edge through a second bolt at the position of the bottom mounting edge, so that self fixed mounting is realized; the position of the disc edge baffle on the axial end face of the head part is axially pressed with the rear end face of the turbine blade disc joggle joint structure, so that the axial displacement of the low-pressure turbine blade is limited;
the conical shell baffle is connected with the head mounting edge of the bearing oil cavity through a third bolt at the bottom mounting edge, so that self fixed mounting is realized; the head joint edge of the cone shell baffle is in joint with the bearing inner ring of the rear frame, so that the deformation resistance of the cone shell baffle is enhanced;
the top outer ring surface of the disc edge baffle is opposite to the inner ring surface of the shoulder part of the cone shell baffle in front and back in the installation state, and a certain distance exists between the top outer ring surface and the inner ring surface in the axial direction.
Further, when the aeroengine works normally, a certain axial distance is kept between the outer annular surface of the top of the disc edge baffle and the inner annular surface of the shoulder part of the cone shell baffle, so that contact friction is avoided;
when the low-pressure shaft breaks, the low-pressure turbine rotor component moves backwards, the disc edge baffle moves backwards along with the low-pressure turbine disc, the outer annular surface of the top of the disc edge baffle presses against the inner annular surface of the shoulder of the cone-shell baffle, the first boss is embedded into the second pit to form clamping stagnation, the cone-shell baffle is caused to deform severely, kinetic energy of the low-pressure turbine rotor component is consumed, the rotor is limited to rise, and quick braking is realized, so that the integrity of the low-pressure turbine component is ensured.
Further, the wavy outer annular surface at the top of the disc edge baffle and the wavy inner annular surface at the shoulder of the cone shell baffle are both conical surfaces, the cone angle of the outer annular surface is larger than that of the inner annular surface, and the larger the backward movement amount of the low-pressure turbine rotor component is, the tighter the clamping stagnation between the outer annular surface and the inner annular surface is, and the stronger the rotation speed limitation on the low-pressure turbine rotor component is.
Further, the method is applied to the double-rotor aero-engine.
Compared with the prior art, the invention has the beneficial effects that:
the disc edge conical shell clamping stagnation structure for preventing the turbine of the aeroengine from over-rotating can rapidly limit the speed of the turbine after the rotating shaft of the aeroengine is broken, so that catastrophic accidents such as turbine over-rotating and turbine breaking are avoided, and the integrity of the turbine is ensured.
The disk edge conical shell clamping structure for preventing the turbine of the aero-engine from over-rotating has the advantages of simple structure configuration, small influence on the existing structure of the engine, easiness in processing and low cost.
The disk edge conical shell clamping stagnation structure for preventing the turbine of the aeroengine from over-rotating not only can be used for a low-pressure turbine, but also can be used for a high-pressure turbine, and can be applied to a double-rotor engine, and also can be applied to aeroengines with other structural layouts such as single rotors or double rotors.
Drawings
The above and other features, properties and advantages of the present invention will be more clearly described by the following description in conjunction with the accompanying drawings and examples, in which:
FIG. 1 is a schematic structural view of a typical dual rotor aircraft engine.
FIG. 2 is a partial schematic view of a low pressure turbine component using an example of the present invention.
FIG. 3 is a partial schematic view of a rim cone-shell clamping structure using an example of the present invention.
Fig. 4 is a partial schematic view of a rim plate using an example of the present invention.
Fig. 5 is a partial schematic view of a cone shell baffle using one example of the present invention.
Detailed Description
The invention is described in detail below with reference to the drawings and examples.
FIG. 1 shows a schematic diagram of a typical dual rotor aircraft engine 100. The twin-spool aeroengine 100 includes a low-pressure compressor component 11, a high-pressure compressor component 12, a combustion chamber 13, a high-pressure turbine component 14, a low-pressure turbine component 15, and the like.
Wherein the low pressure compressor part 11 comprises a low pressure compressor rotor part 11A and a low pressure compressor stator part 11B, the high pressure compressor part 12 comprises a high pressure compressor rotor part 12A and a high pressure compressor stator part 12B, the high pressure turbine part 14 comprises a high pressure turbine rotor part 14A and a high pressure turbine stator part 14B, and the low pressure turbine part 15 comprises a low pressure turbine rotor part 15A and a low pressure turbine stator part 15B.
The high pressure compressor rotor component 12A is driven by a high pressure turbine rotor component 14A, which are connected by a high pressure shaft 16; the low pressure compressor rotor part 11A is driven by a low pressure turbine rotor part 15A, which is connected by a low pressure shaft 17.
The low-pressure compressor rotor part 11A is supported by the first roller bearing 1 and the first ball bearing 2, the high-pressure compressor rotor part 12A is supported by the second ball bearing 3, and axial force and radial force on the three bearings are transmitted outwards through the front bearing frame 18; the high pressure turbine rotor component 14A is supported by the second roller bearing 4, the radial forces on the bearing being transmitted outwards through the high pressure turbine aft frame 19; the low pressure turbine rotor part 15A is supported by a third roller bearing 5, the radial forces on which are transmitted outwards through the low pressure turbine aft frame 32.
During operation of the dual-rotor aircraft engine 100, high-temperature and high-energy fuel gas is discharged from the combustion chamber 13, and impacts the high-pressure turbine part 14 and the low-pressure turbine part 15 backward (in the direction of arrow a), and drives the low-pressure turbine part 15A and the high-pressure turbine part 14A to rotate, and further, the high-pressure turbine part 14A drives the high-pressure compressor part 12A to rotate, and the low-pressure turbine part 15A drives the low-pressure compressor part 11A to rotate.
When a fracture fault occurs in the low-pressure shaft 17, for example, at the fracture position 171, on the one hand, the low-pressure turbine rotor component 15A suddenly throws away the load of the low-pressure compressor rotor component 11A, and the rotation speed rapidly rises under the continuous driving of the high-energy fuel gas in the flow passage; on the other hand, since the breaking point 171 is located behind the first ball bearing 2, causing the low pressure shaft 17 to lose axial constraint, the low pressure shaft 17 and the low pressure turbine rotor component 15A will move axially rearward (in the direction of arrow a) under the urging of the high energy gas.
In fig. 2, the low pressure turbine component 15 of an aircraft engine 101 utilizes a rim cone and shell stuck structure 102 of the present invention that prevents over-rotation of the aircraft engine turbine. Aero-engine 101 is similar in structure to birotor aero-engine 100, and component numbers for birotor aero-engine 100 are used herein.
The low pressure turbine component 15 includes a low pressure turbine rotor component 15A and a low pressure turbine stator component 15B. The low pressure turbine rotor component 15A mainly comprises a low pressure turbine disk 21 and low pressure turbine blades 22, the low pressure turbine blades 22 are mounted on the low pressure turbine disk 21 through a joggle structure 23, a pressing disc 24 is mounted on the front side of the joggle structure 23, and a disc edge baffle 25 is mounted on the rear side of the joggle structure. Wherein the hold-down disc 24 is mounted on the low pressure turbine disc 21 by a first bolt 26 and the rim retainer 25 is mounted on the low pressure turbine disc 21 by a second bolt 27.
The low-pressure turbine stator part 15B mainly includes a bearing housing 31 for the third roller bearing 5, a low-pressure turbine rear frame 32, and a bearing oil chamber 33.
Fig. 3-5 illustrate the specific construction of an example of the invention and its manner of installation in a low pressure turbine component 15.
The rim retainer 25 has a bottom mounting edge 251, a head axial end face 252 and a top outer annular surface 253, wherein the top outer annular surface 253 has a taper and is wave-shaped with first pockets 254 and first bosses 255 circumferentially uniformly spaced therearound.
Cone shell baffle 35 has a bottom mounting edge 351, a head overlap edge 352, and a shoulder inner annular surface 353, wherein shoulder inner annular surface 353 has a taper, but a taper less than that of top outer annular surface 253; the shoulder inner annular surface 353 has a wave shape and is provided with second recesses 354 and second bosses 355 circumferentially and uniformly spaced apart.
The disk edge baffle 25 is connected with the turbine disk mounting edge 211 at the position of the bottom mounting edge 251 through a second bolt 27, so that the self-fixing mounting is realized; the disk edge baffle 25 is compressed against the aft end 231 of the dovetail 23 at the location of the head axial end 252 to limit axial movement of the low pressure turbine blade 22.
The cone shell baffle 35 is connected with the head mounting edge 331 of the bearing oil cavity 33 at the bottom mounting edge 351 through a third bolt 34, so that self fixed mounting is realized; the head overlap edge 352 of the cone shell baffle 35 also overlaps the load bearing inner ring 321 of the low pressure turbine aft frame 32, enhancing the deformation resistance of the cone shell baffle 35.
After the assembly is completed, the top outer annular surface 253 of the rim retainer 25 is axially spaced from the shoulder inner annular surface 353 of the cone retainer 35 by a distance.
The working principle of the invention is as follows:
during normal operation of aircraft engine 101, top outer annular surface 253 of rim retainer 25 is axially spaced from shoulder inner annular surface 353 of cone retainer 35 so that no contact friction occurs.
After the low pressure shaft 17 breaks, the low pressure turbine rotor component 15A moves backward, the disc edge baffle 25 moves backward along with the low pressure turbine disc 21, the top outer annular surface 253 of the disc edge baffle moves backward along with the low pressure turbine disc 21, and the shoulder inner annular surface 353 of the cone shell baffle 35 is pressed by the first boss 255, at this time, the second boss 354 is embedded into the second pit 354 to form clamping stagnation, and the cone shell baffle 35 is caused to deform severely, so that the kinetic energy of the low pressure turbine rotor component 15A is consumed, the rotating speed of the low pressure turbine rotor component 15A is limited, rapid braking is realized, and the integrity of the low pressure turbine component 15 is ensured.
The top outer annular surface 253 of the rim retainer 25 and the shoulder inner annular surface 353 of the cone shell retainer 35 are both tapered surfaces, and the cone angle of the top outer annular surface 253 is larger than the shoulder inner annular surface 353, so that the higher the amount of aft movement of the low pressure turbine rotor component 15A, the tighter the clamping between the top outer annular surface 253 and the shoulder inner annular surface 353, and the stronger the rotational speed limitation of the low pressure turbine rotor component 15A.
The disc edge conical shell clamping stagnation structure 102 for preventing the turbine of the aeroengine from over-rotating has the advantages of simple structure configuration, easiness in processing and manufacturing, low cost and small influence on the existing structure of the engine, and can be used for a low-pressure turbine and a high-pressure turbine, and can be applied to a double-rotor engine and an aeroengine with other structural layouts such as a single rotor or a three-rotor.
The above examples are only for illustrating the technical scheme of the present invention and are not limited thereto. While the invention has been described in detail with reference to the preferred embodiments, those skilled in the art will appreciate that: modifications may be made to the specific embodiments of the present invention or equivalents may be substituted for part of the technical features thereof; without departing from the spirit of the technical proposal of the invention, the invention is covered by the scope of the technical proposal of the invention.
Claims (4)
1. A dish edge cone shell clamping stagnation structure for preventing aeroengine turbine is changeed, its characterized in that: comprises a disc edge baffle and a cone shell baffle; the disc edge conical shell clamping structure is arranged at the low-pressure turbine part, and limits the rotating speed of the low-pressure turbine rotor part after the low-pressure rotating shaft is broken, so that the over-rotation of the low-pressure turbine rotor part and the over-rotation of the low-pressure turbine disc are avoided;
the disc edge baffle is provided with a bottom mounting edge, a head axial end surface and a top outer annular surface, wherein the outer annular surface is provided with a certain taper and is wavy, and first pits and first bosses are uniformly distributed on the outer annular surface at intervals in the circumferential direction;
the cone shell baffle is provided with a bottom mounting edge, a head joint edge and a shoulder inner ring surface, wherein the inner ring surface is provided with taper, but smaller than the outer ring surface in taper, the inner ring surface is wavy, and second pits and second bosses are uniformly distributed on the inner ring surface at intervals in the circumferential direction;
the disc edge baffle is connected with the turbine disc mounting edge through a second bolt at the position of the bottom mounting edge, so that self fixed mounting is realized; the position of the disc edge baffle on the axial end face of the head part is axially pressed with the rear end face of the turbine blade disc joggle joint structure, so that the axial displacement of the low-pressure turbine blade is limited;
the conical shell baffle is connected with the head mounting edge of the bearing oil cavity through a third bolt at the bottom mounting edge, so that self fixed mounting is realized; the head joint edge of the cone shell baffle is in joint with the bearing inner ring of the rear frame, so that the deformation resistance of the cone shell baffle is enhanced;
the top outer ring surface of the disc edge baffle is opposite to the inner ring surface of the shoulder part of the cone shell baffle in front and back in the installation state, and a certain distance exists between the top outer ring surface and the inner ring surface in the axial direction.
2. A disk rim cone shell stuck structure for preventing over-rotation of an aircraft engine turbine as claimed in claim 1, wherein: when the aeroengine works normally, a certain axial distance is kept between the outer annular surface of the top of the disc edge baffle and the inner annular surface of the shoulder part of the cone shell baffle, so that contact friction is avoided;
when the low-pressure shaft breaks, the low-pressure turbine rotor component moves backwards, the disc edge baffle moves backwards along with the low-pressure turbine disc, the outer annular surface of the top of the disc edge baffle presses against the inner annular surface of the shoulder of the cone-shell baffle, the first boss is embedded into the second pit to form clamping stagnation, the cone-shell baffle is caused to deform severely, kinetic energy of the low-pressure turbine rotor component is consumed, the rotor is limited to rise, and quick braking is realized, so that the integrity of the low-pressure turbine component is ensured.
3. A rim cone shell stuck structure for preventing over-rotation of an aero-engine turbine according to claim 1 or 2, wherein: the wavy outer annular surface at the top of the disc edge baffle and the wavy inner annular surface at the shoulder of the cone shell baffle are both conical surfaces, the cone angle of the outer annular surface is larger than that of the inner annular surface, and the higher the backward movement amount of the low-pressure turbine rotor component is, the tighter the clamping stagnation between the outer annular surface and the inner annular surface is, and the stronger the rotation speed limitation on the low-pressure turbine rotor component is.
4. A rim cone shell stuck structure for preventing over-rotation of an aero-engine turbine according to claim 1 or 2, wherein: the bearing frame is applied to an aeroengine with a bearing frame behind the last-stage turbine rotor.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202311074068.XA CN116877213A (en) | 2023-08-24 | 2023-08-24 | A dish edge cone shell clamping stagnation structure for preventing aeroengine turbine is changeed |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202311074068.XA CN116877213A (en) | 2023-08-24 | 2023-08-24 | A dish edge cone shell clamping stagnation structure for preventing aeroengine turbine is changeed |
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| Publication Number | Publication Date |
|---|---|
| CN116877213A true CN116877213A (en) | 2023-10-13 |
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| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202311074068.XA Pending CN116877213A (en) | 2023-08-24 | 2023-08-24 | A dish edge cone shell clamping stagnation structure for preventing aeroengine turbine is changeed |
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| CN (1) | CN116877213A (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN117848709A (en) * | 2024-03-08 | 2024-04-09 | 成都晨发泰达航空科技股份有限公司 | Device and method for testing turbine rotor over-rotation of ultra-high revolution aero-engine |
-
2023
- 2023-08-24 CN CN202311074068.XA patent/CN116877213A/en active Pending
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN117848709A (en) * | 2024-03-08 | 2024-04-09 | 成都晨发泰达航空科技股份有限公司 | Device and method for testing turbine rotor over-rotation of ultra-high revolution aero-engine |
| CN117848709B (en) * | 2024-03-08 | 2024-05-14 | 成都晨发泰达航空科技股份有限公司 | Device and method for testing turbine rotor over-rotation of ultra-high revolution aero-engine |
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