CN114776402A - Bearing cavity sealing system of aircraft engine and control method thereof - Google Patents
Bearing cavity sealing system of aircraft engine and control method thereof Download PDFInfo
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- CN114776402A CN114776402A CN202110089204.7A CN202110089204A CN114776402A CN 114776402 A CN114776402 A CN 114776402A CN 202110089204 A CN202110089204 A CN 202110089204A CN 114776402 A CN114776402 A CN 114776402A
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- flow path
- exhaust
- bearing cavity
- bleed air
- aircraft engine
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- 238000007789 sealing Methods 0.000 title claims abstract description 49
- 238000000034 method Methods 0.000 title claims abstract description 13
- 238000004891 communication Methods 0.000 claims description 12
- 239000010687 lubricating oil Substances 0.000 abstract description 3
- 230000001276 controlling effect Effects 0.000 description 3
- 238000010586 diagram Methods 0.000 description 2
- 238000012986 modification Methods 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 238000001816 cooling Methods 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 230000007613 environmental effect Effects 0.000 description 1
- 230000014509 gene expression Effects 0.000 description 1
- 230000001105 regulatory effect Effects 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
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Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/16—Arrangement of bearings; Supporting or mounting bearings in casings
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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
- F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
- F01D11/02—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type
- F01D11/04—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type using sealing fluid, e.g. steam
- F01D11/06—Control thereof
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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
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
The invention discloses a bearing cavity sealing system of an aircraft engine and a control method thereof. The bearing cavity sealing system of the aircraft engine comprises a bleed air flow path and at least two exhaust flow paths, wherein the bleed air flow path is used for leading sealing air from a compressor into a middle bearing cavity and a rear bearing cavity, the exhaust back pressures of the at least two exhaust flow paths are different, and the bleed air flow path is selectively communicated with at least one of the at least two exhaust flow paths. According to the sealing system for the bearing cavity of the aircraft engine, the plurality of exhaust flow paths with different exhaust back pressures are arranged, so that the exhaust back pressure of the bleed air flow path can be controlled according to the running state of the aircraft engine, and the flow of the sealed gas can be further adjusted. When the aero-engine is in a high Mach number pneumatic parameter state, the exhaust back pressure is reduced, the horizontal distribution of the on-way pressure of the bleed air flow path is further reduced, the sealing pressure difference of the bearing cavity is effectively reduced, and the consumption of lubricating oil is reduced.
Description
Technical Field
The invention relates to a bearing cavity sealing system of an aircraft engine and a control method thereof.
Background
In order to ensure the sealing pressure difference of the aircraft engine in a low-Mach-number state on the ground or high altitude, the bleed air at a relatively higher-level position of an engine compressor needs to be taken, but the sealing pressure difference of the aircraft engine in a high-Mach-number pneumatic parameter state is larger, and the increase of the sealing pressure difference brings the risk of exceeding the pressure difference limit value and increases the consumption of lubricating oil. If bleed air at a relatively low-level position of an engine compressor is led in order to reduce the risk of over-limit of the sealing pressure difference and reduce the consumption of lubricating oil, the problem of insufficient sealing pressure difference of the aeroengine in a low-Mach number pneumatic parameter state on the ground or at high altitude can be caused. Therefore, the design of the sealed flow path has a contradiction.
In addition, the air flow of the bearing sealing flow path is small, if the flow path simultaneously plays a role in cooling the high-pressure compressor disk core and flows through the high-pressure compressor disk core, and in a high-altitude and large-mach-number state, the high-speed rotation of the disk shaft can bring high wind resistance temperature rise to the sealing air flow, so that great risk is brought to the over-temperature of the sealing air temperature.
Disclosure of Invention
The invention aims to provide a bearing cavity sealing system of an aircraft engine and a control method thereof, so as to reduce the temperature of a sealing airflow in a high-altitude high-Mach number state.
The invention provides a bearing cavity sealing system of an aircraft engine, which comprises:
the air guide flow path is used for guiding the sealed air from the air compressor to the middle bearing cavity and the rear bearing cavity; and
at least two exhaust gas flow paths having different exhaust backpressure, the bleed air flow path being selectively in communication with at least one of the at least two exhaust gas flow paths.
In some embodiments, the at least two exhaust gas flow paths comprise:
a first exhaust flow path communicating with the low pressure turbine outlet; and
a second exhaust flow path in communication with the atmosphere, and a bleed air flow path configured to communicate with at least one of the first and second exhaust flow paths.
In some embodiments, an opening degree adjusting member is provided on the second exhaust flow path, the opening degree adjusting member being configured to adjust a flow area of the second exhaust flow path.
In some embodiments, the bearing cavity sealing system includes an exhaust pipe, a first end of the exhaust pipe is communicated with the bleed air flow path, a second end of the exhaust pipe extends to the atmosphere, an inner cavity of the exhaust pipe forms a second exhaust flow path, and the opening degree adjusting piece is arranged on the exhaust pipe.
In some embodiments, the opening adjuster is an on-off control valve.
In some embodiments, the opening adjuster is an opening adjustment valve.
In some embodiments, the bleed air flowpath passes sequentially through the compressor hub and the air conduit into the center and aft bearing cavities.
The second aspect of the invention provides a control method of a bearing cavity sealing system based on the aircraft engine, which comprises the following steps: and controlling the bleed air flow path to be communicated with at least one of the at least two exhaust gas flow paths according to the running state of the aircraft engine.
In some embodiments, the at least two exhaust gas flow paths include a first exhaust gas flow path in communication with the low pressure turbine outlet and a second exhaust gas flow path in communication with the atmosphere, and controlling the bleed air flow path to communicate with at least one of the at least two exhaust gas flow paths based on operating conditions of the aircraft engine includes:
controlling the bleed air flow path to communicate with both the first exhaust flow path and the second exhaust flow path in a first state; and
and in the second state, the bleed air flow path is controlled to be communicated with the first exhaust flow path and the bleed air flow path is controlled to be disconnected from the second exhaust flow path.
In some embodiments, the opening of the second exhaust flow path is controlled in a first state to adjust the airflow rate of the bleed air flow path.
Based on the aspects provided by the invention, the bearing cavity sealing system of the aircraft engine comprises a bleed air flow path and at least two exhaust flow paths, wherein the bleed air flow path is used for guiding sealing gas from a compressor to the middle bearing cavity and the rear bearing cavity, the exhaust back pressures of the at least two exhaust flow paths are different, and the bleed air flow path is selectively communicated with at least one of the at least two exhaust flow paths. According to the sealing system for the bearing cavity of the aircraft engine, the plurality of exhaust flow paths with different exhaust back pressures are arranged, so that the exhaust back pressure of the bleed air flow path can be controlled according to the running state of the aircraft engine, and the flow of the sealed gas can be further adjusted. For example, when the aircraft engine is in a high mach number aerodynamic parameter state, the bleed air flow path is controlled to be communicated with the first exhaust flow path and the second exhaust flow path, so that exhaust back pressure is reduced, the flow of the seal gas is increased, and the risk of over-temperature of a compressor disk or a shaft caused by high-speed rotation is effectively reduced. And the reduction of the exhaust back pressure can further reduce the horizontal distribution of the on-way pressure of the bleed air flow path, thereby effectively reducing the sealing pressure difference of the bearing cavity.
Other features of the present invention and advantages thereof will become apparent from the following detailed description of exemplary embodiments thereof, which proceeds with reference to the accompanying drawings.
Drawings
The accompanying drawings, which are included to provide a further understanding of the invention and are incorporated in and constitute a part of this application, illustrate embodiment(s) of the invention and together with the description serve to explain the invention without limiting the invention. In the drawings:
FIG. 1 is a schematic flow diagram of a bearing cavity sealing system of an aircraft engine according to an embodiment of the invention;
FIG. 2 is a schematic view of the bleed air flow path of FIG. 1;
FIG. 3 is a schematic view of the exhaust flow path of FIG. 1.
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 relative arrangement of the components and steps, the numerical expressions and numerical values set forth in these embodiments do not limit the scope of the present invention unless specifically stated otherwise. Meanwhile, it should be understood that the sizes of the respective portions shown in the drawings are not drawn in an actual proportional relationship for the convenience of description. Techniques, methods, and apparatus known to one of ordinary skill in the relevant art may not be discussed in detail, but are intended to be part of the specification where appropriate. In all examples shown and discussed herein, any particular value should be construed as merely illustrative, and not limiting. Thus, other examples of the exemplary embodiments may have different values. It should be noted that: like reference numbers and letters refer to like items in the following figures, and thus, once an item is defined in one figure, further discussion thereof is not required in subsequent figures.
Spatially relative terms, such as "above … …", "above … …", "above … …", "above", and the like, may be used herein for ease of description to describe one device or feature's spatial relationship to another device or feature as illustrated in the figures. It will be understood that the spatially relative terms are intended to encompass different orientations of the device in use or operation in addition to the orientation depicted in the figures. For example, if a device in the figures is turned over, devices described as "above" or "on" other devices or configurations would then be oriented "below" or "under" the other devices or configurations. Thus, the exemplary term "above … …" can include both an orientation of "above … …" and "below … …". The device may be otherwise variously positioned and the spatially relative descriptors used herein interpreted accordingly.
Referring to fig. 1, an aircraft engine of some embodiments includes a forward bearing cavity 1, a middle bearing cavity 2, an aft bearing cavity 3, a compressor blade root 4, a compressor blade tip 5, a compressor hub 6, an air duct 7, and a high-pressure turbine 8.
The bearing cavity sealing system of the aircraft engine comprises a gas guiding flow path and at least two exhaust flow paths, wherein the gas guiding flow path is used for guiding sealing gas from a gas compressor to a middle bearing cavity and a rear bearing cavity. The exhaust backpressure of the at least two exhaust gas flow paths are different, and the bleed air flow path is selectively in communication with at least one of the at least two exhaust gas flow paths. According to the bearing cavity sealing system of the aircraft engine, provided by the embodiment of the invention, the plurality of exhaust flow paths with different exhaust back pressures are arranged, so that the exhaust back pressures of the bleed air flow paths can be controlled according to the running state of the aircraft engine, and the flow of the sealed gas can be further regulated.
Referring to fig. 1 to 3, a bearing cavity sealing system of an aircraft engine according to an embodiment of the present invention includes:
the air-entraining flow path 13 is used for guiding the sealed air from the air compressor to the middle bearing cavity 2 and the rear bearing cavity 3;
the first exhaust flow path 11 flows through the rear side of the rear journal of the low-pressure turbine and is communicated with the outlet 10 of the low-pressure turbine;
a second exhaust flow path 9 in communication with the atmosphere, and a bleed air flow path configured to communicate with at least one of the first exhaust flow path 11 and the second exhaust flow path 9.
The bearing cavity sealing system comprises a first exhaust flow path 11 communicated with a low-pressure turbine outlet 10 and a second exhaust flow path 9 communicated with the atmosphere, wherein the second exhaust flow path 9 is communicated with the atmosphere, and when an aircraft engine is in a high Mach number pneumatic parameter state, the bleed air flow path is controlled to be communicated with the first exhaust flow path 11 and the second exhaust flow path 9, so that exhaust back pressure is reduced, the flow of sealing gas is increased, and the risk of over-temperature of a compressor disc or a shaft caused by high-speed rotation is effectively reduced. And the reduction of the exhaust back pressure can further reduce the horizontal distribution of the on-way pressure of the bleed air flow path, thereby effectively reducing the sealing pressure difference of the bearing cavity.
In embodiments not shown in the other figures, the first and second exhaust gas flow paths may also communicate with other lower pressure locations of the aircraft engine (e.g., may also be axial locations) as long as the first and second exhaust gas flow paths are made to have different exhaust backpressure. Of course, in other embodiments, more than two exhaust gas flow paths with different exhaust gas counter pressures may be provided.
In some embodiments, the opening degree adjusting member 12 is provided on the second exhaust gas flow path 9, and the opening degree adjusting member 12 is configured to adjust the flow area of the second exhaust gas flow path 9. Specifically, in one embodiment, the opening degree adjuster 12 is an on-off control valve provided on the second exhaust flow path 9, and when the on-off control valve is opened, the second exhaust flow path 9 is in a communicating state, and at this time, the gas in the bleed air flow path can be communicated to the atmosphere through the second exhaust flow path 9; when the on-off control valve is closed, the second exhaust flow path 9 is in a disconnected state, in which the bleed air flow path is not able to effect exhaust through the second exhaust flow path 9. In another embodiment, the opening degree adjusting member 12 is an opening degree adjusting valve provided on the second exhaust gas flow path 9, and the opening degree adjusting valve can adjust the flow area of the second exhaust gas flow path 9 from zero to full opening, that is, the opening degree adjusting valve can not only control the opening and closing of the second exhaust gas flow path 9, but also control the opening degree of the second exhaust gas flow path when the second exhaust gas flow path 9 is in a communicating state, for example, to make the second exhaust gas flow path 9 in a 50% open state.
In some embodiments, referring to fig. 1-3, the bearing cavity sealing system includes an exhaust pipe having a first end in communication with the bleed air flow path and a second end extending to the atmosphere, an inner cavity of the exhaust pipe forming a second exhaust flow path 9, and an opening adjuster 12 disposed on the exhaust pipe.
In some embodiments, referring to fig. 1, after the leading air flow path is sealed off from the front bearing cavity 1, the leading air flow path flows into the middle bearing cavity 2 and the rear bearing cavity 3 through the compressor disk core 6 and the air conduit 7 in sequence. And the bleed air flow path leads out the seal air from the compressor blade root 4 or the compressor blade tip 5, as shown by the arrow a in fig. 1, and the seal air flows into the middle bearing cavity 2 and the rear bearing cavity 3 along the compressor disk core 6 and the air duct 7 in sequence.
The invention also provides a control method of the bearing cavity sealing system based on the aero-engine, which comprises the following steps:
in a first state, the bleed air flow path is controlled to be communicated with the first exhaust flow path 11 and the second exhaust flow path 9; and
in a second state the bleed air flow path is controlled to be in communication with the first exhaust air flow path 11 and to be disconnected from the second exhaust air flow path 9.
In the embodiment, the first exhaust flow path 11 is communicated with the low-pressure turbine outlet 10, the second exhaust flow path 9 is communicated with the atmosphere, and because the atmospheric environmental pressure is far lower than the pressure of the low-pressure turbine outlet 10, in a first state, the second exhaust flow path 9 is opened, so that the first exhaust flow path 11 and the second exhaust flow path 9 work simultaneously, and because the second exhaust flow path 9 is opened, the exhaust back pressure is reduced, the airflow flow in the bearing sealing flow path is increased, and therefore the influence of wind resistance temperature rise caused by the rotation of a compressor disc or a shaft is effectively reduced; meanwhile, due to the opening of the second exhaust flow path 9, the exhaust back pressure is further reduced, the horizontal distribution of the on-way pressure in the flow path is further reduced, and the sealing pressure difference of the bearing cavity is effectively reduced.
In the second state, because the engine pressure and temperature parameter level are lower, therefore can close second exhaust flow path 9, only first exhaust flow path 11 works, and the risk that the bearing obturation air current is overtemperature is lower, promotes along-the-way pressure level distribution simultaneously, increases the differential pressure of obturation.
In some embodiments, the opening of the second exhaust flow path is controlled in the first state to adjust the airflow rate of the bleed air flow path.
Fig. 1 shows a schematic structural diagram of a bearing cavity sealing system of an aircraft engine according to an embodiment of the invention. The sealed gas is guided into a middle bearing cavity and a rear bearing cavity from a certain stage blade tip or blade root of the compressor, in the process, the sealed gas is guided from a certain stage blade tip or blade root 4 of the compressor, flows through parts such as a compressor disk core 6, an air conduit 7 and the like after being sealed for a front bearing cavity 1, then sequentially flows to the middle bearing cavity 2 and the sealing cavity of the rear bearing cavity 3 to be sealed and insulated, and then is converged with an outer cavity low-pressure cavity of the front sealing cavity of the middle bearing cavity 2 after being coiled by a high-pressure turbine 8. The exhaust gas flow path includes a first exhaust gas flow path 11 and a second exhaust gas flow path 9, the first exhaust gas flow path 11 communicating with an outlet 10 of the low pressure turbine; the second exhaust gas flow path 9 communicates with the atmosphere.
Finally, it should be noted that: the above examples are only intended to illustrate the technical solution of the present invention and not to limit it; although the present invention has been described in detail with reference to preferred embodiments, those skilled in the art will understand that: modifications to the specific embodiments of the invention or equivalent substitutions for parts of the technical features may be made; without departing from the spirit of the invention, it is intended to cover all modifications within the scope of the invention as claimed.
Claims (10)
1. The utility model provides an aeroengine's bearing bore system of obturating which characterized in that includes:
the air guide flow path is used for guiding the sealed air from the air compressor to the middle bearing cavity and the rear bearing cavity; and
at least two exhaust air flow paths having different exhaust backpressure, the bleed air flow path being selectively in communication with at least one of the at least two exhaust air flow paths.
2. The aircraft engine bearing cavity sealing system of claim 1 wherein said at least two exhaust gas flow paths comprise:
a first exhaust flow path communicating with the low pressure turbine outlet; and
a second exhaust flow path in communication with the atmosphere, and the bleed air flow path is configured to communicate with at least one of the first and second exhaust flow paths.
3. The aeroengine bearing cavity sealing system of claim 2, wherein an opening adjuster is disposed on the second exhaust gas flow path, the opening adjuster configured to adjust a flow area of the second exhaust gas flow path.
4. The aircraft engine bearing cavity sealing system of claim 3, wherein the bearing cavity sealing system comprises an exhaust pipe, a first end of the exhaust pipe is communicated with the bleed air flow path, a second end of the exhaust pipe extends to the atmosphere, an inner cavity of the exhaust pipe forms the second exhaust flow path, and the opening degree adjusting member is arranged on the exhaust pipe.
5. The aeroengine bearing cavity sealing system of claim 3, wherein the opening adjuster is an on-off control valve.
6. The aeroengine bearing cavity sealing system of claim 3, wherein the opening adjuster is an opening adjuster valve.
7. The aircraft engine bearing cavity sealing system of claim 1 wherein said bleed air flowpath passes sequentially through a compressor hub and an air conduit into said middle and aft bearing cavities.
8. A control method of a bearing cavity sealing system of an aircraft engine based on any one of claims 1 to 7 is characterized by comprising the following steps: and controlling the bleed air flow path to be communicated with at least one of the at least two exhaust gas flow paths according to the running state of the aircraft engine.
9. The method of controlling a bearing cavity sealing system of an aircraft engine of claim 8, wherein the at least two exhaust flow paths include a first exhaust flow path in communication with the low pressure turbine outlet and a second exhaust flow path in communication with the atmosphere, and controlling the bleed air flow path to communicate with at least one of the at least two exhaust flow paths based on an operating condition of the aircraft engine comprises:
controlling the bleed air flow path to communicate with both the first exhaust flow path and the second exhaust flow path in a first state; and
and in a second state, the bleed air flow path is controlled to be communicated with the first exhaust flow path and the bleed air flow path is controlled to be disconnected from the second exhaust flow path.
10. The method of claim 9, wherein the opening of the second exhaust flow path is controlled in a first state to regulate the airflow rate of the bleed air flow path.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110089204.7A CN114776402A (en) | 2021-01-22 | 2021-01-22 | Bearing cavity sealing system of aircraft engine and control method thereof |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110089204.7A CN114776402A (en) | 2021-01-22 | 2021-01-22 | Bearing cavity sealing system of aircraft engine and control method thereof |
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| CN114776402A true CN114776402A (en) | 2022-07-22 |
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| CN202110089204.7A Pending CN114776402A (en) | 2021-01-22 | 2021-01-22 | Bearing cavity sealing system of aircraft engine and control method thereof |
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| CN103282606A (en) * | 2011-03-31 | 2013-09-04 | 三菱重工业株式会社 | Low pressure steam turbine |
| CN103939626A (en) * | 2013-01-23 | 2014-07-23 | 波音公司 | Dual door fan air modulating valve |
| CN107636258A (en) * | 2015-05-07 | 2018-01-26 | 劳斯莱斯有限公司 | Gas-turbine unit |
| CN107636295A (en) * | 2015-02-20 | 2018-01-26 | 普拉特 - 惠特尼加拿大公司 | Engine charge component with selector valve |
| CN109139122A (en) * | 2018-11-07 | 2019-01-04 | 哈尔滨电气股份有限公司 | A kind of inner cooling system of 2 grades of turbine rotors of gas turbine |
-
2021
- 2021-01-22 CN CN202110089204.7A patent/CN114776402A/en active Pending
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103282606A (en) * | 2011-03-31 | 2013-09-04 | 三菱重工业株式会社 | Low pressure steam turbine |
| CN103939626A (en) * | 2013-01-23 | 2014-07-23 | 波音公司 | Dual door fan air modulating valve |
| CN107636295A (en) * | 2015-02-20 | 2018-01-26 | 普拉特 - 惠特尼加拿大公司 | Engine charge component with selector valve |
| CN107636258A (en) * | 2015-05-07 | 2018-01-26 | 劳斯莱斯有限公司 | Gas-turbine unit |
| CN109139122A (en) * | 2018-11-07 | 2019-01-04 | 哈尔滨电气股份有限公司 | A kind of inner cooling system of 2 grades of turbine rotors of gas turbine |
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