CN116291874A - Method and system for extracting front-end power of large bypass ratio aero-engine - Google Patents
Method and system for extracting front-end power of large bypass ratio aero-engine Download PDFInfo
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- CN116291874A CN116291874A CN202310162597.9A CN202310162597A CN116291874A CN 116291874 A CN116291874 A CN 116291874A CN 202310162597 A CN202310162597 A CN 202310162597A CN 116291874 A CN116291874 A CN 116291874A
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- 238000000034 method Methods 0.000 title abstract description 14
- 238000000605 extraction Methods 0.000 claims abstract description 84
- 238000013461 design Methods 0.000 claims abstract description 23
- 238000001816 cooling Methods 0.000 claims description 27
- 230000005540 biological transmission Effects 0.000 claims description 26
- 238000010438 heat treatment Methods 0.000 claims description 11
- 239000012528 membrane Substances 0.000 claims description 4
- 239000002184 metal Substances 0.000 description 6
- 230000008878 coupling Effects 0.000 description 3
- 238000010168 coupling process Methods 0.000 description 3
- 238000005859 coupling reaction Methods 0.000 description 3
- 238000011161 development Methods 0.000 description 3
- 238000005516 engineering process Methods 0.000 description 3
- 230000005284 excitation Effects 0.000 description 3
- 238000009434 installation Methods 0.000 description 3
- 238000006243 chemical reaction Methods 0.000 description 2
- 230000005611 electricity Effects 0.000 description 2
- 230000008569 process Effects 0.000 description 2
- 238000006467 substitution reaction Methods 0.000 description 2
- 238000012546 transfer Methods 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 1
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C6/00—Plural gas-turbine plants; Combinations of gas-turbine plants with other apparatus; Adaptations of gas-turbine plants for special use
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/12—Cooling of plants
- F02C7/16—Cooling of plants characterised by cooling medium
- F02C7/18—Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
- F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
- F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
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- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Connection Of Motors, Electrical Generators, Mechanical Devices, And The Like (AREA)
Abstract
The invention provides a method and a system for extracting front power of an aeroengine with a large bypass ratio, wherein the system comprises a low-voltage rotor front power extraction subsystem, the low-voltage rotor front power extraction subsystem comprises a supporting structure, the supporting structure comprises an outer casing and an inner casing which are coaxial, the outer casing is connected with a fan casing, the front end of the inner casing is connected with the outer casing through a bearing support plate, the rear end of the inner casing is in lap joint with a low-voltage rotor disc, and an inner cavity at the front end of the inner casing is provided with a generator. The method comprises a front power extraction mode design step and a front power extraction system design step; the front power extraction mode design step comprises the following steps: high-pressure rotor front-end power and low-pressure rotor front-end power are extracted and distributed. The method and the system designed by the invention extract the low-voltage rotor power in a preposed power extraction mode, ensure the power consumption requirement of the aeroengine under the condition of ensuring the margin of the air compressor, make up the blank of the industry and have important guiding significance.
Description
Technical Field
The invention belongs to the field of aeroengines, relates to an engine power extraction design technology, and particularly relates to a method and a system for extracting front-end power of an aeroengine with a large bypass ratio.
Background
With the development of communication technology, computer technology and on-board laser weapons, aircraft demand for electrical power has increased, such that aircraft engines are facing increasingly greater power extraction (i.e., the generation of electrical energy by engine-driven generators, the portion of the energy consumed by the engine being referred to as power extraction) demands.
For example, when an aircraft performs a flight mission at high altitude, when work generated by a turbine is reduced, the power extraction capability of an engine is reduced, and if high power extraction is continued, the surge margin of the engine is insufficient to influence the flight safety, and if the flight safety is ensured, the high power extraction requirement of the aircraft cannot be met. For another example, modern aero-engines mainly adopt a double-rotor structure, engine power extraction mainly comes from a high-pressure rotor, wherein when a medium and small thrust engine works in a high-altitude state, the margin of a compressor is obviously reduced when enough power is extracted from the high-pressure rotor due to small power emitted by a turbine.
Meanwhile, a front-mounted low-voltage power extraction method for the aero-engine with the large bypass ratio is not designed at present.
Disclosure of Invention
The invention aims to disclose a method and a system for extracting front-mounted power of an aeroengine with a large bypass ratio, which can meet the power requirement of an aircraft on the premise that the flight safety cannot be ensured, and can also solve the problems of installation and working reliability of a power extraction system.
The technical scheme for realizing the aim of the invention is as follows:
in a first aspect, the present invention provides a high bypass ratio aircraft engine forward power extraction system comprising a low pressure rotor forward power extraction subsystem comprising a support structure, a forward power transfer device.
The support structure comprises an outer casing and an inner casing which are coaxial, the outer casing is connected with the fan casing, the front end of the inner casing is connected with the outer casing through a bearing support plate, the rear end of the inner casing is in lap joint with the low-pressure rotor disc, and a generator is assembled in an inner cavity of the front end of the inner casing. Because the low-voltage rotor has lower rotating speed and low extractable power, the generator is heavier and larger in size, and therefore, in order to ensure the reliable installation and force transmission of the generator, the generator is installed at the bearing support plate, and the axial distance between the bearing support plate and the fan rotor blade is longer, so that the length of the power transmission device between the motor and the low-voltage rotor is longer.
Further, the axial distance between the bearing support plate and the fan rotor blade is 1-2 times the chord length of the fan rotor blade so as to ensure the aerodynamic stability of the fan rotor.
Further, the front power transmission device is located in the inner cavity of the rear end of the inner casing, one end of the front power transmission device is connected with the generator, and the other end of the front power transmission device is connected with the low-voltage rotor blade disc.
Still further, leading power transmission device is located between low pressure rotor and the generator, solves the installation and the transmission stability that transmission device length is longer arouses simultaneously, designs leading power transmission system, and leading power transmission system includes the drive shaft that connects gradually, the drive shaft other end sets up on low pressure rotor leaf dish, the other end with power output shaft connection, just the other end of power output shaft with the generator is connected.
Preferably, a membrane disc coupler is arranged between the driving shaft and the power output shaft.
Further, the low-pressure rotor preposed power extraction subsystem further comprises a cooling structure, the cooling structure comprises a cooling exhaust pipe and a cooling air supply pipe which are positioned in the inner cavity of the bearing support plate, and through holes through which the cooling exhaust pipe and the cooling air supply pipe penetrate are formed in the outer casing and the inner casing.
Further, the low pressure rotor front power extraction subsystem further comprises an air intake cap disposed at the front end of the support structure;
the air inlet cap cover and/or the bearing support plate is/are provided with a heating element, and the heating element is any one of a metal resistance wire, a metal resistance sheet and a metal resistance film.
In a second aspect, the invention provides a method for extracting front-end power of an aeroengine with a large bypass ratio, which comprises a front-end power extraction mode design step and a front-end power extraction system design step;
the front power extraction mode design step comprises the following steps: high-pressure rotor power extraction and low-pressure rotor power extraction and distribution;
the front-end power extraction system design includes: any one or more of a power output structure, a supporting structure and a cooling structure in the low-pressure rotor front-end power extraction subsystem are designed.
Further, the high pressure rotor power extraction and low pressure rotor power extraction distribution includes: and according to the engine performance, the high-pressure rotor power limit value and the matching of the high-pressure rotor and the low-pressure rotor performance, acquiring a low-pressure rotor power extraction range under the condition of meeting the engine component efficiency and the instability boundary margin of the engine rotor component.
Compared with the prior art, the invention has the beneficial effects that: according to the engine front-end power extraction system and method, partial front-end power is extracted through the low-voltage rotor in an auxiliary mode, and electricity consumption requirements of an aeroengine are guaranteed under the condition that the margin of a gas compressor is reduced. Meanwhile, the low-voltage rotor front-end power extraction subsystem is designed in terms of system engineering, and particularly, a supporting structure, a cooling structure, a front-end power transmission structure and the like of the low-voltage rotor front-end power extraction subsystem are designed, so that the front-end power extraction system designed by the invention has the following advantages:
1. the technical problems that the rotation speed of a low-voltage rotor of the large bypass ratio aero-engine is low, the motor scheme is large and the internal arrangement cannot be carried out are solved through the design of the supporting structure, and the problems of pneumatic excitation, anti-icing and the like caused by the conventional supporting structure are solved;
2. the front power transmission structure adopts a three-section design, so that the rotor dynamics problem caused by the electromechanical conversion of the aero-engine is solved;
the system and the method designed by the invention are applied in the development process of a certain turbofan engine, are high-reliability large-bypass-ratio aeroengine power extraction technology, make up the blank of the industry and have important guiding significance.
Drawings
In order to more clearly illustrate the technical solution of the embodiments of the present invention, the drawings that are needed in the description of the embodiments will be briefly described.
FIG. 1 is a schematic diagram of a low pressure rotor front-end power extraction subsystem of a high bypass ratio aircraft engine front-end power extraction system in an embodiment;
1, a fan casing; 2. an outer casing; 3. a fan rotor blade; 4. an inner casing; 5. a drive shaft; 6. film disc coupling; 7. a power output shaft; 8. a bearing support plate; 9. cooling the exhaust pipe; 10. a heating member; 11. a generator; 12. an air intake cap; 13. a heating cable; 14. cooling the gas supply tube; 15. a power transmission cable; 16. and a motor control cable.
Detailed Description
The invention will be further described with reference to specific embodiments, and advantages and features of the invention will become apparent from the description. These examples are merely exemplary and do not limit the scope of the invention in any way. It will be understood by those skilled in the art that various changes and substitutions of details and forms of the technical solution of the present invention may be made without departing from the spirit and scope of the present invention, but these changes and substitutions fall within the scope of the present invention.
Example 1:
the embodiment provides a front-end power extraction method of a large bypass ratio aeroengine, which comprises a front-end power extraction mode design step and a front-end power extraction system design step;
the front power extraction mode design step comprises the following steps: high pressure rotor power extraction and low pressure rotor power extraction distribution.
In this step, the high-pressure rotor power extraction and low-pressure rotor power extraction and distribution includes: and according to the engine performance, the high-pressure rotor power limit value and the matching of the high-pressure rotor and the low-pressure rotor performance, acquiring a low-pressure rotor power extraction range under the condition of meeting the engine component efficiency and the instability boundary margin of the engine rotor component.
Wherein the front-end power extraction system design comprises: any one or more of a power output structure, a supporting structure and a cooling structure in the low-pressure rotor front-end power extraction subsystem are designed.
Alternatively, because the large bypass is lower than the rotation speed of the aero-engine and the extractable power is low, a permanent magnet type or three-stage motor with larger volume is selected to extract power in the embodiment.
Optionally, the air cooling mode is selected to meet the cooling requirement of the motor by comprehensively considering factors such as the heat productivity of the motor, structural design, external system layout and the like and the condition that the pressure of the position where the engine can bleed air is high.
Example 2:
the embodiment provides a front-end power extraction system of an aeroengine with a large bypass ratio, wherein the front-end power extraction system of the aeroengine comprises a front-end power extraction subsystem of a high-voltage rotor, and the front-end power extraction subsystem of the high-voltage rotor adopts an existing design mode and is not described in the embodiment.
The engine front-end power extraction system also includes a low-pressure rotor front-end power extraction subsystem including a support structure, a front-end power transfer device.
In this embodiment, the generator 11 is installed by adopting a front support and rear fixing manner, specifically, as shown in fig. 1, the supporting structure includes an outer casing 2 and an inner casing 4 that are coaxial, the outer casing 2 is connected with the fan casing 1, and the front end of the inner casing 4 is connected with the outer casing 2 through a bearing support plate 8. The rear end of the inner casing 4 is in lap joint with the low-voltage rotor disc, and the inner cavity of the front end of the inner casing 4 is provided with the generator 11. The load bearing support plates 8 are circumferentially distributed and can reduce the vibration response of the generator 11 when the engine is in operation. An outer flow passage air inlet wall surface is formed between the outer casing 2 and the inner casing 4, and the outer casing 2 is connected with the fan casing 1 through a bolt, so that load generated by the whole supporting structure can be transmitted to an engine force transmission system.
In this embodiment, the bearing support plate 8 is a hollow structure, and is used for lead arrangement of cables such as a power transmission cable 15 and a motor control cable 16.
Further, when the engine works, exciting force generated by air flow separation can affect the vibration of the fan rotor, and the pneumatic analysis result shows that on the premise of reasonably designing the shape of the bearing support plate 8, the arrangement of the axial distance between the bearing support plate 8 and the fan rotor blade 3 can effectively reduce the pneumatic excitation influence, so that in the embodiment, the axial distance between the bearing support plate 8 and the fan rotor blade 3 is 1-2 times the chord length of the fan rotor blade 3, and the axial distance between the bearing support plate 8 and the fan rotor blade 3 is optimal when 1.5-2 times.
Further, referring to fig. 1, the front power transmission device is located in the inner cavity of the rear end of the inner casing 4, one end of the front power transmission device is connected with the generator 11, and the other end is connected with the low-pressure rotor blade disc.
Still further, referring to fig. 1, the front power transmission device is located between the low-voltage rotor and the generator 11, and includes a driving shaft 5 sequentially connected, the other end of the driving shaft 5 is disposed on the low-voltage rotor disc, the other end of the driving shaft is connected to the power output shaft 7, and the other end of the power output shaft 7 is connected to the generator 11.
Preferably, as shown in fig. 1, a membrane disc coupling 6 is disposed between the driving shaft 5 and the power output shaft 7, and the membrane disc coupling 6 can generate flexible deformation during extraction of the front power, so as to counteract the problems of eccentricity of the low-pressure rotor and unstable operation of the front power transmission device.
Further, the low-pressure rotor front-mounted power extraction subsystem further comprises a cooling structure, as shown in fig. 1, the cooling structure comprises a cooling exhaust pipe 9 and a cooling air supply pipe 14 which are positioned in the inner cavity of the bearing support plate 8, and through holes through which the cooling exhaust pipe 9 and the cooling air supply pipe 14 penetrate are formed in the outer casing 2 and the inner casing 4.
Further, referring to fig. 1, the low pressure rotor front power extraction subsystem further includes an intake cap 12, the intake cap 12 being disposed at the forward end of the support structure.
Still further, referring to fig. 1, the air intake cap 12 and/or the bearing support plate 8 are provided with a heating element 10, and when the aeroengine needs to perform anti-icing, the heating element 10 can transmit a control signal through the engine, and the generator is powered to perform heating anti-icing. In this embodiment, the heating element 10 may be any one of a metal resistance wire, a metal resistance sheet, and a metal resistance film, and the heating element 10 is connected to the generator 11 via the heating cable 13.
According to the engine front-end power extraction system and method, partial front-end power is extracted through the low-voltage rotor in an auxiliary mode, and electricity consumption requirements of an aeroengine are guaranteed under the condition that the margin of a gas compressor is reduced. Meanwhile, the low-voltage rotor front-end power extraction subsystem is designed in terms of system engineering, and particularly, a supporting structure, a cooling structure, a front-end power transmission structure and the like of the low-voltage rotor front-end power extraction subsystem are designed, so that the front-end power extraction system designed by the invention has the following advantages:
1. the technical problems that the rotation speed of a low-voltage rotor of the large bypass ratio aero-engine is low, the motor scheme is large and the internal arrangement cannot be carried out are solved through the design of the supporting structure, and the problems of pneumatic excitation, anti-icing and the like caused by the conventional supporting structure are solved;
2. the front power transmission structure adopts a three-section design, so that the rotor dynamics problem caused by the electromechanical conversion of the aero-engine is solved;
the system and the method designed by the invention are applied in the development process of a certain turbofan engine, are high-reliability large-bypass-ratio aeroengine power extraction technology, make up the blank of the industry and have important guiding significance.
The foregoing description of the preferred embodiments of the invention is not intended to be limiting, but rather is intended to cover all modifications, equivalents, alternatives, and improvements that fall within the spirit and scope of the invention.
Furthermore, it should be understood that although the present disclosure describes embodiments, not every embodiment is provided with a separate embodiment, and that this description is provided for clarity only, and that the disclosure is not limited to the embodiments described in detail below, and that the embodiments described in the examples may be combined as appropriate to form other embodiments that will be apparent to those skilled in the art.
Claims (9)
1. The high bypass ratio aeroengine front-end power extraction system is characterized by comprising a low-pressure rotor front-end power extraction subsystem, wherein the low-pressure rotor front-end power extraction subsystem comprises a support structure and a front-end power transmission device;
the supporting structure comprises an outer casing (2) and an inner casing (4) which are coaxial, the outer casing (2) is connected with the fan casing (1), the front end of the inner casing (4) is connected with the outer casing (2) through a bearing support plate (8), the rear end of the inner casing (4) is lapped with the low-pressure rotor disc, and an inner cavity of the front end of the inner casing (4) is provided with a generator (11).
2. The high bypass ratio aircraft engine forward power extraction system of claim 1 wherein: the axial distance between the bearing support plate (8) and the fan rotor blade (3) is 1-2 times the chord length of the fan rotor blade (3).
3. The high bypass ratio aircraft engine forward power extraction system of claim 1 wherein: the front power transmission device is located in the inner cavity of the rear end of the inner casing (4), one end of the front power transmission device is connected with the generator (11), and the other end of the front power transmission device is connected with the low-voltage rotor blade disc.
4. The high bypass ratio aircraft engine forward power extraction system of claim 3 wherein: the preposed power transmission device is located between the low-voltage rotor and the generator and comprises a driving shaft (5) which is sequentially connected, the other end of the driving shaft (5) is arranged on a low-voltage rotor disc, the other end of the driving shaft is connected with a power output shaft (7), and the other end of the power output shaft (7) is connected with the generator.
5. The high bypass ratio aircraft engine forward power extraction system of claim 4 wherein: a membrane disc coupler (6) is arranged between the driving shaft (5) and the power output shaft (7).
6. The high bypass ratio aircraft engine forward power extraction system of any one of claims 1 to 5 wherein: the low-pressure rotor preposed power extraction subsystem further comprises a cooling structure, the cooling structure comprises a cooling exhaust pipe (9) and a cooling air supply pipe (14) which are positioned in the inner cavity of the bearing support plate (8), and through holes through which the cooling exhaust pipe (9) and the cooling air supply pipe (14) penetrate are formed in the outer casing (2) and the inner casing (4).
7. The high bypass ratio aircraft engine forward power extraction system of claim 6 wherein: the low pressure rotor front power extraction subsystem further comprises an air inlet cap (12), wherein the air inlet cap (12) is arranged at the front end of the supporting structure;
and a heating element (10) is arranged on the air inlet cap cover (12) and/or the bearing support plate (8).
8. The front-end power extraction method of the large bypass ratio aero-engine is characterized by comprising a front-end power extraction mode design step and a front-end power extraction system design step;
the front power extraction mode design step comprises the following steps: high-pressure rotor power extraction and low-pressure rotor power extraction and distribution;
the front-end power extraction system design includes: any one or more of a power output structure, a supporting structure and a cooling structure in the low-pressure rotor front-end power extraction subsystem are designed.
9. The high bypass ratio aircraft engine forward power extraction method of claim 8, wherein: the high pressure rotor power extraction and low pressure rotor power extraction distribution comprises: and according to the engine performance, the high-pressure rotor power limit value and the matching of the high-pressure rotor and the low-pressure rotor performance, acquiring a low-pressure rotor power extraction range under the condition of meeting the engine component efficiency and the instability boundary margin of the engine rotor component.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310162597.9A CN116291874A (en) | 2023-02-24 | 2023-02-24 | Method and system for extracting front-end power of large bypass ratio aero-engine |
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| Application Number | Priority Date | Filing Date | Title |
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| CN202310162597.9A CN116291874A (en) | 2023-02-24 | 2023-02-24 | Method and system for extracting front-end power of large bypass ratio aero-engine |
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| CN116291874A true CN116291874A (en) | 2023-06-23 |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN118150172A (en) * | 2024-05-13 | 2024-06-07 | 中国航发四川燃气涡轮研究院 | Automatic linkage test run method for aeroengine power extraction system |
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2023
- 2023-02-24 CN CN202310162597.9A patent/CN116291874A/en active Pending
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN118150172A (en) * | 2024-05-13 | 2024-06-07 | 中国航发四川燃气涡轮研究院 | Automatic linkage test run method for aeroengine power extraction system |
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