CN114017388A - Device for improving working Mach number of aviation turbine engine - Google Patents
Device for improving working Mach number of aviation turbine engine Download PDFInfo
- Publication number
- CN114017388A CN114017388A CN202111401287.5A CN202111401287A CN114017388A CN 114017388 A CN114017388 A CN 114017388A CN 202111401287 A CN202111401287 A CN 202111401287A CN 114017388 A CN114017388 A CN 114017388A
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- turbine engine
- engine
- mach number
- bleed
- aircraft turbine
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- 238000002485 combustion reaction Methods 0.000 claims abstract description 7
- 230000007246 mechanism Effects 0.000 claims description 3
- 230000003028 elevating effect Effects 0.000 claims 1
- 238000002347 injection Methods 0.000 abstract description 6
- 239000007924 injection Substances 0.000 abstract description 6
- 239000000446 fuel Substances 0.000 abstract description 4
- 238000001816 cooling Methods 0.000 abstract description 3
- 239000002826 coolant Substances 0.000 description 4
- 230000006872 improvement Effects 0.000 description 3
- 238000000034 method Methods 0.000 description 3
- 230000000694 effects Effects 0.000 description 2
- 238000012986 modification Methods 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 239000000243 solution Substances 0.000 description 2
- 230000009471 action Effects 0.000 description 1
- 230000002411 adverse Effects 0.000 description 1
- 238000006243 chemical reaction Methods 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 238000009434 installation Methods 0.000 description 1
- 230000009467 reduction Effects 0.000 description 1
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/40—Casings; Connections of working fluid
- F04D29/403—Casings; Connections of working fluid especially adapted for elastic fluid pumps
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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 application belongs to the technical field of engine design, and particularly relates to a device for improving the working Mach number of an aviation turbine engine. The device comprises an air entraining pipe (2), wherein the air entraining pipe (2) is arranged on the outer side of an engine host, one end of the air entraining pipe (2) is connected with a compressor middle stage of the aircraft turbine engine, the other end of the air entraining pipe (2) is connected with a combustion chamber of the aircraft turbine engine, and an air entraining valve (3) is arranged at the position where the air entraining pipe (2) is connected with the compressor middle stage. The turbine engine provided by the application does not need to bear heavy cooling media, containers and injection devices, effectively reduces the takeoff weight of the airplane, and can improve the fuel economy and increase the voyage of the airplane.
Description
Technical Field
The application belongs to the technical field of engine design, and particularly relates to a device for improving the working Mach number of an aviation turbine engine.
Background
In recent years, high-speed aircrafts have become the key research direction in the field of aviation, and the power technology of the high-speed aircrafts is one of the difficulties in developing the high-speed aircrafts. At present, the limit of the working Mach number of the domestic turbine engine is about Ma2, if the flight Mach number is continuously increased, the inlet temperature of the engine is greatly increased, the converted rotating speed of the engine is sharply reduced, the flow of the engine is greatly reduced, and the thrust of the engine cannot meet the requirement of an airplane. If the working envelope of the engine is further expanded, the technical improvement on the existing turbine engine scheme is needed.
The prior technical scheme mainly adopts a jet flow precooling technology, a cooling medium injection device is additionally arranged in an air inlet channel of an airplane, and the cooling medium is injected to cool incoming flow in a high-Mach number state of an engine, so that the inlet temperature of the engine is reduced, and the conversion flow of the engine is increased, thereby increasing the thrust under the high-Mach number of the engine and realizing high-speed flight.
The jet flow precooling technology is adopted to realize the work Mach number promotion of the engine, but the scheme also has certain defects:
the cooling medium injection device needs to be additionally arranged in the aircraft air inlet channel, and the aircraft needs to take off with a large amount of cooling medium and a container thereof, so that the take-off weight of the aircraft is greatly increased, the aircraft needs to carry a large amount of dead weight in flight, and adverse effects are brought to the performance exertion of the aircraft and the fuel economy.
Meanwhile, the installation of the injection device in the air inlet channel can cause various problems such as increase of total pressure loss of the air inlet channel, temperature/pressure distortion of air flow at the inlet of the engine and the like, so that the thrust of the engine is reduced, and certain influence is generated on the working stability of the engine.
Disclosure of Invention
In view of the limitation of a jet flow precooling technology, the method for air entraining of the intermediate-stage bypass of the air compressor is provided for the problem of insufficient engine thrust in a high-Mach number state, partial flow is enabled to cross the subsequent stages of the air compressor and the turbine in the high-Mach number state, the flow capacity of a main engine of the engine is improved, and the flow of the engine is increased. Meanwhile, the led-out flow directly flows into an afterburner to participate in combustion, and the thrust of the engine is improved. Under the combined action of the two aspects of lifting effects, the thrust performance under the high Mach number of the engine is comprehensively improved, and the working envelope of the engine is expanded.
The application provides a promote device of aviation turbine engine work mach number, including bleed pipe, bleed pipe sets up in the engine host computer outside, and the compressor intermediate level of aviation turbine engine is connected to its one end, and the other end is connected to aviation turbine engine's combustion chamber, bleed pipe is connecting the position department of compressor intermediate level is provided with the bleed valve.
Preferably, the bleed air duct comprises a plurality of bleed air ducts which are arranged circumferentially and uniformly around the circumference of the aircraft turbine engine.
Preferably, the bleed air pipe is provided with a lateral opening at a position connecting the compressor middle stage, and the bleed valve can extend into or out of the bleed air pipe from the lateral opening.
Preferably, the bleed valve includes a link ring which is driven by the drive mechanism to move axially along the engine, the link ring having a plurality of baffles extending axially along the engine, each baffle being inserted into or out of the bleed duct from a lateral opening of the bleed duct.
Preferably, the bleed air duct is supported outside the aircraft turbine engine by a bracket.
Preferably, a swirling device is arranged in the bleed air pipe, and the swirling device is a plurality of staggered inclined blades.
Compared with a jet flow precooling scheme, the method has the advantages that:
1) the airplane does not need to bear heavy cooling media, containers and injection devices, and the takeoff weight is effectively reduced;
2) the reduction of the dead weight in the air can improve the maneuvering performance of the airplane, improve the fuel economy and increase the range of the airplane;
3) the method has the advantages that the flow field at the inlet of the engine is not influenced, the extra total pressure loss and temperature/pressure distortion at the inlet of the engine are not caused, and the aerial work stability of the engine is not deteriorated.
Drawings
Fig. 1 is a schematic view of the closing of bleed valves of the device for raising the operating mach number of an aircraft turbine engine according to the present invention.
Figure 2 is a schematic view of the embodiment of the invention shown in figure 1 with the bleed valves open.
The system comprises an engine host, an air guide pipe, an air guide valve, a support and a rotational flow device, wherein the engine host is 1, the air guide pipe is 2, the air guide valve is 3, the support is 4, and the rotational flow device is 5.
Detailed Description
In order to make the implementation objects, technical solutions and advantages of the present application clearer, the technical solutions in the embodiments of the present application will be described in more detail below with reference to the accompanying drawings in the embodiments of the present application. In the drawings, the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The described embodiments are some, but not all embodiments of the present application. The embodiments described below with reference to the drawings are exemplary and intended to be used for explaining the present application, and should not be construed as limiting the present application. All other embodiments obtained by a person of ordinary skill in the art without any inventive work based on the embodiments in the present application are within the scope of protection of the present application. Embodiments of the present application will be described in detail below with reference to the drawings.
The application provides a device for improving work Mach number of an aircraft turbine engine, as shown in figures 1 and 2, the device comprises an air guide pipe 2, the air guide pipe 2 is arranged on the outer side of an engine host, one end of the air guide pipe is connected with a compressor middle stage of the aircraft turbine engine, the other end of the air guide pipe is connected with a combustion chamber of the aircraft turbine engine, and the air guide pipe 2 is connected with an air guide valve 3 arranged at the position of the compressor middle stage.
At low mach numbers, the engine operating modes are shown in fig. 1: the engine main unit 1 has high working efficiency, and bypass air bleed is not needed at the moment, so that the air bleed valve 3 is closed, and no flow flows through the air bleed pipe 2;
at high mach numbers, the engine operating modes are shown in fig. 2: the working efficiency of the engine host 1 is reduced, bypass air entraining is needed at this time, the opening degree of the air entraining valve 3 is adjusted, controllable adjustment of air entraining flow under different Mach numbers is realized, the guided flow is discharged into the afterburner through the air entraining pipe 2, and the working capacity of the engine under high Mach numbers is improved.
In some alternative embodiments, the bleed air ducts comprise a plurality of bleed air ducts, which are arranged circumferentially and uniformly around the circumference of the aircraft turbine engine, for example 8 bleed air ducts arranged uniformly outside the engine in the axial direction of the engine.
In some alternative embodiments, the bleed air duct 2 is provided with a lateral opening at the location of the connection to the compressor intermediate stage, from which the bleed valve 3 can project into or out of the bleed air duct.
In some alternative embodiments, the bleed valve 3 comprises a linkage ring that is driven by a drive mechanism to move axially along the engine, and the linkage ring has a plurality of baffles extending axially along the engine, each baffle being inserted into or out of the bleed duct 2 from a lateral opening thereof.
In this embodiment, the valve openings of a plurality of bleed air ducts 2 can be controlled simultaneously by means of a linkage ring.
In some alternative embodiments, the bleed air duct 2 is supported outside the aircraft turbine engine by brackets 4. The structure is strengthened through the support 4, and the stability of the engine main body is prevented from being damaged by vibration of the air guide pipe 2.
In some optional embodiments, a swirling device 5 is arranged in the air-entraining pipe 2, and the swirling device is a plurality of staggered inclined blades, so that the air flow entering the combustion chamber is more uniform through the swirling device, and the combustion efficiency is improved.
The turbine engine provided by the application does not need to bear heavy cooling media, containers and injection devices, effectively reduces the takeoff weight of the airplane, and can improve the fuel economy and increase the voyage of the airplane.
Although the present application has been described in detail with respect to the general description and specific embodiments, it will be apparent to those skilled in the art that certain modifications or improvements may be made based on the present application. Accordingly, such modifications and improvements are intended to be within the scope of this invention as claimed.
Claims (6)
1. The device for improving the working Mach number of the aviation turbine engine is characterized by comprising an air guide pipe (2), wherein the air guide pipe (2) is arranged on the outer side of a main engine, one end of the air guide pipe is connected with a compressor middle stage of the aviation turbine engine, the other end of the air guide pipe is connected with a combustion chamber of the aviation turbine engine, and an air guide valve (3) is arranged at the position, connected with the compressor middle stage, of the air guide pipe (2).
2. An arrangement for elevating an operating mach number of an aircraft turbine engine according to claim 1, wherein said bleed air duct comprises a plurality of bleed air ducts circumferentially uniformly disposed about the outer periphery of the aircraft turbine engine.
3. Device for increasing the operating mach number of an aircraft turbine engine according to claim 1, wherein said bleed air duct (2) is provided with a lateral opening at the location of the connection to the compressor intermediate stage, from which said bleed air valve (3) can protrude into or out of the bleed air duct.
4. Device for increasing the operating mach number of an aircraft turbine engine according to claim 3, wherein said bleed valve (3) comprises a link ring driven in its axial movement along the engine by a drive mechanism, said link ring having a plurality of baffles extending in the axial direction of the engine, each baffle being inserted into or out of the bleed duct (2) from a lateral opening thereof.
5. Device for increasing the operating mach number of an aircraft turbine engine according to claim 1, wherein said bleed air duct (2) is supported outside said aircraft turbine engine by means of a bracket (4).
6. Device for increasing the operating mach number of an aircraft turbine engine according to claim 1, characterized in that swirl means (5) are provided in said bleed air duct (2), said swirl means being a plurality of staggered inclined blades.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202111401287.5A CN114017388A (en) | 2021-11-19 | 2021-11-19 | Device for improving working Mach number of aviation turbine engine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202111401287.5A CN114017388A (en) | 2021-11-19 | 2021-11-19 | Device for improving working Mach number of aviation turbine engine |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN114017388A true CN114017388A (en) | 2022-02-08 |
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Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202111401287.5A Pending CN114017388A (en) | 2021-11-19 | 2021-11-19 | Device for improving working Mach number of aviation turbine engine |
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| Country | Link |
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| CN (1) | CN114017388A (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115680903A (en) * | 2022-08-31 | 2023-02-03 | 中国航发四川燃气涡轮研究院 | Recyclable bypass bleed air control method |
Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| DE102016120682A1 (en) * | 2016-10-28 | 2018-05-03 | Rolls-Royce Deutschland Ltd & Co Kg | Aircraft bleed air system and method for providing bleed air in an aircraft engine |
| CN109252981A (en) * | 2018-10-25 | 2019-01-22 | 中国人民解放军空军工程大学 | Turbine/shock wave converges pinking combined engine |
| CN109339875A (en) * | 2018-09-21 | 2019-02-15 | 南京航空航天大学 | A kind of mixing diffuser of band bypass bleed |
| CN110440288A (en) * | 2019-07-26 | 2019-11-12 | 中国航发沈阳发动机研究所 | It is a kind of for premixing the inlet duct of fuel gas |
| US20190368421A1 (en) * | 2018-05-31 | 2019-12-05 | Rolls-Royce Deutschland Ltd & Co Kg | Gas turbine with rotating duct |
| CN111207005A (en) * | 2020-02-14 | 2020-05-29 | 中国航发沈阳发动机研究所 | Variable cycle engine mode control mechanism and intermediate casing structure with same |
-
2021
- 2021-11-19 CN CN202111401287.5A patent/CN114017388A/en active Pending
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| DE102016120682A1 (en) * | 2016-10-28 | 2018-05-03 | Rolls-Royce Deutschland Ltd & Co Kg | Aircraft bleed air system and method for providing bleed air in an aircraft engine |
| US20190368421A1 (en) * | 2018-05-31 | 2019-12-05 | Rolls-Royce Deutschland Ltd & Co Kg | Gas turbine with rotating duct |
| CN109339875A (en) * | 2018-09-21 | 2019-02-15 | 南京航空航天大学 | A kind of mixing diffuser of band bypass bleed |
| CN109252981A (en) * | 2018-10-25 | 2019-01-22 | 中国人民解放军空军工程大学 | Turbine/shock wave converges pinking combined engine |
| CN110440288A (en) * | 2019-07-26 | 2019-11-12 | 中国航发沈阳发动机研究所 | It is a kind of for premixing the inlet duct of fuel gas |
| CN111207005A (en) * | 2020-02-14 | 2020-05-29 | 中国航发沈阳发动机研究所 | Variable cycle engine mode control mechanism and intermediate casing structure with same |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115680903A (en) * | 2022-08-31 | 2023-02-03 | 中国航发四川燃气涡轮研究院 | Recyclable bypass bleed air control method |
| CN115680903B (en) * | 2022-08-31 | 2024-05-03 | 中国航发四川燃气涡轮研究院 | Recoverable bypass bleed air control method |
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