CN116293799A - Afterburner of aero-engine and control method thereof - Google Patents
Afterburner of aero-engine and control method thereof Download PDFInfo
- Publication number
- CN116293799A CN116293799A CN202310157842.7A CN202310157842A CN116293799A CN 116293799 A CN116293799 A CN 116293799A CN 202310157842 A CN202310157842 A CN 202310157842A CN 116293799 A CN116293799 A CN 116293799A
- Authority
- CN
- China
- Prior art keywords
- oil
- jet ignition
- oil injection
- outer casing
- support plate
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Pending
Links
- 238000000034 method Methods 0.000 title claims abstract description 10
- 238000002347 injection Methods 0.000 claims abstract description 52
- 239000007924 injection Substances 0.000 claims abstract description 52
- 238000002485 combustion reaction Methods 0.000 claims abstract description 28
- 238000001816 cooling Methods 0.000 claims description 21
- 239000000446 fuel Substances 0.000 description 32
- 238000001514 detection method Methods 0.000 description 4
- 230000000694 effects Effects 0.000 description 3
- 239000000243 solution Substances 0.000 description 3
- 238000006467 substitution reaction Methods 0.000 description 2
- 238000000889 atomisation Methods 0.000 description 1
- 238000001704 evaporation Methods 0.000 description 1
- 230000008020 evaporation Effects 0.000 description 1
- 230000000149 penetrating effect Effects 0.000 description 1
- 230000000750 progressive effect Effects 0.000 description 1
- 239000007921 spray Substances 0.000 description 1
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
- F23R3/283—Attaching or cooling of fuel injecting means including supports for fuel injectors, stems, or lances
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
- F23R3/286—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply having fuel-air premixing devices
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Fluidized-Bed Combustion And Resonant Combustion (AREA)
Abstract
The application belongs to the technical field of aircraft engine afterburner design, and particularly relates to an aircraft engine afterburner and a control method thereof, wherein the aircraft engine afterburner comprises: an outer casing; the converging ring is arranged in the outer casing and forms an outer culvert with the outer casing; the inner cone is arranged in the converging ring and forms connotation with the converging ring; the support plates are supported between the converging ring and the inner cone along the circumferential direction, the inner part of the support plates is hollow, and the side walls of the support plates are provided with a plurality of oil injection holes; the oil injection rods penetrate through the outer casing and the converging ring; each oil injection rod correspondingly stretches into one support plate; the jet ignition oil nozzles penetrate through the outer wall of the main combustion chamber and extend into the rear end of the main combustion chamber; the number of jet ignition oil nozzles is half of the number of oil injection rods, and each jet ignition oil nozzle is correspondingly positioned between two adjacent oil injection rods.
Description
Technical Field
The application belongs to the technical field of aircraft engine afterburner design, and particularly relates to an aircraft engine afterburner and a control method thereof.
Background
The rear end of the main combustion chamber in the aero-engine is connected with the turbine, the afterburner is connected with the rear end of the turbine, and fuel is injected into the afterburner for re-combustion, so that the thrust of the aero-engine is increased, the outlet temperature of the combustion chamber is higher and higher along with the improvement of the performance of the aero-engine, the detectable performance of the aero-engine is seriously damaged, and the aero-engine is easy to detect.
The present application has been made in view of the existence of the drawbacks of the prior art.
It should be noted that the above disclosure of the background art is only for aiding in understanding the inventive concept and technical solution of the present invention, which is not necessarily prior art to the present application, and should not be used for evaluating the novelty and the creativity of the present application in the case where no clear evidence indicates that the above content has been disclosed at the filing date of the present application.
Disclosure of Invention
It is an object of the present application to provide an aircraft engine afterburner and a method of operating the same which overcome or mitigate at least one of the known technical drawbacks.
The technical scheme of the application is as follows:
in one aspect, an aircraft engine afterburner is provided, comprising:
an outer casing;
the converging ring is arranged in the outer casing and forms an outer culvert with the outer casing;
the inner cone is arranged in the converging ring and forms connotation with the converging ring;
the support plates are supported between the converging ring and the inner cone along the circumferential direction, the inner part of the support plates is hollow, and the side walls of the support plates are provided with a plurality of oil injection holes;
the oil injection rods penetrate through the outer casing and the converging ring; each oil injection rod correspondingly stretches into one support plate;
the jet ignition oil nozzles penetrate through the outer wall of the main combustion chamber and extend into the rear end of the main combustion chamber;
the number of jet ignition oil nozzles is half of the number of oil injection rods, and each jet ignition oil nozzle is correspondingly positioned between two adjacent oil injection rods.
According to at least one embodiment of the present application, in the aircraft engine afterburner described above, the cross-section of each of the struts is curved, shielding the turbine.
According to at least one embodiment of the present application, in the afterburner of an aeroengine, the converging ring is provided with a plurality of cooling air inlets, and each cooling air inlet is correspondingly communicated with the inside of one support plate;
the side wall of each support plate is provided with a plurality of support plate cooling air outlet holes;
the inner cone body is hollow, the side wall is provided with a plurality of cooling air vents, the front end is provided with a plurality of inner cone cooling air outlet holes distributed along the circumferential direction, and each cooling air vent is correspondingly communicated with the inside of one support plate.
According to at least one embodiment of the present application, in the aircraft engine afterburner described above, further comprising:
the vibration-proof heat shield is arranged in the outer casing and is positioned at the rear end of the outer casing.
Another aspect provides a method of aircraft engine afterburner manipulation comprising:
two oil injection rods corresponding to one jet ignition oil nozzle are combined into a group;
in the full stress state, each group of oil injection rods is used for supplying oil, and each jet ignition oil nozzle is used for jet ignition and combustion;
in the small stress state, a group of oil injection rods are used for supplying oil, and corresponding jet ignition oil nozzles are used for jet ignition and combustion;
and in the intermediate state between the full stress state and the small stress state, the oil is supplied by half groups of oil injection rods, the corresponding jet ignition oil nozzle is used for jet ignition and combustion, and the oil injection rods of the non-oil supply group and the oil injection rods of the oil supply group are distributed alternately.
Drawings
FIG. 1 is a schematic illustration of the operation of an aircraft engine afterburner provided in an embodiment of the present application;
FIG. 2 is a partial schematic view of an aircraft engine afterburner provided in an embodiment of the present application;
FIG. 3 is a schematic cross-sectional view of a support plate provided in an embodiment of the present application;
wherein:
1-an outer casing; 2-confluence ring; 3-an inner cone; 4-supporting plates; 5-an oil injection rod; 6-jet ignition oil nozzle; 7-a main combustion chamber; 8-a turbine; 9-vibration-proof heat shields.
For the purpose of better illustrating the present embodiments, certain elements of the drawings may be omitted, enlarged or reduced and do not represent the actual product dimensions, and furthermore, the drawings are for illustrative purposes only and are not to be construed as limiting the present patent.
Detailed Description
In order to make the technical solution of the present application and the advantages thereof more apparent, the technical solution of the present application will be more fully described in detail below with reference to the accompanying drawings, it being understood that the specific embodiments described herein are only some of the embodiments of the present application, which are for explanation of the present application, not for limitation of the present application. It should be noted that, for convenience of description, only the portion relevant to the present application is shown in the drawings, and other relevant portions may refer to a general design, and without conflict, the embodiments and technical features in the embodiments may be combined with each other to obtain new embodiments.
Furthermore, unless defined otherwise, technical or scientific terms used in the description of this application should be given the ordinary meaning as understood by one of ordinary skill in the art to which this application belongs. The terms "upper," "lower," "left," "right," "center," "vertical," "horizontal," "inner," "outer," and the like as used in this description are merely used to indicate relative directions or positional relationships, and do not imply that a device or element must have a particular orientation, be configured and operated in a particular orientation, and that the relative positional relationships may be changed when the absolute position of the object being described is changed, and thus should not be construed as limiting the present application. The terms "first," "second," "third," and the like, as used in the description herein, are used for descriptive purposes only and are not to be construed as indicating or implying any particular importance to the various components. The use of the terms "a," "an," or "the" and similar referents in the description of the invention are not to be construed as limited in number to the precise location of at least one. As used in this description, the terms "comprises," "comprising," or the like are intended to cover an element or article that appears before the term and that is listed after the term and its equivalents, without excluding other elements or articles.
Furthermore, unless specifically stated and limited otherwise, the terms "mounted," "connected," and the like in the description herein are to be construed broadly and refer to either a fixed connection, a removable connection, or an integral connection, for example; can be mechanically or electrically connected; can be directly connected or indirectly connected through an intermediate medium, and can also be communicated with the inside of two elements, and the specific meaning of the two elements can be understood by a person skilled in the art according to specific situations.
The present application is described in further detail below in conjunction with fig. 1-3.
In one aspect, an aircraft engine afterburner is provided, comprising:
an outer casing 1;
the converging ring 2 is arranged in the outer casing 1 and forms an outer culvert with the outer casing 1;
the inner cone 3 is arranged in the converging ring 2, forms connotation with the converging ring 2, and is connected with the converging ring 2 and the inner cone 3 on corresponding parts at the rear end of the turbine 8;
the support plates 4 are supported between the converging ring 2 and the inner cone 3 along the circumferential direction, are hollow in the converging ring, and are provided with a plurality of oil spray holes on the side walls, and are distributed close to the tail edges of the support plates 4;
a plurality of oil injection rods 5 penetrating the outer casing 1 and the converging ring 2; each oil injection rod 5 correspondingly stretches into one support plate 4;
the jet ignition oil nozzles 6 penetrate through the outer wall of the main combustion chamber 7 and extend into the rear end of the main combustion chamber 7, and the rear end of the main combustion chamber 7 is connected with the front end of the turbine 8;
the number of jet ignition fuel nozzles 6 is half of the number of fuel injection rods 5, and each jet ignition fuel nozzle 6 is correspondingly positioned between two adjacent fuel injection rods 5.
According to the afterburner of the aeroengine disclosed by the embodiment, when the aeroengine works, fuel can be supplied through each fuel injection rod 5, fuel injected by each fuel injection rod 5 enters the corresponding support plate 4 and is injected from the fuel injection holes on the side wall of the support plate 4, so that the fuel is subjected to the action of connotation air flow, evaporation and atomization are carried out, the fuel is mixed with connotation air flow, and enters a backflow area formed after the tail edge of each support plate 4, so that mixed fuel is formed, meanwhile, each jet ignition fuel injection nozzle 6 injects fuel into the main combustion chamber 7, the fuel in the main combustion chamber is combusted under the action of high temperature, so that a flame is formed, each flame passes through the turbine 8 and enters the afterburner, and the mixed fuel in the backflow area corresponding to the tail edges of the two support plates 4 is ignited by adopting full-space jet ignition without Zhou Xianglian flame, so that the possibility of radar detection can be effectively reduced.
The afterburner of the aeroengine disclosed in the above embodiment can be designed to be operated by the following method when the aeroengine is in operation:
two fuel injection rods 5 corresponding to one jet ignition fuel injection nozzle 6 are in a group;
in the full stress state, each group of oil injection rods 5 is used for supplying oil, and each jet ignition oil nozzle 6 is used for jet ignition and combustion;
in the small stress state, a group of oil injection rods 5 are used for supplying oil, and the corresponding jet ignition oil nozzle 6 is used for jet ignition and combustion;
in the intermediate state between the full stress state and the small stress state, the fuel injection rods 5 of a half group are used for fuel supply, the corresponding jet ignition fuel injection nozzle 6 is used for jet ignition and combustion, and the fuel injection rods 5 of the fuel supply group are distributed alternately.
According to the afterburner of the aeroengine disclosed by the embodiment, when the aeroengine works, according to different stress application states, corresponding partial or all oil injection rods 5 can be flexibly selected for oil supply, and the corresponding jet ignition oil nozzles 6 are used for jet ignition and combustion, so that different stress application states can be well adapted, and the fuel utilization efficiency is improved.
In some alternative embodiments, in the aircraft engine afterburner described above, the cross-section of each support plate 4 is curved, shielding the turbine 8, i.e. the high temperature components at the aft end, thereby reducing the likelihood of infrared detection.
In some alternative embodiments, in the afterburner of the aeroengine described above, the converging ring 2 has a plurality of cooling inlets, each of which communicates with the inside of one of the brackets 4;
the side wall of each support plate 4 is provided with a plurality of support plate cooling air outlet holes;
the inner cone 3 is hollow, the side wall is provided with a plurality of cooling air vents, the front end is provided with a plurality of inner cone cooling air outlet holes distributed along the circumferential direction, and each cooling air vent is correspondingly communicated with the inside of one support plate 4.
According to the aeroengine afterburner disclosed by the embodiment, when the aeroengine works, the outer culvert air flow is introduced into the support plate 4 through the cooling air inlet on the converging ring 2, part of the outer culvert air flow flows out through the support plate cooling air outlet holes on the side wall of the support plate 4, part of the outer culvert air flow enters the inner cone 3 through the cooling air inlet on the side wall of the inner cone 3, and flows out through the inner cone cooling air outlet holes at the front end of the inner cone 3, so that the support plate 4 and the inner cone 3 can be cooled by the outer culvert air flow efficiently, the support plate 4 and the inner cone 3 are prevented from being ablated by high temperature, and the possibility of infrared detection can be reduced.
In some alternative embodiments, the aircraft engine afterburner described above further comprises:
the vibration-proof heat shield 9 is arranged in the outer casing 1 and positioned at the rear end of the outer casing 1, and can be connected to the inner wall of the rear end of the outer casing 1 through a bracket, so that the vibration-proof heat shield has vibration-proof and heat-insulating effects.
In some alternative embodiments, in the aircraft engine afterburner described above, the trailing edge of each strut 4 and its inner thrust body 3 are of a radar-modified design, thereby reducing the likelihood of detection by radar.
Another aspect provides a method of aircraft engine afterburner manipulation comprising:
two fuel injection rods 5 corresponding to one jet ignition fuel injection nozzle 6 are in a group;
in the full stress state, each group of oil injection rods 5 is used for supplying oil, and each jet ignition oil nozzle 6 is used for jet ignition and combustion;
in the small stress state, a group of oil injection rods 5 are used for supplying oil, and the corresponding jet ignition oil nozzle 6 is used for jet ignition and combustion;
in the intermediate state between the full stress state and the small stress state, the fuel injection rods 5 of a half group are used for fuel supply, the corresponding jet ignition fuel injection nozzle 6 is used for jet ignition and combustion, and the fuel injection rods 5 of the fuel supply group are distributed alternately.
For the control method of the afterburner of the aeroengine disclosed in the foregoing embodiment, the description is simpler, and specific relevant points can be referred to the description of the afterburner of the aeroengine, and the technical effects of the afterburner of the aeroengine can also be referred to the technical effects of the relevant parts of the afterburner of the aeroengine, which are not repeated herein.
In the description, each embodiment is described in a progressive manner, and each embodiment is mainly described by the differences from other embodiments, so that the same similar parts among the embodiments are mutually referred.
Having thus described the technical aspects of the present application with reference to the preferred embodiments illustrated in the accompanying drawings, it should be understood by those skilled in the art that the scope of the present application is not limited to the specific embodiments, and those skilled in the art may make equivalent changes or substitutions to the relevant technical features without departing from the principles of the present application, and those changes or substitutions will now fall within the scope of the present application.
Claims (5)
1. An aircraft engine afterburner, comprising:
an outer casing (1);
the converging ring (2) is arranged in the outer casing (1) and forms an outer culvert with the outer casing (1);
an inner cone (3) is arranged in the converging ring (2) and forms connotation with the converging ring (2);
the support plates (4) are supported between the converging ring (2) and the inner cone (3) along the circumferential direction, the inner part of the support plates is hollow, and the side walls of the support plates are provided with a plurality of oil injection holes;
the oil injection rods (5) penetrate through the outer casing (1) and the converging ring (2); each oil injection rod (5) correspondingly extends into one support plate (4);
a plurality of jet ignition oil nozzles (6) penetrate through the outer wall of the main combustion chamber (7) and extend into the rear end of the main combustion chamber (7);
the number of jet ignition oil nozzles (6) is half of the number of the oil injection rods (5), and each jet ignition oil nozzle (6) is correspondingly positioned between two adjacent oil injection rods (5).
2. The aircraft engine afterburner of claim 1,
the cross section of each support plate (4) is bent to cover the turbine (8).
3. The aircraft engine afterburner of claim 1,
the converging ring (2) is provided with a plurality of cooling air inlets, and each cooling air inlet is correspondingly communicated with the inside of one support plate (4);
the side wall of each support plate (4) is provided with a plurality of support plate cooling air outlet holes;
the inner cone (3) is hollow, the side wall is provided with a plurality of cooling air vents, the front end is provided with a plurality of inner cone cooling air outlet holes distributed along the circumferential direction, and each cooling air vent is correspondingly communicated with the inside of one support plate (4).
4. The aircraft engine afterburner of claim 1,
further comprises:
the vibration-proof heat shield (9) is arranged in the outer casing (1) and is positioned at the rear end of the outer casing (1).
5. A method of aircraft engine afterburner operation, comprising:
two injection rods (5) corresponding to one jet ignition injection nozzle (6) are grouped;
in the full stress state, oil is supplied by each group of oil injection rods (5), and jet ignition and combustion are carried out by each jet ignition oil nozzle (6);
in a small stress state, a group of oil injection rods (5) are used for supplying oil, and corresponding jet ignition oil nozzles (6) are used for jet ignition and combustion;
in the intermediate state between the full stress state and the small stress state, a half group of oil injection rods (5) are used for supplying oil, and the corresponding jet ignition oil nozzles (6) are used for jet ignition and combustion, and the oil injection rods (5) of the non-oil supply group and the oil injection rods (5) of the oil supply group are distributed alternately.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310157842.7A CN116293799A (en) | 2023-02-23 | 2023-02-23 | Afterburner of aero-engine and control method thereof |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310157842.7A CN116293799A (en) | 2023-02-23 | 2023-02-23 | Afterburner of aero-engine and control method thereof |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN116293799A true CN116293799A (en) | 2023-06-23 |
Family
ID=86782706
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202310157842.7A Pending CN116293799A (en) | 2023-02-23 | 2023-02-23 | Afterburner of aero-engine and control method thereof |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN116293799A (en) |
-
2023
- 2023-02-23 CN CN202310157842.7A patent/CN116293799A/en active Pending
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| US5385015A (en) | Augmentor burner | |
| EP0315486B1 (en) | Aircraft engine frame construction | |
| US10101030B2 (en) | Gas turbine engines with plug resistant effusion cooling holes | |
| KR101211814B1 (en) | Combustion burner | |
| US4009569A (en) | Diffuser-burner casing for a gas turbine engine | |
| EP3018416B1 (en) | Impingement film-cooled floatwall with backside feature | |
| CN107270328B (en) | Closed trapped vortex cavity pilot for gas turbine engine amplifier | |
| EP3039340B1 (en) | Vena contracta swirling dilution passages for gas turbine engine combustor | |
| EP2180256A2 (en) | Dual wall structure for use in a combustor of a gas turbine engine | |
| JP5080815B2 (en) | Exhaust duct flow splitter system | |
| US7908868B2 (en) | Device for mounting an air-flow dividing wall in a turbojet engine afterburner | |
| CN115200037B (en) | Afterburner of aeroengine | |
| US3633361A (en) | Burners for reheat combustion chambers | |
| EP3039345A2 (en) | Dual fuel nozzle with liquid filming atomization for a gas turbine engine | |
| CN110726157A (en) | Fuel nozzle cooling structure | |
| CN115200043B (en) | Annular stabilizer ignition afterburner | |
| CN108019778B (en) | Fuel nozzle assembly with impingement purge | |
| EP2573464B1 (en) | Combustion sections of gas turbine engines with convection shield assemblies | |
| JP2014219195A (en) | Wake manipulating structure for turbine system | |
| CN116293799A (en) | Afterburner of aero-engine and control method thereof | |
| CN115164230B (en) | Afterburner of aircraft engine | |
| CN115200045B (en) | External culvert ignition afterburner | |
| CN115183270A (en) | Wall-mounted stabilizer ignition afterburning chamber | |
| US11092076B2 (en) | Turbine engine with combustor | |
| CN116025926B (en) | Afterburner of aeroengine |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination |