CN116025926A - Afterburner of aeroengine - Google Patents
Afterburner of aeroengine Download PDFInfo
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- CN116025926A CN116025926A CN202310229388.1A CN202310229388A CN116025926A CN 116025926 A CN116025926 A CN 116025926A CN 202310229388 A CN202310229388 A CN 202310229388A CN 116025926 A CN116025926 A CN 116025926A
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- radial
- wall
- stabilizer
- radial stabilizer
- cooling
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- 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
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Abstract
The application belongs to general combustion engine and aeroengine technical field, concretely relates to aeroengine afterburner, include: an outer casing; the converging ring is arranged in the outer casing, an outer culvert is formed between the converging ring and the outer casing, and a plurality of cooling air inlets distributed along the circumferential direction are formed in the side wall of the tail edge of the converging ring; the inner cone is arranged in the converging ring and forms connotation with the converging ring; the outer walls of the radial stabilizers are connected in the converging ring and point to the inner cone, the cross section of the radial stabilizers is V-shaped, the openings of the radial stabilizers are backwards, and tooth grooves distributed in a staggered manner are formed in the tail edges of the two side walls; the radial stabilizer inner walls are connected in the converging ring and correspondingly connected in the tail edges of the outer walls of the radial stabilizers, and cooling cavities are formed between the radial stabilizer inner walls and the corresponding radial stabilizer outer walls; each cooling cavity is correspondingly communicated with one cooling air inlet.
Description
Technical Field
The application belongs to the technical field of non-positive-displacement engines and jet propulsion device designs thereof, and particularly relates to an aircraft engine afterburner.
Background
In an aeroengine, an afterburner is connected to the rear end of a turbine, fuel is injected into the afterburner for reburning, so that the thrust is increased, the length of the afterburner is shorter and shorter along with the improvement of the performance requirement of the afterburner, and the temperature in the afterburner is higher and higher during operation, so that the following defects are generated:
1) The size of the afterburner is shortened, so that the content gas diffusion is insufficient, the uniformity state is difficult to achieve, the airflow speed of a back flow area behind the tail edge of the stabilizer is excessive, and ignition and flame connection in the afterburner are affected;
2) The temperature in the afterburner is increased, so that the risk of high-temperature ablation of the stabilizer is greatly increased, and therefore, the stabilizer is cooled by introducing external culvert gas, the external culvert gas flows out through the tail edge of the stabilizer and directly enters the rear reflux zone of the tail edge of the stabilizer, so that the temperature in the rear reflux zone of the tail edge of the stabilizer is reduced, ignition and continuous flame in the afterburner are influenced, and oscillation combustion in the combustion chamber is initiated;
3) The temperature in the afterburner is increased, the residual air coefficient is smaller, and during operation, periodic pressure pulsation and heat release pulsation are easy to occur, so that oscillation combustion is generated, strong mechanical vibration of the afterburner and an aeroengine main engine of the afterburner is caused, local overheating is generated, and local ablation is caused.
The present application has been made in view of the existence of the above-mentioned technical drawbacks.
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 that overcomes or mitigates at least one of the known technical drawbacks.
The technical scheme of the application is as follows:
an aircraft engine afterburner comprising:
an outer casing;
the converging ring is arranged in the outer casing, an outer culvert is formed between the converging ring and the outer casing, and a plurality of cooling air inlets distributed along the circumferential direction are formed in the side wall of the tail edge of the converging ring;
the inner cone is arranged in the converging ring and forms connotation with the converging ring;
the outer walls of the radial stabilizers are connected in the converging ring and point to the inner cone, the cross section of the radial stabilizers is V-shaped, the openings of the radial stabilizers are backwards, and tooth grooves distributed in a staggered manner are formed in the tail edges of the two side walls;
the radial stabilizer inner walls are connected in the converging ring and correspondingly connected in the tail edges of the outer walls of the radial stabilizers, and cooling cavities are formed between the radial stabilizer inner walls and the corresponding radial stabilizer outer walls; each cooling cavity is correspondingly communicated with one cooling air inlet.
According to at least one embodiment of the present application, in the aircraft engine afterburner described above, each radial stabilizer inner wall is V-shaped, opening rearward.
According to at least one embodiment of the present application, in the afterburner of an aeroengine, the inner cone has a hollow structure, the side wall has a plurality of cooling vent holes, and the cone end has a plurality of cooling vent holes;
the outer wall of each radial stabilizer and the inner wall of each radial stabilizer are connected to the inner cone;
each cooling cavity is correspondingly communicated with one cooling air vent.
According to at least one embodiment of the present application, in the aircraft engine afterburner described above, each radial stabilizer has a bleed air inlet on its outer wall, towards the front end;
the inner wall of each radial stabilizer is provided with a bleed air outlet which is opposite to the bleed air inlet;
the aircraft engine afterburner further comprises:
the plurality of bleed air pipes are correspondingly arranged in each cooling cavity, supported between the outer wall of the corresponding radial stabilizer and the inner wall of the radial stabilizer, and communicated with the corresponding bleed air inlet and bleed air outlet.
According to at least one embodiment of the present application, in the aircraft engine afterburner described above, there are a plurality of bleed air inlets on the outer wall of each radial stabilizer, and a plurality of bleed air outlets, bleed air ducts on the inner wall of the corresponding radial stabilizer, arranged radially.
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 edge of the outer casing.
The application has at least the following beneficial technical effects:
the utility model provides an aeroengine afterburner, its design comprises radial stabilizer outer wall, radial stabilizer inner wall, at aeroengine during operation, the fuel is inwards contained in the accessible oil spout pole, the fuel takes place evaporation, the atomizing under the effect of containing the air current, and can mix with containing the air current, enter into the back flow area behind each radial stabilizer trailing edge, and then ignite with the ignition mouth, or adopt the mode of jet ignition to ignite, continuous flame, and further organize the burning, simultaneously, can introduce partial outer air current into the cooling chamber of each radial stabilizer through each cooling air inlet, utilize outer air current to cool off each radial stabilizer, protect each radial stabilizer not to receive the high temperature ablation, in addition, design the tooth's socket that has the staggered distribution on each radial stabilizer outer wall two lateral wall trailing edges, can make the front of flame radially staggered distribution behind each radial stabilizer outer wall two lateral wall trailing edges, thereby make the front disturbance of flame offset each other behind each radial stabilizer outer wall two lateral wall trailing edges, thereby reduce the amplitude that causes the pressure pulsation in the afterburning chamber, can effectively restrain the amplitude of afterburning flame in the afterburning chamber, the afterburning effect can be guaranteed, the afterburning, even combustion can be guaranteed to the afterburning, the afterburning is evenly is easy to be burnt.
Drawings
FIG. 1 is a schematic illustration of an aircraft engine afterburner provided in an embodiment of the present application;
FIG. 2 is a view in the direction A of FIG. 1;
FIG. 3 is a partial cross-sectional view of FIG. 2;
wherein:
1-an outer casing; 2-confluence ring; 3-an inner cone; 4-radial stabilizer outer wall; 5-radial stabilizer inner wall; 6, an induced draft tube; 7-vibration-proof heat shield.
For better illustration of the present embodiment, some parts of the drawings may be omitted, enlarged or reduced, and do not represent the actual product size, and furthermore, the drawings are for illustrative purposes only and are not to be construed as limiting the present application.
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.
An aircraft engine afterburner comprising:
an outer casing 1;
the converging ring 2 is arranged in the outer casing 1, forms an outer culvert with the outer casing 1, and is provided with a plurality of cooling air inlets distributed along the circumferential direction on the side wall of the tail edge;
the inner cone 3 is arranged in the converging ring 2 and forms connotation with the converging ring 2;
the outer walls 4 of the radial stabilizers are connected in the converging ring 2 and point to the inner cone 3, the cross section of the radial stabilizers is V-shaped, the openings of the radial stabilizers are backward, and tooth grooves distributed in a staggered manner are formed in the tail edges of the two side walls;
a plurality of radial stabilizer inner walls 5 connected in the converging ring 2 and correspondingly connected in the trailing edge of each radial stabilizer outer wall 4, and cooling cavities are formed between the corresponding radial stabilizer outer walls 4; each cooling cavity is correspondingly communicated with one cooling air inlet.
In the afterburner of the aeroengine disclosed in the above embodiment, the radial stabilizer is designed to be formed by the outer wall 4 of the radial stabilizer and the inner wall 5 of the radial stabilizer, when the aeroengine works, fuel oil can be sprayed into the inner container through the fuel injection rod, the fuel oil is evaporated and atomized under the action of the inner container air flow and can be mixed with the inner container air flow, and the fuel oil enters into a backflow area behind the tail edge of each radial stabilizer to be ignited by an ignition nozzle or to be ignited and continuously flame in a jet ignition mode, and further the combustion is organized, meanwhile, part of the outer container air flow can be introduced into a cooling cavity of each radial stabilizer through each cooling air inlet, and each radial stabilizer is cooled by the outer container air flow to protect each radial stabilizer from high-temperature ablation.
According to the afterburner of the aeroengine disclosed by the embodiment, when the aeroengine works, the tooth grooves which are distributed in a staggered way are formed in the tail edges of the two side walls of the outer wall 4 of each radial stabilizer, so that the fronts of flames behind the tail edges of the two side walls of the outer wall 4 of each radial stabilizer are distributed in a staggered way along the radial direction, and the disturbance of the fronts of the flames behind the tail edges of the two side walls of the outer wall 4 of each radial stabilizer are mutually counteracted, so that the amplitude of pressure pulsation caused by the flames in the afterburner is reduced, the oscillation combustion in the afterburner can be effectively restrained, the uniform diffusion of content airflow is facilitated, and the ignition and flame connecting capacity of the afterburner is ensured.
In some alternative embodiments, in the afterburner of the aeroengine described above, each radial stabilizer inner wall 5 is V-shaped, opening backwards, i.e. a radial groove is configured in the radial stabilizer trailing edge, so as to be able to expand the range of the backflow zone behind the radial stabilizer trailing edge, ensuring the ignition and flame-carrying capacity of the afterburner.
In some alternative embodiments, in the afterburner of the aeroengine, the inner cone 3 has a hollow structure, the side wall has a plurality of cooling vent holes, and the cone end has a plurality of cooling vent holes;
each radial stabilizer outer wall 4 and each radial stabilizer inner wall 5 are connected to the inner cone 3;
each cooling cavity is correspondingly communicated with one cooling air vent.
According to the afterburner of the aeroengine disclosed by the embodiment, when the aeroengine works, the external air flow entering the inner cooling cavity of each radial stabilizer carries out the along-way cooling on each radial stabilizer, then the external air flow can flow into the inner cone body 3 through the corresponding cooling vent holes to cool the inner cone body 3, the inner cone body 3 is protected from high-temperature ablation, and finally the external air flow is discharged through each cooling vent hole at the conical end of the inner cone body 3, and the external air flow is not directly entering the backflow area behind the tail edges of each radial stabilizer any more, so that the ignition and flame connection capacities of the afterburner are ensured.
In some alternative embodiments, in the aircraft engine afterburner described above, each radial stabilizer outer wall 4 has bleed air inlets thereon, towards the front, i.e. at the leading edge of the radial stabilizer outer wall 4;
the inner wall 5 of each radial stabilizer is provided with a bleed air outlet, and the bleed air outlet is opposite to the bleed air inlet and is positioned at the front edge part of the inner wall 5 of the radial stabilizer;
the aircraft engine afterburner further comprises:
and a plurality of bleed air pipes 6 are correspondingly arranged in each cooling cavity, supported between the corresponding radial stabilizer outer walls 4 and the corresponding radial stabilizer inner walls 5, and communicated with the corresponding bleed air inlets and bleed air outlets.
According to the afterburner of the aeroengine disclosed by the embodiment, when the aeroengine works, after part of connotation air flow is directly introduced into the tail edges of the corresponding radial stabilizers through the air introduction pipes 6, the temperature reduction of the backflow areas affected by the connotation air flow after the tail edges of the radial stabilizers is compensated, the temperature of the backflow areas after the tail edges of the radial stabilizers is improved, the ignition and flame connection capability of the afterburner is ensured, flame stable combustion is facilitated, the initiation of oscillation combustion in the afterburner is prevented, and in addition, the flame lifting height of the tail edges of the radial stabilizers is improved, so that high-temperature flames are far away from the tail edges of the radial stabilizers, wall-hanging combustion is avoided, and the radial stabilizers are protected from high-temperature ablation.
In some alternative embodiments, in the aircraft engine afterburner described above, there are a plurality of bleed air inlets on each radial stabilizer outer wall 4 and a corresponding plurality of bleed air outlets on the radial stabilizer inner wall 5, bleed air pipes 6, arranged radially.
In some alternative embodiments, the aircraft engine afterburner described above further comprises:
the vibration-proof heat shield 7 is disposed in the outer casing 1 at the rear edge of the outer casing 1 to further suppress the oscillating combustion in the afterburner and to protect the rear edge of the outer casing 1 from high temperature ablation.
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 (6)
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), and the side wall of the tail edge of the converging ring is provided with a plurality of cooling air inlets distributed along the circumferential direction;
an inner cone (3) is arranged in the converging ring (2) and forms connotation with the converging ring (2);
the outer walls (4) of the radial stabilizers are connected in the converging ring (2) and are directed towards the inner cone (3), the cross section of the radial stabilizers is V-shaped, the openings of the radial stabilizers are backward, and tooth grooves which are distributed in a staggered manner are formed in the tail edges of the two side walls;
a plurality of radial stabilizer inner walls (5) connected in the converging ring (2) and correspondingly connected in the tail edge of each radial stabilizer outer wall (4), and cooling cavities are formed between the radial stabilizer inner walls and the corresponding radial stabilizer outer walls (4); each cooling cavity is correspondingly communicated with one cooling air inlet.
2. The aircraft engine afterburner of claim 1,
the inner wall (5) of each radial stabilizer is V-shaped, and the opening is backward.
3. The aircraft engine afterburner of claim 1,
the inner cone (3) is of a hollow structure, the side wall of the inner cone is provided with a plurality of cooling vent holes, and the cone end of the inner cone is provided with a plurality of cooling vent holes;
each radial stabilizer outer wall (4) and each radial stabilizer inner wall (5) are connected to the inner cone (3);
each cooling cavity is correspondingly communicated with one cooling air vent.
4. The aircraft engine afterburner of claim 1,
each radial stabilizer outer wall (4) is provided with a bleed air inlet towards the front end;
each radial stabilizer inner wall (5) is provided with a bleed air outlet which is opposite to the bleed air inlet;
the aircraft engine afterburner further comprises:
and a plurality of bleed air pipes (6) are correspondingly arranged in each cooling cavity and are supported between the corresponding outer wall (4) and the corresponding inner wall (5) of the radial stabilizer, and are communicated with the corresponding bleed air inlet and bleed air outlet.
5. The aircraft engine afterburner of claim 4,
the plurality of bleed air inlets on the outer wall (4) of each radial stabilizer, and the plurality of bleed air outlets and bleed air pipes (6) on the inner wall (5) of the corresponding radial stabilizer are arranged along the radial direction.
6. The aircraft engine afterburner of claim 1,
further comprises:
the vibration-proof heat shield (7) is arranged in the outer casing (1) and is positioned at the rear edge of the outer casing (1).
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310229388.1A CN116025926A (en) | 2023-03-10 | 2023-03-10 | Afterburner of aeroengine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310229388.1A CN116025926A (en) | 2023-03-10 | 2023-03-10 | Afterburner of aeroengine |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN116025926A true CN116025926A (en) | 2023-04-28 |
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Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202310229388.1A Pending CN116025926A (en) | 2023-03-10 | 2023-03-10 | Afterburner of aeroengine |
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| CN (1) | CN116025926A (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN117109029A (en) * | 2023-08-25 | 2023-11-24 | 西南科技大学 | Blunt body flame stabilizer and aeroengine combustion assembly |
-
2023
- 2023-03-10 CN CN202310229388.1A patent/CN116025926A/en active Pending
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN117109029A (en) * | 2023-08-25 | 2023-11-24 | 西南科技大学 | Blunt body flame stabilizer and aeroengine combustion assembly |
| CN117109029B (en) * | 2023-08-25 | 2024-02-02 | 西南科技大学 | Blunt body flame stabilizer and aeroengine combustion assembly |
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