阅读说明：本技术 一种航空发动机燃烧室头部dbd等离子体叶片式轴向旋流器 (Aircraft engine combustion chamber head DBD plasma vane type axial swirler ) 是由 邓俊 王平 罗艳春 吕雪燕 崔连柱 张华磊 焦胜博 周一鹏 王宇天 孙杨 于 2021-09-01 设计创作，主要内容包括：本发明公开了一种航空发动机燃烧室头部DBD等离子体叶片式轴向旋流器,包括内环接地电极、外环高压电极、旋流器主体件、导线；旋流器主体件为回转体结构,通过DLP光固化3D陶瓷打印一体成型；外环高压电极粘贴于外环绝缘阻挡介质层的外表面,内环接地电极粘贴于内环绝缘阻挡介质层的内表面；外环绝缘阻挡介质层与内环绝缘阻挡介质层之间为旋流器叶片；导线一端焊接在外环高压电极的外侧面,另一端连接等离子体电源输出的高压端接线柱。通过在航空发动机燃烧室头部安装本发明,可在不影响燃烧室头部原有气流通道结构的情况下,在旋流器叶片的气流通道位置生成高浓度的活性粒子,从而达到强化燃烧室燃烧、扩大点熄火边界和提高燃烧效率的目的。(The invention discloses a head DBD plasma blade type axial swirler of an aircraft engine combustion chamber, which comprises an inner ring grounding electrode, an outer ring high-voltage electrode, a swirler main body part and a lead, wherein the inner ring grounding electrode is arranged on the outer ring high-voltage electrode; the cyclone main body part is of a revolving body structure and is integrally formed by DLP photocuring 3D ceramic printing; the outer ring high-voltage electrode is pasted on the outer surface of the outer ring insulation barrier dielectric layer, and the inner ring grounding electrode is pasted on the inner surface of the inner ring insulation barrier dielectric layer; a swirler vane is arranged between the outer ring insulation blocking dielectric layer and the inner ring insulation blocking dielectric layer; one end of the lead is welded on the outer side surface of the outer ring high-voltage electrode, and the other end of the lead is connected with a high-voltage terminal output by the plasma power supply. By installing the invention on the head of the combustion chamber of the aeroengine, high-concentration active particles can be generated at the position of the airflow channel of the swirler vane under the condition of not influencing the original airflow channel structure of the head of the combustion chamber, thereby achieving the purposes of strengthening combustion of the combustion chamber, expanding the point flameout boundary and improving the combustion efficiency.)

1. A DBD plasma vane type axial swirler of a head part of a combustion chamber of an aeroengine is characterized by comprising an inner ring grounding electrode (1), an outer ring high-voltage electrode (2), a swirler main body part (3) and a lead (4); the cyclone main body part (3) is of a revolving body structure and is integrally formed by DLP photocuring 3D ceramic printing; the cyclone main body part (3) comprises a cyclone main body part body, an outer ring insulation blocking dielectric layer (8) is arranged on the outer ring of the cyclone main body part body, an inner ring insulation blocking dielectric layer (6) is arranged on the inner ring of the cyclone main body part body, and cyclone blades (7) are arranged between the outer ring insulation blocking dielectric layer (8) and the inner ring insulation blocking dielectric layer (6); the outer ring high-voltage electrode (2) is adhered to the outer surface of the outer ring insulation barrier dielectric layer (8), and the inner ring grounding electrode (1) is adhered to the inner surface of the inner ring insulation barrier dielectric layer (6); one end of a lead (4) is welded on the outer side surface of the outer ring high-voltage electrode (2), and the other end of the lead (4) is connected with a high-voltage terminal of the plasma power supply output.

2. An aircraft engine combustion chamber head DBD plasma vane axial swirler as claimed in claim 1, characterized in that the swirler body further provides a swirler inlet nozzle seat (5), a swirler outlet seat (9); the outer ring insulation blocking medium layer (8) extends upwards to form a cyclone outlet seat (9), a shoulder is arranged between the outer ring insulation blocking medium layer (8) and the cyclone outlet seat (9), a flange is arranged at the bottom of the outer ring insulation blocking medium layer (8), and radial cyclone main body part assembling and fixing holes (10) are symmetrically formed in the upper end of the cyclone outlet seat (9); the inner ring insulation blocking dielectric layer (6) extends downwards to form a swirler inlet nozzle seat (5); after a fuel nozzle (14) of the combustion chamber of the aircraft engine is sleeved with a swirler inlet nozzle seat (5), the outlet of the fuel nozzle is flush with the outlet plane of a swirler vane (7).

3. The aircraft engine combustion chamber head DBD plasma blade type axial swirler of claim 1, characterized in that the inner ring grounding electrode (1) is a copper foil with a certain thickness, the outer side surface of the inner ring grounding electrode (1) is a glue surface and is adhered to the inner surface of the inner ring insulation barrier medium layer (6), and the inner side surface of the inner ring grounding electrode (1) is grounded after being contacted with a metal fuel nozzle; the outer ring high-voltage electrode (2) is a copper foil with a certain thickness, the inner side surface of the outer ring high-voltage electrode (2) is a rubber surface and is adhered to the outer surface of the outer ring insulation barrier dielectric layer (8), and the outer side surface of the outer ring high-voltage electrode (2) is sealed through silicon rubber.

4. The aircraft engine combustion chamber head DBD plasma vane type axial swirler of claim 3, characterized in that the shape of the inner ring grounding electrode (1) is a cylindrical cylinder, and the copper foil thickness d of the inner ring grounding electrode is₁0.01-0.5 mm, inner diameter D₁16-20 mm in length L₁Is 6-20 mm.

5. An aircraft engine combustion chamber head DBD plasma vane axial swirler as claimed in claim 3, characterized in that the outer ring high voltage electrode (2) is cylindrical in shape with copper foil thickness d₂0.01-0.5 mm, inner diameter D₂30-50 mm, length L₂6-20 mm.

6. An aircraft engine combustion chamber head DBD plasma vane axial swirler as claimed in claim 1, characterized in that the swirler inner diameter d of the swirler body (3) is_s16-24 mm, the outer diameter D of the cyclone_sIs 30 to 50 mm.

7. An aircraft engine combustion chamber head DBD plasma vane axial swirler as claimed in claim 1, characterized in that the swirler vanes (7) are shaped as straight vanes or twisted vanes.

8. The aircraft engine combustion chamber head DBD plasma vane axial swirler of claim 1, characterized in that the number n of swirler vanes (7) is 5-12.

9. An aircraft engine combustion chamber head DBD plasma vane axial swirler as claimed in claim 1, characterized in that the mounting angle β of the swirler vanes (7) is 40 ° to 70 °; the width delta d of a discharge gap between the swirler vanes (7) is 3-10 mm.

10. The aircraft engine combustion chamber head DBD plasma vane axial swirler of claim 1, characterized in that the thickness δ of the inner ring insulation barrier medium layer (6) is₃0.5-2 mm, the thickness delta of the outer ring insulation barrier dielectric layer (8)₄Is 0.5 to 2 mm.

Technical Field

The invention relates to a novel intensified combustion technology in the field of aircraft engines, in particular to a head dielectric barrier discharge plasma blade type axial swirler of an aircraft engine combustion chamber, wherein dielectric barrier discharge is called DBD for short.

Background

Along with the rise of flying height such as aircraft, unmanned aerial vehicle, under the adverse conditions such as high altitude, little table speed, aeroengine will face the unstable, the flame-out border narrows down, combustion efficiency step down scheduling problem of burning. In order to ensure stable work of the combustion chamber of the aircraft engine, at present, most researches attempt to improve the combustion process through a local structure of the combustion chamber, such as a multi-annular-cavity flame tube structure, a novel fuel nozzle design, a multi-stage swirl head and the like. By changing the structure, the combustion stability range of the combustion chamber is widened, and the aircraft engine with higher reliability is obtained. With the development of aircraft engine technology, the structure of the engine combustion chamber becomes more and more complex. As a novel intensified combustion technology, the plasma combustion-supporting technology has great prospect for being used for the combustion performance of an aircraft engine combustion chamber, is concerned by many researchers, and provides a new possible way for improving the performance of the aircraft engine combustion chamber. Dielectric barrier discharge, a form of plasma generation, has many application scenarios for combustion assistance due to its advantage of generating uniform, large-volume plasma. Due to the limitation of the traditional mechanical processing, numerical control processing and material process, the application difficulty of the dielectric barrier discharge plasma combustion-supporting mode in the aircraft engine is still large.

In aircraft engine combustion chambers, the swirler at the nose is a very important component. The performance of the combustion chamber directly and comprehensively influences the comprehensive performance of the combustion chamber. The swirler generates high-speed rotating jet flow at the head of the flame tube to form a low-pressure area, so that a hot reflux area is formed, and the stability of flame in the combustion chamber is ensured; at the same time, the rotating jet improves fuel atomization and oil-gas mixing. In addition, the swirler provides a proper amount of air for the combustion chamber head, and the head is ensured to have a proper residual air coefficient. The swirler is arranged at the head of the flame tube in the combustion chamber, is close to the flame area of the combustion chamber, and has higher working environment temperature. The swirler is well matched with the fuel nozzle and the head of the flame tube of the combustion chamber, so that the combustion chamber obtains good combustion performance. Vane-type axial swirlers are still used by most aircraft engines today. The axial swirler has small volume and large difficulty in special-shaped machining, and the middle of the swirler is directly matched with a fuel nozzle. According to the position and the structural characteristics of the axial swirler, if the combustion performance of the combustion chamber is improved by applying plasma to the head of the combustion chamber, not only the safety problem caused by high-voltage discharge of a plasma device is considered, but also the problem that the processing technology of the special-shaped insulator is difficult needs to be overcome.

Disclosure of Invention

In order to solve the problems that the flameout boundary of a combustion chamber of an aero-engine is narrowed and the combustion efficiency is reduced under severe conditions of high altitude, small surface speed and the like, the invention provides a DBD plasma blade type axial swirler at the head of the aero-engine combustion chamber.

The purpose of the invention is realized by the following technical scheme, which is combined with the attached drawings:

a DBD plasma blade type axial swirler of a head part of a combustion chamber of an aircraft engine comprises an inner ring grounding electrode 1, an outer ring high-voltage electrode 2, a swirler main body part 3 and a lead 4; the cyclone main body part 3 is of a revolving body structure and is integrally formed by DLP photocuring 3D ceramic printing; the cyclone main body part 3 comprises a cyclone main body part body, an outer ring insulation blocking dielectric layer 8 is arranged on the outer ring of the cyclone main body part body, an inner ring insulation blocking dielectric layer 6 is arranged on the inner ring of the cyclone main body part body, and cyclone blades 7 are arranged between the outer ring insulation blocking dielectric layer 8 and the inner ring insulation blocking dielectric layer 6; the outer ring high-voltage electrode 2 is adhered to the outer surface of the outer ring insulation barrier dielectric layer 8, and the inner ring grounding electrode 1 is adhered to the inner surface of the inner ring insulation barrier dielectric layer 6; one end of a lead 4 is welded on the outer side surface of the outer ring high-voltage electrode 2, and the other end of the lead 4 is connected with a high-voltage terminal of the plasma power supply output.

Further, the cyclone main body part body is also provided with a cyclone inlet nozzle seat 5 and a cyclone outlet seat 9; the outer ring insulation blocking medium layer 8 extends upwards to form a cyclone outlet seat 9, a shoulder is arranged between the outer ring insulation blocking medium layer 8 and the cyclone outlet seat 9, a flange is arranged at the bottom of the outer ring insulation blocking medium layer 8, and radial cyclone main body part assembling and fixing holes 10 are symmetrically formed in the upper end of the cyclone outlet seat 9; the inner ring insulation barrier medium layer 6 extends downwards to form a swirler inlet nozzle seat 5; after a fuel nozzle 14 of the aircraft engine combustion chamber is sleeved with the swirler inlet nozzle seat 5, the outlet of the fuel nozzle is flush with the outlet plane of the swirler vane 7.

Furthermore, the inner ring grounding electrode 1 is a copper foil with a certain thickness, the outer side surface of the inner ring grounding electrode 1 is a glue surface and is adhered to the inner surface of the inner ring insulation barrier dielectric layer 6, and the inner side surface of the inner ring grounding electrode 1 is in contact with the metal fuel nozzle and then is grounded; the outer ring high-voltage electrode 2 is a copper foil with a certain thickness, the inner side surface of the outer ring high-voltage electrode 2 is a rubber surface and is adhered to the outer surface of the outer ring insulation barrier dielectric layer 8, and the outer side surface of the outer ring high-voltage electrode 2 is sealed through silicon rubber.

Preferably, the inner ring ground electrode 1 has a cylindrical shape with a copper foil thickness d ₁0.01-0.5 mm, inner diameter D₁16-20 mm in length L₁6-20 mm;

preferably, the outer ring high voltage electrode 2 is cylindrical and has a copper foil thickness d₂0.01-0.5 mm, inner diameter D₂30-50 mm, length L₂6-20 mm.

Preferably, the cyclone body member 3 has a cyclone inner diameter d_s16-24 mm, the outer diameter D of the cyclone_sIs 30 to 50 mm.

Further, the swirler vanes 7 of the swirler body 3 are shaped and configured as straight vanes or twisted vanes.

Preferably, the number n of the swirler vanes 7 is 5 to 12.

Preferably, the installation angle β of the swirler vane 7 is 40 ° to 70 °; the width delta d of a discharge gap between the swirler vanes 7 is 3-10 mm.

Preferably, the thickness delta of the inner ring insulation barrier dielectric layer 6₃0.5-2 mm, and the thickness delta of the outer ring insulation barrier dielectric layer 8₄Is 0.5 to 2 mm.

The working principle of the invention is as follows:

the working medium of the invention is air. The inner ring grounding electrode 1, the outer ring high-voltage electrode 2, the inner ring insulation barrier medium layer and the outer ring insulation barrier medium layer of the cyclone main body part 3 form a double-medium-layer dielectric barrier discharge combustion-supporting exciter device. Discharge is driven between the inner ring grounding electrode 1 and the outer ring high-voltage electrode 2 through a single high-voltage plasma power supply, and a plasma discharge area is formed between airflow passage positions of the swirler vanes of the swirler main body part 3. When the gas passes through the dielectric barrier discharge area of the airflow channel of the cyclone blade, a high-voltage alternating current power supply is applied to two ends of the electrode to discharge the gas in the discharge area, and active particles (such as ozone, free oxygen atoms, excited molecular atoms, active groups and the like) with large volume and high concentration are formed. Under the blowing action of the airflow, the gas mixed with the plasma moves towards the outlet of the cyclone, enters a main combustion area at the head of a flame tube of the combustion chamber, and then participates in the processes of atomized oil-gas mixing, fuel cracking, combustion and the like, so that the chemical reaction of combustion is accelerated, and the purpose of improving the combustion performance of the combustion chamber is achieved.

The invention has the following beneficial effects:

the invention has simple manufacturing and installation process and simultaneously combines the structural characteristics of the coaxial circular tube double-layer dielectric barrier discharge plasma generator and the blade type axial swirler. The plasma combustion-supporting exciter device is arranged at the head part of the combustion chamber. The device provided by the invention is driven by a single high-voltage power supply on the premise of not changing the airflow channel structure of the head of the combustion chamber of the original aero-engine, and can generate large-volume high-concentration plasma active particles at the position of the airflow channel of the blade of the swirler, so that the device can be used for widening the combustion stability range, improving the uniformity of the temperature field at the outlet of the combustion chamber and improving the combustion efficiency, thereby strengthening the combustion effect of the combustion chamber of the aero-engine and improving the combustion performance of the combustion chamber.

The invention has simple mechanical structure, and breaks through the limitation of the manufacturing process of the traditional metal and alloy parts by adopting the DLP photocuring ceramic 3D printing technology. Under the condition of high temperature, the ceramic material has good mechanical strength and lower density, and the oxidation resistance is better than that of metals and alloys under the condition of being heated. By modifying the head swirler of the combustion chamber of the aircraft engine into the dielectric barrier discharge plasma swirler made of ceramic materials, the combustion-supporting performance of the plasma combustion-supporting exciter can be obtained, the weight of the combustion chamber can be effectively reduced, and the oxidation resistance of the head swirler of the combustion chamber is improved.

The invention has strong practicability and universality, can be used in most of aeroengines using the axial swirler, and is also suitable for other combustors with blade type axial swirlers.

Drawings

In order to more clearly illustrate the technical solutions in the embodiments of the present invention, the drawings used in the description of the embodiments of the present invention will be briefly described below, and it is obvious that the drawings in the following description are only some embodiments of the present invention, and it is obvious for those skilled in the art that other drawings can be obtained according to the contents of the embodiments of the present invention and the drawings without creative efforts.

FIG. 1 is a schematic sectional structural view of a DBD plasma vane type axial swirler of a combustion chamber head of an aircraft engine in accordance with the present invention;

FIG. 2 is a left side view of the structure of the present invention;

FIG. 3 is an isometric view of a cutaway structure of the present invention;

FIG. 4 is a schematic structural view of the inner ring ground electrode of the present invention;

FIG. 5 is a schematic structural diagram of the outer ring high voltage electrode according to the present invention;

FIG. 6(a) is a cross-sectional view of the cyclone body assembly of the present invention;

FIG. 6(b) is an isometric view of a cutaway view of a cyclone body member according to the present invention;

FIG. 7(a) is a schematic view of a swirler straight vane type structure of a swirler body according to the present invention;

FIG. 7(b) is a schematic view of a twisted blade type structure of a swirler of the swirler body according to the present invention;

FIG. 8 is a schematic view of the installation of the present invention in an aircraft engine combustion chamber;

in the figure:

1-inner ring ground electrode; 2-outer ring high voltage electrode; 3-a swirler body member; 4-a wire; 5-swirler inlet nozzle seat; 6-inner ring insulation barrier dielectric layer; 7-swirler vanes; 8-outer ring insulation barrier dielectric layer; 9-a swirler outlet seat; 10-assembling and fixing holes of the cyclone main body part; 11-DBD plasma discharge region; 12-DBD plasma vane axial swirler (invention); 13-an igniter; 14-a fuel nozzle; 15-combustion chamber outlet; 16-a flame tube; 17-a combustor casing; 18-main burning holes; 19-mixing holes.

Detailed Description

The present invention will be described in further detail with reference to the accompanying drawings and examples. It is to be understood that the specific embodiments described herein are merely illustrative of the invention and are not limiting of the invention. It should be further noted that, for the convenience of description, only some of the structures related to the present invention are shown in the drawings, not all of the structures.

As shown in fig. 1 to 3, the present embodiment is an aero-engine combustor head DBD plasma vane axial swirler, and for convenience of description, the flow direction of air is defined from bottom to top. The present embodiment is composed of 4 components, specifically, an inner ring grounding electrode 1, an outer ring high voltage electrode 2, a cyclone body component 3, and a lead 4. Wherein: the cyclone main body part 3 is of a revolving body structure and is integrally formed by DLP photocuring 3D ceramic printing; the cyclone main body part 3 comprises a cyclone main body part body, an outer ring insulation blocking dielectric layer 8 is arranged on the outer ring of the cyclone main body part body, an inner ring insulation blocking dielectric layer 6 is arranged on the inner ring of the cyclone main body part body, and cyclone blades 7 are arranged between the outer ring insulation blocking dielectric layer 8 and the inner ring insulation blocking dielectric layer 6; the outer ring high-voltage electrode 2 is adhered to the outer surface of the outer ring insulation barrier dielectric layer 8, and the inner ring grounding electrode 1 is adhered to the inner surface of the inner ring insulation barrier dielectric layer 6; one end of a lead 4 is welded on the outer side surface of the outer ring high-voltage electrode 2, and the other end of the lead 4 is connected with a high-voltage terminal of the plasma power supply output.

As shown in fig. 4 to 7, the cyclone main body member 3 is a rotary body structure, and is integrally formed by DLP photocuring 3D ceramic printing technology using a ceramic material having high temperature resistance and good insulation properties. The ceramic material is silicon carbide, silicon nitride or aluminum oxide. In this embodiment, adopt 99 aluminium oxide ceramic material, print integrated into one piece through DLP photocuring 3D pottery.

As shown in fig. 4 to 8, the cyclone main body 3 includes a cyclone main body, and for convenience of description, the cyclone main body is further subdivided into a cyclone inlet nozzle seat 5, an inner ring insulating barrier medium layer 6, cyclone vanes 7, an outer ring insulating barrier medium layer 8 and a cyclone outlet seat 9 according to different functions of each position of the cyclone main body 3; the outer ring of the main body of the cyclone main body is an outer ring insulation blocking medium layer 8, the bottom of the outer ring insulation blocking medium layer 8 is provided with a flange, the outer ring insulation blocking medium layer 8 extends upwards to form a cyclone outlet seat 9, a shoulder is arranged between the outer ring insulation blocking medium layer 8 and the cyclone outlet seat 9, the upper end of the cyclone outlet seat 9 is symmetrically provided with radial cyclone main body assembly fixing holes 10, and the cyclone main body assembly fixing holes 10 are used for fixing the device on a flame tube 16 at the head of an aircraft engine combustion chamber through screws; the inner ring of the main body of the cyclone main body part is an inner ring insulation blocking medium layer 6, and the inner ring insulation blocking medium layer 6 extends downwards to form a cyclone inlet nozzle seat 5; and a swirler vane 7 is arranged between the outer ring insulating barrier medium layer 8 and the inner ring insulating barrier medium layer 6. After the fuel nozzle 14 is sleeved with the swirler inlet nozzle seat 5, the outlet of the fuel nozzle is flush with the outlet plane of the swirler vane 7.

The inner ring grounding electrode 1 is a copper foil with a certain thickness, the outer side surface of the inner ring grounding electrode 1 is a sticky glue surface and is attached to the inner surface of the inner ring insulation barrier dielectric layer 6, and the inner side surface of the inner ring grounding electrode 1 is in contact with the metal fuel nozzle and then is grounded.

As shown in fig. 4, the inner ring ground electrode 1 has a cylindrical shape, is made of a metal foil having good conductivity and ductility, and has a thickness d₁0.01-0.5 mm, inner diameter D₁16-20 mm in length L₁6-20 mm. In this embodiment, the inner ring ground electrode 1 is made of copper foil with a thickness d₁0.4mm, inner diameter D₁Is 19.6mm and has a length L₁Is 10 mm.

The outer ring high-voltage electrode 2 is a copper foil with a certain thickness, the inner side surface of the outer ring high-voltage electrode 2 is a sticky rubber surface and is attached to the outer surface of the outer ring insulation barrier dielectric layer 8. The outer side surface of the outer ring high-voltage electrode 2 is sealed by silicon rubber, so that the metal outer side surface of the outer ring high-voltage electrode 2 and air are isolated, and the outer ring high-voltage electrode 2 is prevented from creepage and oxidation.

As shown in fig. 5, the outer ring high voltage electrode 2 is cylindrical, made of a metal foil with good conductivity and ductility, and has a thickness d₂0.01-0.5 mm, inner diameter D₂30-50 mm, length L ₂6-20 mm. In this embodiment, the inner ring ground electrode 1 is made of copper foil with a thickness d₂0.4mm, inner diameter D₂Is 40mm, length L₂Is 16 mm.

As shown in fig. 6(a), the cyclone inner diameter d of the cyclone main body 3_s16-24 mm, the outer diameter D of the cyclone_sIs 30 to 50 mm. In this embodiment, the swirler inner diameter d_sIs 20mm, the outer diameter D of the swirler_sIs 40 mm.

As shown in fig. 7(a) and 7(b), the swirler vanes of the swirler body 3 are shaped and structured as straight vanes or twisted vanes. In this embodiment, straight blades are used.

The number n of the swirler vanes of the swirler main body 3 is 5-12. In this embodiment, the number of the swirler vanes is 12.

The mounting angle beta of the swirler vanes 7 is 40-70 degrees. In this embodiment, the swirler vane setting angle β is 60 °.

As shown in fig. 6(a), the width Δ d of the discharge gap between the swirler vanes 7 of the swirler body 3 is 3 to 10 mm. In this embodiment, the discharge gap width Δ d is 5 mm.

As shown in fig. 6(a), the thickness δ of the inner ring insulating barrier dielectric layer 6₃0.5-2 mm, and the thickness delta of the outer ring insulation barrier dielectric layer 8₄Is 0.5 to 2 mm. In this embodiment, the thickness δ of the inner ring dielectric barrier layer ₃And outer ring insulation barrier dielectric layer thickness delta₄All are 1 mm.

The working medium of this embodiment is air.

The working principle of the invention is described as follows:

air from the compressor of an aircraft engine is divided into two main streams entering the interior of the combustor liner 16 of the combustor. The first air flow enters the main combustion area of the combustion chamber from the swirler at the head of the flame tube and the main combustion holes 18 and the cooling holes at the front half section of the flame tube respectively, and then is mixed and combusted with the fuel. When air entering from an axial swirler at the head part of the combustion chamber passes through an airflow channel at the position of an axial swirler vane 7, under the driving action of applying high-voltage alternating current to a plasma power supply, dielectric barrier discharge is formed between two electrodes of the device. Because the inner and outer ring insulation barrier dielectric layers have larger dielectric constants, the dielectric layers cannot be broken down, and the generation of independent strong arcs is blocked, and the gas between the electrodes can be broken down continuously to generate countless tiny discharge channels, so that a plasma discharge area 11 with uniform and stable discharge and large volume is formed in a space part (not including a swirler blade entity and an inner and outer ring insulation barrier dielectric layer entity) through which the gas flow passes between the two inner and outer ring electrodes. The air is broken down and ionized when passing through the plasma discharge area, and active particles with strong oxidizing property, such as high-concentration ozone, free oxygen atoms, excited molecular atoms, active groups and the like, are generated, so that the oxidizing capability of the air is greatly improved. In addition, the discharge process also increases the temperature of the air due to the temperature rise effect of the plasma. Meanwhile, through the swirler vanes with deflection angles, the air mixed with plasma and passing through the swirler vanes is converted into high-speed rotating jet flow by the axial swirler, and then enters the main combustion area at the head of the flame tube of the combustion chamber to form a backflow area, so that fuel atomization, oil-gas mixing and combustion are improved, the effect of combustion strengthening is achieved, and finally the combustion performances of the combustion stability range, outlet temperature field uniformity, combustion efficiency and the like of the combustion chamber of the aircraft engine are improved. The second air flow firstly enters two channels between the flame tube 16 and the combustion chamber shell 17, and then enters the mixing area through the mixing hole 19 and the cooling hole on the rear half section of the flame tube for reducing the temperature of the fuel gas and cooling the wall surface of the flame tube.

In the present invention, the terms "upper", "lower", "left", "right", and the like are used in an orientation or positional relationship based on those shown in the drawings only for convenience of description and simplification of operation, and do not indicate or imply that the referred device or element must have a specific orientation, be constructed in a specific orientation, and be operated, and thus, should not be construed as limiting the present invention.

The terms "comprises," "comprising," or any other variation thereof, are intended to cover a non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements does not include only those elements but may include other elements not expressly listed or inherent to such process, method, article, or apparatus.

Although embodiments of the present invention have been shown and described, it will be appreciated by those skilled in the art that changes, modifications, substitutions and alterations can be made in these embodiments without departing from the principles and spirit of the invention, the scope of which is defined in the appended claims and their equivalents.