阅读说明：本技术 一种组合型介质阻挡放电等离子体航空发动机燃烧室头部 (Combined dielectric barrier discharge plasma aircraft engine combustion chamber head ) 是由 赵兵兵 胡长淮 何立明 于锦禄 邓俊 雷健平 时广昊 赵志宇 曾昊 于 2021-09-14 设计创作，主要内容包括：一种组合型介质阻挡放电等离子体航空发动机燃烧室头部,二级文氏管套在一级文氏管的外圆周上。在二级文氏管的环形槽上固定有平板高压电极和圆环高压电极。平板高压电极和圆环高压电极通过连接导线连通并接入等离子体单高压电源的高压端；组合型介质阻挡放电等离子体航空发动机燃烧室头部的接地电极通过导线直接与地线相连中,使平板型DBD等离子体放电结构与圆管型DBD等离子体放电结构结合并能够进行两次电离,对进入燃烧室的气流以及燃油施加介质阻挡放电等离子体后,能够使燃烧更加充分,扩宽燃烧的熄火边界,且结构简单、制作和安装方便、通用性强,同样适用于其他具有类似结构的燃烧室。(A combined dielectric barrier discharge plasma aircraft engine combustion chamber head is characterized in that a secondary venturi tube is sleeved on the outer circumference of a primary venturi tube. A flat high-voltage electrode and a circular high-voltage electrode are fixed on the annular groove of the secondary venturi. The flat high-voltage electrode and the circular high-voltage electrode are communicated through a connecting wire and are connected to a high-voltage end of the plasma single high-voltage power supply; the grounding electrode of the combined dielectric barrier discharge plasma aircraft engine combustion chamber head is directly connected with the ground wire through the lead, so that the flat-plate DBD plasma discharge structure is combined with the circular-tube DBD plasma discharge structure and can be ionized twice, after dielectric barrier discharge plasma is applied to air flow and fuel oil entering the combustion chamber, the combustion can be more sufficient, the flameout boundary of the combustion can be widened, and the combined dielectric barrier discharge aircraft engine combustion chamber is simple in structure, convenient to manufacture and install and high in universality, and is also suitable for other combustion chambers with similar structures.)

1. A combined dielectric barrier discharge plasma aircraft engine combustion chamber head is characterized by comprising a secondary venturi, secondary radial swirler blades, a secondary radial swirler upper flat plate, a primary venturi, a primary inclined-cut hole swirler, a secondary radial swirler lower flat plate, a flat high-voltage electrode and a circular high-voltage electrode, wherein the secondary venturi is connected with the primary radial swirler upper flat plate through a connecting rod; the lower flat plates of the primary venturi tube, the primary inclined cut hole swirler and the secondary radial swirler are integrally processed and formed by metal materials; the secondary venturi tube, the secondary radial swirler vanes and the secondary radial swirler upper flat plate are bonded and fastened with the primary venturi tube, the primary inclined-cut hole swirler and the secondary radial swirler lower flat plate through high-temperature-resistant glue; the second-stage venturi tube is sleeved on the outer circumference of the first-stage venturi tube and is arranged on the upper surface of the lower flat plate of the second-stage radial swirler; a flat high-voltage electrode is fixed on the inner bottom surface of one side of the annular groove of the secondary venturi, and a circular high-voltage electrode is fixed on the inner surface of the groove wall of the inner side of the annular groove of the secondary venturi; the flat high-voltage electrode and the circular high-voltage electrode are communicated through a connecting wire and are connected to a high-voltage end of the plasma single high-voltage power supply; the grounding electrode of the head of the combustion chamber of the combined dielectric barrier discharge plasma aeroengine is directly connected with the ground wire through a wire; the primary oblique cut hole swirler is positioned in the middle of the primary venturi.

2. The combination dbd-pdp aircraft engine combustor head of claim 1, wherein the height h of the annular outer wall outside the outer circumference of the upper surface of the secondary swirler upper plate₁Is 5-10 mm, and the outer diameter D of the annular outer wall₁50-55 mm, thickness delta of annular outer wall₃0.5-2 mm; the height h from the top end of the secondary venturi to the upper surface of the upper flat plate of the secondary radial swirler₂6-12 mm, and the outer diameter D of the outlet of the secondary venturi tube₂30-35 mm in thickness delta₄0.2-0.8 mm; inner diameter D of annular groove of two-stage venturi₃20-25 mm, thickness delta₂Is 1 to 3 mm.

3. The composite dbd-pdp aircraft engine combustor head as claimed in claim 1, wherein the outer diameter D of the lower plate of the secondary radial swirler₄40-50 mm in thickness delta₅Is 1 &3 mm; height h from lower plate of first-stage venturi tube to second-stage radial swirler₃5-10 mm, the outer diameter D of the first-stage venturi tube₅18-23 mm, the radius R of the inner side arc of the first-stage venturi tube₁3-4 mm, inner diameter D of throat part of the first-stage venturi tube₆13-15 mm; inner diameter D of nozzle port₇15-20 mm in thickness delta₆Is 1-4 mm.

4. The composite dielectric barrier discharge plasma aircraft engine combustor head of claim 1, wherein the outer diameter D of said flat high voltage electrode₉30 to 50mm in thickness d₁0.1 to 3mm in width d₄10-40 mm.

5. The composite dielectric barrier discharge plasma aircraft engine combustor head of claim 1, wherein the outer diameter D of said annular high voltage electrode₈10 to 30mm, height d₂0.1 to 5mm in thickness d₃Is 0.1 to 3 mm.

6. The combination dielectric barrier discharge plasma aircraft engine combustor head of claim 1, wherein the gap between the annular grooves of the primary venturi and the secondary venturi is the discharge gap of said circular tube type DBD plasma discharge structure; the discharge gap is B₂1-10 mm; the upper flat plate of the secondary cyclone is used as an insulating medium layer of the flat DBD discharge structure, and the thickness delta of the upper flat plate of the secondary cyclone is₁Is 0.5 to 3 mm.

7. The combined type dielectric barrier discharge plasma aircraft engine combustion chamber head as claimed in claim 1, wherein the number of the inclined holes of the primary inclined hole type swirler is 3-20, and an included angle theta between the center line of each inclined hole and the tangent of the circular surface where the outlet is located₄Is 40 to 80 degrees.

8. A combined dielectric barrier discharge plasma aircraft engine combustor head as claimed in claim 1, whereinThe number of the second-stage radial swirler vanes is 3-20, the swirler vanes are straight vanes or curved vanes, the vane thickness delta is 0.4-1.2 mm, and the height B of the second-stage radial swirler vanes₂1-10 mm, blade installation angle theta₃40-80 degrees, and the width B of the channel at the outlet between the two blades₃Is 1 mm-20 mm.

9. The composite dielectric barrier discharge plasma aircraft engine combustor head of claim 1, wherein said primary venturi has an opening angle θ₁Is 30-80 degrees, and the opening angle theta of the secondary venturi tube₂Is 30-80 degrees.

Technical Field

The invention relates to a plasma intensified combustion technology in the field of aircraft engines, in particular to a combined dielectric barrier discharge plasma aircraft engine combustion chamber head.

Background

The combustion chamber is one of the core components of an aircraft engine and functions to convert chemical energy in fuel into heat energy through combustion release. High-temperature gas generated by combustion in the combustion chamber expands in the turbine and the spray pipe to do work, so that thrust is generated. Therefore, the performance of the combustion chamber will directly affect engine thrust and performance, operational stability, and reliability. In 1987, the united states started the "integrated high performance turbine engine technology project" (IHPTET) to ensure efficient and reliable operation of next generation aero-power systems under full operating envelope. The aeroengine combustion chamber mainly comprises a diffuser, a casing, a fuel nozzle, a swirler, a flame tube and the like. At present, the combustion chamber of the aero-engine makes many research progresses in the aspects of a novel fuel nozzle technology, a lean oil premixing and pre-evaporation technology, a rich oil-fast mixing-lean oil technology, a multi-stage cyclone head technology and the like. Much of the research in these technologies has focused on improvements and optimizations in existing combustor structures. However, these studies have still failed to meet the rapidly increasing aircraft engine performance demands. With the continuous improvement of the performance of the aero-engine, the structure of the aero-engine combustion chamber becomes more and more complex.

In order to meet the requirements of advanced aviation aircraft propulsion systems in the future, a plurality of technical bottlenecks which restrict efficient, stable, reliable and safe operation of an aircraft engine combustion chamber must be broken through. As a novel combustion chamber technology with great prospect, compared with other traditional intensified combustion technologies, the plasma intensified combustion technology has great advantages in the aspects of improving the combustion efficiency of the combustion chamber under all working conditions, widening the flameout boundary of the combustion chamber under extremely severe conditions, enhancing the capability of high-altitude re-ignition, improving the quality of the outlet temperature field of the combustion chamber and the like. In recent years, the plasma enhanced combustion technology has become a research hotspot in the field of domestic and foreign combustion. In 2009-2013, the North Atlantic convention group implemented a research plan for improving the performance of military aircrafts by using plasma. In 2009-2014, the U.S. department of defense implemented a research project of plasma-assisted combustion multidisciplinary university. A great deal of research is carried out on plasma combustion-supporting discharge abroad, and the research mainly focuses on the verification experiment research on the ignition combustion-supporting of the dielectric barrier discharge plasma. The research on plasma ignition combustion supporting of an aeroengine combustion chamber is late and still in the stages of principle verification and key technology attack. The research on the application of dielectric barrier discharge to plasma ignition combustion supporting is relatively slow. Due to the fact that dielectric barrier plasma discharge has the problems of difficult discharge under the condition of large gap, difficult high-voltage shielding, difficult long-time stable discharge and the like, and the structure and the working condition of the combustion chamber of the aero-engine are complex, the application of the technology to the combustion chamber of the aero-engine is limited, and documents related to mature application are few.

The patent with the publication number of CN102162644B of the institute of engineering thermal physics of Chinese academy of sciences discloses a structure of a dielectric barrier discharge plasma cyclone. According to the structure, a plurality of pairs of electrodes are arranged on two sides of the expansion section or the nozzle outlet section of the combustor in a staggered mode to form a plurality of pairs of surface dielectric barrier discharge plasma combustion-supporting exciters, so that gas is ionized, and the swirling effect of air is enhanced to assist combustion. However, most of the atomized fuel oil with the structure does not pass through an over-discharge area, and ignition is difficult; secondly, the insulating medium with the structure can not resist higher temperature, so that the excitation effect is limited.

Anokhin E M of the physical Technology college of Moscow, Russia, et al, Ignition of hydrocarbons by a nano controlled surface dielectric barrier discharge (Plasma Sources Science and Technology,2015,24(4):045014.) disclose a circular tube type dielectric barrier discharge Plasma exciter, one using a cylindrical central high voltage electrode and the other using a saw-tooth shaped central high voltage electrode, which constitutes a planar dielectric barrier discharge. Ignition of the mixed gas passing through the discharge region is achieved by arranging an insulating dielectric layer between the high voltage electrode and the ground electrode. The ignition mode is similar to the traditional spark ignition, mainly carries out ignition in a single-point area of a main combustion area, and has a limited excitation effect.

The university of air force engineering discloses a plate type aircraft engine combustion chamber dielectric barrier discharge plasma combustion-supporting exciter in the invention creation with the publication number of CN108412616A, and a plate electrode is arranged on the adjacent surface of an insulating plate to realize plasma discharge of incoming flow of an air inlet channel. In the discharge mode, the plate electrode is away from the combustion chamber, most of the incoming flow does not pass through the plasma discharge region, and ignition cannot be performed at the same time, so the action effect needs to be improved.

The dielectric barrier plasma discharge structure developed by the institute of engineering thermophysics of the academy of sciences of the Chinese academy of sciences, the physical technology college of the Russian department, and the university of air force engineering has a limited discharge action area, and in order to solve the problems that a plasma combustion-supporting exciter of dielectric barrier discharge is difficult to arrange and has a limited discharge area and improve the ignition combustion-supporting capability, the invention develops the combined DBD plasma aircraft engine combustion chamber head.

Disclosure of Invention

In order to overcome the defect that the combustion supporting of dielectric barrier discharge plasma applied to an aero-engine combustion chamber is difficult in the prior art, the invention provides a combined dielectric barrier discharge plasma aero-engine combustion chamber head.

The invention comprises a secondary venturi, secondary radial swirler vanes, a secondary radial swirler upper flat plate, a primary venturi, a primary inclined-cut hole swirler, a secondary radial swirler lower flat plate, a flat high-voltage electrode and a circular high-voltage electrode. The lower flat plates of the first-stage venturi tube, the first-stage oblique cutting hole swirler and the second-stage radial swirler are integrally processed and formed by metal materials. The secondary venturi, the secondary radial swirler vanes and the secondary radial swirler upper flat plate are bonded and fastened with the primary venturi, the primary inclined-cut hole swirler and the secondary radial swirler lower flat plate through high-temperature-resistant glue. The secondary venturi is sleeved on the outer circumference of the primary venturi and is arranged on the upper surface of the lower flat plate of the secondary radial swirler. A flat high-voltage electrode is fixed on the inner bottom surface of one side of the annular groove of the secondary venturi, and a circular high-voltage electrode is fixed on the inner surface of the groove wall of the inner side of the annular groove of the secondary venturi. The flat high-voltage electrode and the circular high-voltage electrode are communicated through a connecting wire and are connected to a high-voltage end of the plasma single high-voltage power supply; the grounding electrode of the combined dielectric barrier discharge plasma aircraft engine combustion chamber head is directly connected with the ground wire through a wire. The primary oblique cut hole swirler is positioned in the middle of the primary venturi.

The height h of the annular outer wall of the outer circumference of the upper surface of the upper flat plate of the secondary cyclone₁Is 5-10 mm, and the outer diameter D of the annular outer wall₁50-55 mm, thickness delta of annular outer wall₃Is 0.5 to 2 mm. The height h from the top end of the secondary venturi to the upper surface of the upper flat plate of the secondary radial swirler₂6-12 mm, and the outer diameter D of the outlet of the secondary venturi tube₂30-35 mm in thickness delta₄0.2-0.8 mm. Inner diameter D of annular groove of two-stage venturi₃20-25 mm, thickness delta₂Is 1 to 3 mm.

Outer diameter D of lower flat plate of secondary radial swirler₄40-50 mm in thickness delta₅Is 1 to 3 mm. Height h from lower plate of first-stage venturi tube to second-stage radial swirler₃5-10 mm, the outer diameter D of the first-stage venturi tube₅18-23 mm, the radius R of the inner side arc of the first-stage venturi tube₁3-4 mm, inner diameter D of throat part of the first-stage venturi tube₆Is 13-15 mm. Inner diameter D of nozzle port₇15-20 mm in thickness delta₆Is 1-4 mm.

The outer diameter D of the flat high-voltage electrode₉30 to 50mm in thickness d₁0.1 to 3mm in width d₄10-40 mm.

The outer diameter D of the ring high-voltage electrode₈10 to 30mm, height d₂0.1 to 5mm in thickness d₃Is 0.1 to 3 mm.

The width of the gap between the annular grooves of the first-stage venturi and the second-stage venturi is the discharge gap of the circular tube type DBD plasma discharge structure, and the discharge gap is B₂1-10 mm; the upper flat plate of the secondary cyclone is used as an insulating medium layer of the flat DBD discharge structure, and the thickness delta of the upper flat plate of the secondary cyclone is₁Is 0.5 to 3 mm.

The number of the inclined cutting holes of the primary inclined cutting hole type swirler is 3-20, and the included angle theta between the center line of each inclined cutting hole and the tangent line of the circular surface where the outlet is located₄Is 40 to 80 degrees.

The number of the second-stage radial swirler vanes is 3-20, the swirler vanes are straight vanes or curved vanes, the vane thickness delta is 0.4-1.2 mm, and the height B of the second-stage radial swirler vanes₂1-10 mm, blade installation angle theta₃40-80 degrees, and the width B of the channel at the outlet between the two blades₃Is 1 mm-20 mm.

Opening angle theta of the first-stage venturi₁Is 30-80 degrees, and the opening angle theta of the secondary venturi tube₂Is 30-80 degrees.

The invention takes a certain type of aircraft engine combustion chamber head as a prototype, and designs a combined DBD plasma aircraft engine combustion chamber head suitable for an aircraft engine. By adopting the 3D ceramic printing technology and the 3D metal printing technology, the dielectric barrier discharge plasma device is combined with the swirler of the combustion chamber of the aero-engine, so that the integrated design is realized, and according to the structural characteristics of the head of the combustion chamber of the aero-engine, the flat dielectric barrier plasma discharge structure and the circular tube type dielectric barrier plasma discharge structure are integrated to the head of a single combustion chamber to form the head of the combustion chamber of the DBD plasma aero-engine with a combined structure, so that the engineering application capability of the combustion-supporting exciter of the dielectric barrier discharge plasma is greatly improved. The DBD is an abbreviation of dielectric barrier discharge.

The invention comprises a secondary venturi, secondary radial swirler vanes, a secondary radial swirler upper flat plate, a primary venturi, a primary inclined-cut hole swirler, a secondary radial swirler lower flat plate, a flat plate high-voltage electrode, a ring high-voltage electrode and a connecting wire.

Wherein the secondary venturi, the secondary radial swirler vanes and the secondary radial swirler upper flat plate are integrally formed by ceramic insulating materials, as shown in fig. 4; the lower flat plates of the primary venturi tube, the primary inclined cut hole swirler and the secondary radial swirler are integrally formed by metal materials, and are shown in figure 5. The invention adopts a two-stage swirler which is formed by combining the blades of a first-stage oblique-cut hole type swirler and a second-stage radial swirler. Wherein the second-stage venturi tube is sleeved on the outer circumference of the first-stage venturi tube and is arranged on the upper surface of the lower flat plate of the second-stage radial swirler. The combustion-supporting excitation of the combined DBD plasma aircraft engine combustion chamber head comprises a flat plate type dielectric barrier plasma discharge structure and a coaxial circular tube type dielectric barrier plasma discharge structure. The flat plate high-voltage electrode of the flat plate type dielectric barrier plasma discharge structure is arranged on an annular groove on the upper surface of the upper flat plate of the secondary cyclone and is a circular metal foil electrode, and the upper flat plate of the secondary cyclone is an insulating medium layer. The annular outer wall of the outer circumference of the upper surface of the upper flat plate of the secondary cyclone can prevent the creepage phenomenon between the flat plate high-voltage electrode and the lower flat plate of the secondary radial cyclone. Under the drive of a plasma single high-voltage power supply, voltage is applied between a flat plate high-voltage electrode and metal of a lower flat plate of the secondary radial swirler, and a DBD plasma discharge area can be formed in an airflow channel of a blade of the secondary radial swirler between an upper flat plate of the secondary radial swirler and the lower flat plate of the secondary radial swirler; the annular high-voltage electrode of the coaxial circular tube type DBD plasma discharge structure is arranged on the inner surface of the groove wall on the inner side of the annular groove of the secondary venturi tube and is a cylindrical metal foil electrode, and the secondary venturi tube is used as an insulating medium layer. Under the drive of a plasma single high-voltage power supply, voltage is applied between the circular ring high-voltage electrode and the metal of the lower flat plate of the secondary radial swirler, and a DBD plasma discharge area is formed between the secondary venturi and the primary venturi. The flat high-voltage electrode and the circular high-voltage electrode are connected through a connecting wire and are jointly connected to the high-voltage end of the plasma single high-voltage power supply, and the flat plate below the secondary radial swirler is connected to the ground wire. A plasma single high-voltage power supply simultaneously drives a flat DBD plasma discharge structure and a coaxial circular tube type DBD plasma discharge structure to form a combined DBD plasma aircraft engine combustion chamber head. The upper direction refers to the direction from the lower end nozzle inlet of the central shaft in the head part of the combined DBD plasma aircraft engine combustion chamber to the outlet of the secondary venturi. The clockwise direction refers to the clockwise direction when viewed from the lower end nozzle inlet of the central shaft of the combined DBD plasma aircraft engine combustion chamber head to the upper end.

The invention provides a combined DBD plasma aircraft engine combustion chamber head based on a two-stage swirler of an aircraft engine combustion chamber head, wherein a flat DBD plasma discharge structure and a circular tube type DBD plasma discharge structure are combined and simultaneously discharged. In an engine, air from a compressor is diffused through a diffuser to be decelerated, and a part of the decelerated air enters a flame tube through a swirler arranged at the head part of a combustion chamber. The two-stage swirler not only organizes the flow field to stabilize the flame, but also promotes mixing of the fuel with the air, and in addition, is located relatively close to the main combustion zone. Therefore, DBD plasma discharge is carried out at the position, on one hand, the flat-plate-type DBD plasma discharge structure carries out primary ionization on air flowing into the combustion chamber, and a DBD plasma discharge area of an airflow channel in a blade of the secondary radial swirler between an upper flat plate of the secondary radial swirler and a lower flat plate of the secondary radial swirler is beneficial to forming plasma discharge with larger space volume, so that the defects of small DBD plasma discharge gap and small plasma area volume at the position of the primary venturi tube can be overcome; on the other hand, the circular tube type DBD plasma discharge structure can carry out secondary ionization on air and can also carry out ionization on fuel oil sprayed by the nozzle. Meanwhile, the DBD plasma discharge area at the position of the venturi is close to the flame front of the combustion chamber. The dielectric barrier discharge plasma generated at the position enters the combustion chamber more quickly to participate in the combustion process. The working medium of the combined DBD plasma aircraft engine combustion chamber head is used as gas, a flat plate high-voltage electrode of a flat plate type DBD plasma discharge structure is of a ring structure, a secondary venturi, secondary radial swirler vanes and a secondary radial swirler upper flat plate are integrally machined and formed by ceramic insulating materials, a high-voltage alternating current power supply is applied to two ends of the electrode, a DBD plasma discharge area is formed in an airflow channel of the secondary radial swirler vanes between the secondary radial swirler upper flat plate and the secondary radial swirler lower flat plate, and plasma discharge is conducted on air flowing through the DBD plasma discharge area. By utilizing the characteristics of the first-stage venturi tube and according to the outlet shape of the burner head swirler, the annular high-voltage electrode of the coaxial circular tube type DBD plasma discharge structure adopts a cylindrical structure, DBD discharge is carried out at the outlet position of the radial swirler, the action effect in plasma is further improved, and fuel in the area can be ionized. The secondary venturi, the secondary radial swirler vanes and the secondary radial swirler upper flat plate at the position are integrally machined and formed, so that the uniformity of the insulating medium layer can be effectively ensured to avoid the breakdown phenomenon in the discharging process and maintain stable DBD plasma discharging. Therefore, under the action of the head part of the combustion chamber of the combined DBD plasma aircraft engine, air and fuel oil enter the main combustion area of the combustion chamber, the chemical reaction of combustion can be promoted, and the chemical reaction rate of combustion can be improved. The invention carries out the reconstruction of the prior aircraft engine combustion chamber head, does not change the prior combustion chamber head structure, has high application value, and simultaneously adopts two forms of DBD combustion-supporting excitation structures, thereby having better excitation effect.

Fig. 12 is a schematic view of combustion efficiency and combustion efficiency increment in a combustion chamber in a normal state and in a plasma combustion supporting state, and as can be seen from a combustion efficiency 24 of the combustion chamber in the normal state, a combustion efficiency 23 of the combustion chamber in the plasma combustion supporting state and a combustion efficiency increment 25 of the combustion chamber in the plasma combustion supporting state, after the invention is installed on a head part of a combustion chamber of an aero-engine, due to the action of DBD plasma, a combustion process is more sufficient, and when an after-gas coefficient is 0.8, the combustion efficiency increment can be improved to 3.23%; fig. 13 is a flameout boundary curve of the combustion chamber in the normal state and in the plasma combustion supporting state, and is composed of a flameout boundary curve 26 of the combustion chamber in the normal state and a flameout boundary curve 27 of the combustion chamber in the plasma combustion supporting state. The curve reflects the residual gas coefficient at extinction at different combustor inlet velocities. Experimental results show that the flameout boundary of the combustion chamber is widened to 7.1%.

The invention is combined with the head part of the combustion chamber of the aero-engine, the structural size of the invention is completely matched with the combustion chamber, the size and the structure of the original combustion chamber are not changed, and the invention can be directly installed in the combustion chamber for use. According to the record of the monograph 'gas turbine engine combustion' on page P201, two rotational flows 'at the head part of the combustion chamber can enhance the shearing action between fuel oil and atomized air and is favorable for fuel oil atomization and oil-gas mixing', and the oil-gas condition at the position is better. Meanwhile, at the head of the combustion chamber, the oil-gas condition is favorable for the combustion supporting of the dielectric barrier discharge plasma; and the head part of the combustion chamber utilizes the airflow self-sufficiency of the cyclone, and does not need external air supply. The invention combines the flat plate type dielectric barrier discharge structure and the coaxial circular tube type dielectric barrier discharge structure, can carry out ionization twice, can ensure that the combustion is more sufficient after dielectric barrier discharge plasma is applied to the airflow and the fuel oil entering the combustion chamber, widens the flameout boundary of the combustion, has simple structure, convenient manufacture and installation and strong universality, and is also suitable for other combustion chambers with similar structures.

Drawings

FIG. 1 is a schematic diagram of FIG. 1 showing an axial swirling structure of plasma along a surface dielectric barrier discharge developed by institute of engineering thermal physics of the Chinese academy of sciences;

FIG. 2 is a circular tube type dielectric barrier discharge plasma structure along the surface developed by the physical technology academy of Temminck, Russia;

FIG. 3 is a combustion-supporting exciting structure of a guide vane type DBD plasma of an aircraft engine combustion chamber, which is developed by the university of air force engineering;

FIG. 4 is a schematic diagram of the construction of a secondary venturi, secondary radial swirler vanes and secondary radial swirler upper plate of ceramic material, wherein FIG. 4a is a front view; FIG. 4b is a top view of FIG. 4 a;

FIG. 5 is a cross-sectional view A-A of FIG. 4 a;

FIG. 6 is a cross-sectional view B-B of FIG. 4 a;

FIG. 7 is a schematic structural diagram of a lower flat plate of a primary venturi, a primary chamfered hole swirler and a secondary radial swirler made of metal, wherein FIG. 7a is a front view; FIG. 7b is a top view of FIG. 7 a;

FIG. 8 is a schematic diagram of the connection of a flat plate high voltage electrode and a circular ring high voltage electrode;

FIG. 9 is a schematic structural diagram of a flat high voltage electrode;

FIG. 10 is a schematic structural diagram of a ring high voltage electrode;

FIG. 11 is a schematic view of the structure of a combined dielectric barrier discharge plasma aircraft engine combustion head;

FIG. 12 is a schematic view of the structure of a chamfered orifice swirler inlet passage; wherein 12a is a front view of the inlet passage of the chamfered hole type swirler; 12b is a top view of the inlet channel of the chamfered hole type swirler;

FIG. 13 is the installed position of the present invention in the annular combustor;

FIG. 14 is an installed position of the present invention in a reverse flow type combustor;

FIG. 15 is a schematic view of combustion efficiency and increment in a normal state of a combustion chamber and a state of carrying out plasma combustion supporting;

fig. 16 is a flameout boundary curve in a normal state of the combustion chamber and a state of carrying out plasma combustion supporting.

In the figure: 1. a power source; 2. a high-voltage interface; 3. a ground wire interface; 4. a high voltage electrode; 5. a ground electrode; 6. an insulating dielectric layer; 7. a secondary venturi; 8. secondary radial swirler vanes; 9. an upper flat plate of the secondary radial swirler; 10. the direction of air intake; 11. a flat plate high voltage electrode; 12. a circular ring high voltage electrode; 13. connecting a lead; 14. a primary venturi; 15. a first-stage oblique cutting hole swirler; 16. a lower flat plate of the secondary radial swirler; 17. a nozzle interface; 18. an igniter; 19. the mounting position of the head part of the annular combustion chamber; 20. an annular combustion chamber inlet expansion section; 21. a mounting position of the head of the reflux type combustion chamber; 22. an inlet expansion section of the backflow type combustion chamber; 23. the efficiency of the combustion chamber is realized in a plasma combustion-supporting state; 24. combustion efficiency of the combustion chamber in a normal state; 25. combustion efficiency increment of the combustion chamber; 26. a flameout boundary of the combustion chamber in a normal state; 27. and implementing flameout boundary of the combustion chamber in the plasma combustion supporting state.

Detailed Description

The embodiment is a combined dielectric barrier discharge plasma aircraft engine combustion chamber head. For convenience of description, the nozzle port 17 of the metal ground electrode at the lower end of the present invention is defined as an inlet, and the orifice of the secondary venturi 7 of the ceramic insulator at the upper end is defined as an outlet.

The embodiment comprises a secondary venturi 7, secondary radial swirler vanes 8, a secondary radial swirler upper flat plate 9, a primary venturi 14, a primary inclined hole swirler 15, a secondary radial swirler lower flat plate 16, a nozzle interface 17, a flat high-voltage electrode 11, a circular high-voltage electrode 12 and a connecting lead 13. The secondary venturi, the secondary radial swirler vanes and the secondary radial swirler upper flat plate are integrally processed and formed by ceramic insulating materials. The lower flat plates of the first-stage venturi tube, the first-stage oblique cutting hole swirler and the second-stage radial swirler are integrally processed and formed by metal materials. Wherein:

the height h of the annular outer wall of the outer circumference of the upper surface of the upper flat plate 9 of the secondary cyclone₁Is 5-10 mm, and the outer diameter D of the annular outer wall₁50-55 mm, thickness delta of annular outer wall₃Is 0.5 to 2 mm. The height h from the top end of the secondary venturi 7 to the upper surface of the upper flat plate 9 of the secondary radial swirler₂6-12 mm, and the outer diameter D of the outlet of the secondary venturi tube₂30-35 mm in thickness delta₄0.2-0.8 mm. Inner diameter D of annular groove of secondary venturi 7₃20-25 mm, thickness delta₂Is 1 to 3 mm. In the embodiment, the height h of the annular outer wall of the outer circumference of the upper surface of the upper flat plate 9 of the secondary cyclone₁Is 8mm, and the outer diameter D of the annular outer wall₁52mm, thickness delta of the annular outer wall₃Is 1 mm. The height h from the top end of the secondary venturi 7 to the upper surface of the upper flat plate 9 of the secondary radial swirler₂8.5mm, the outer diameter D of the outlet of the secondary venturi tube₂33mm, outlet thickness delta₄Is 0.5 mm. Inner diameter D of annular groove of secondary venturi 7₃Is 23mm, thickness delta₂Is 1.5 mm.

The outer diameter D of the lower flat plate 16 of the secondary radial swirler₄40-50 mm in thickness delta₅Is 1 to 3 mm. Height h of upper surface of lower plate 16 from first-stage venturi 14 to second-stage radial swirler₃5-10 mm, first-stage venturi 14Outer diameter D₅18-23 mm, the inner side arc radius R of the first-stage venturi tube 14₁3-4 mm, inner diameter D of throat part of the first-stage venturi tube₆Is 13-15 mm. Inner diameter D of nozzle port 17₇15-20 mm in thickness delta₆Is 1-4 mm. In this embodiment, the outer diameter D of the lower plate 16 of the secondary radial swirler₄Is 44mm and has a thickness delta₅Is 1.5 mm. Height h of upper surface of lower plate 16 from first-stage venturi 14 to second-stage radial swirler₃5mm, the outer diameter D of the primary venturi 14₅20mm, the radius R of the inner arc of the primary venturi tube 14₁3.3mm, inner diameter D of throat part of the first-stage venturi tube 14₆Is 14 mm. Inner diameter D of nozzle port 17₇Is 16mm and has a thickness delta₆Is 2.5 mm.

The flat high-voltage electrode 11 is a metal foil with good conductivity and ductility. The flat high-voltage electrode 11 of the flat DBD plasma discharge structure is annular, and the outer diameter D of the flat high-voltage electrode₉30 to 50mm in thickness d₁0.1 to 3mm in width d₄10-40 mm. In this embodiment, the flat high voltage electrode 11 of the flat DBD plasma discharge structure is a circular ring with an outer diameter D₉Is 42mm and has a thickness d₁Is 0.5mm, and has a width d₄Is 16 mm.

The material of the annular high-voltage electrode 12 is a metal foil with good conductivity and ductility. The annular high-voltage electrode 12 of the coaxial circular tube type DBD plasma discharge structure is in a cylindrical shape, and the outer diameter D of the annular high-voltage electrode₈10 to 30mm, height d₂0.1 to 5mm in thickness d₃Is 0.1 to 3 mm. In this embodiment, the outer diameter D of the annular high voltage electrode 12 of the coaxial circular tube type DBD plasma discharge structure₈Is 26mm, height d₂Is 3.4mm, thickness d₃Is 0.5 mm.

The secondary venturi 7 is fitted over the outer circumference of the primary venturi 14 and is seated on the upper surface of the lower flat plate 16 of the secondary radial swirler. A flat high-voltage electrode 11 is fixed on the inner bottom surface of one side of the annular groove of the secondary venturi, and a circular high-voltage electrode 12 is fixed on the inner surface of the inner side groove wall of the annular groove of the secondary venturi.

The flat high-voltage electrode 11 and the circular high-voltage electrode 12 are communicated through a connecting wire 13 and are connected to a high-voltage end of the plasma single high-voltage power supply; the grounding electrode 5 of the combined dielectric barrier discharge plasma aircraft engine combustion chamber head is directly connected with the ground wire through a wire.

The discharge gap of the flat-plate DBD plasma discharge structure is B₂1 to 10mm, and a thickness delta thereof₁0.5-3 mm; the gap width between the annular grooves of the first-stage venturi tube 14 and the second-stage venturi tube 7 is the discharge gap B of the circular tube type DBD plasma discharge structure₁Is 1-10 mm. In this embodiment, the discharge gap of the flat-plate DBD plasma discharge structure is B₂Is 4.3mm and has a thickness delta₁Is 1.5 mm; the gap width between the annular grooves of the first-stage venturi tube 14 and the second-stage venturi tube 7 is the discharge gap B of the circular tube type DBD plasma discharge structure₁Is 1.5 mm.

The number of the inclined cutting holes of the primary inclined cutting hole type swirler 15 is 3-20, and an included angle theta between each inclined cutting hole and the direction of the symmetry axis in the figure 7₅The included angle between the central line of the oblique cutting hole and the intersection point tangent line of the circle where the outlet is positioned is 40-80 degrees. In this embodiment, the number of the inclined holes of the first-stage inclined hole type swirler 15 is 8, and an included angle θ between the inclined holes and the direction of the symmetry axis in fig. 7₅The included angle between the central line of the oblique cutting hole and the tangent line of the intersection point of the circle where the outlet is positioned is 45 degrees.

The number of the second-stage radial swirler vanes 8 is 3-20, the swirler vanes can be straight vanes or curved vanes, the vane thickness delta is 0.4-1.2 mm, and the height B of the second-stage radial swirler vanes 8₂1-10 mm, blade installation angle theta₃40-80 degrees, and the width B of the channel at the outlet between the two blades₃Is 1 mm-20 mm. In this embodiment, the number of the second-stage radial swirler vanes 8 is 8, the swirler vanes are curved vanes, the vane thickness δ is 1mm, and the height B of the second-stage radial swirler vanes 8₂Is 4.3mm, blade mounting angle theta₃Is 70 degrees, and the width B of the channel at the outlet between the two blades₃3.745 mm.

Opening angle theta of the first-stage venturi₁Is 30-80 degrees, and the opening angle theta of the secondary venturi tube₂Is 30-80 degrees. In this embodiment, the opening angle θ of the first-stage venturi tube₁Is 45 degrees, and the opening angle theta of the secondary venturi tube₂Is 50 deg..

The secondary venturi, the secondary radial swirler vanes and the secondary radial swirler upper flat plate are made of ceramic insulating materials with high temperature resistance and good insulativity. In this embodiment, the ceramic insulating material is made of 99 alumina ceramic, has a dielectric constant of 11, and is manufactured by integrally processing a photo-cured 3D ceramic by printing, machining, or die casting.

In this embodiment, the secondary venturi, the secondary radial swirler vanes, and the secondary radial swirler upper flat plate are integrally formed by a ceramic insulating material and integrally formed by a metal material, and the primary venturi, the primary inclined hole swirler, and the secondary radial swirler lower flat plate are bonded and fastened by a high temperature resistant adhesive. The test is carried out in the embodiment, and the test device is arranged in a combustion chamber of an axial flow aero-engine of a certain type. Air entering the head of the combustion chamber, wherein a part of air flow enters a bevel hole of the primary bevel hole type swirler to form rotational flow air flow, fuel oil sprayed out from the nozzle is further atomized when passing through the primary venturi tube and the secondary venturi tube, is mixed with the atomized fuel oil, and then enters a main combustion area of the combustion chamber to participate in the combustion process; the other part of the airflow enters a secondary radial swirler, passes through a plasma discharge area at the position of an air inlet channel of the secondary radial swirler, is subjected to primary ionization, then enters a plasma discharge area between a round pipe part of a secondary venturi 7 and a primary venturi 14, is subjected to secondary plasma ionization, is mixed with fuel oil which is also subjected to ionization, and then enters a main combustion area of a combustion chamber to participate in the combustion process.

Experiments show that after the implementation is applied to a laboratory platform of a combustion chamber, the oil-gas mixture is more fully combusted in a main combustion area, the combustion efficiency is improved by 3.23%, and the flameout boundary widening of the combustion chamber reaches 7.1%.