CN116878031A - Trapped vortex combustion chamber and aeroengine - Google Patents

Abstract

The application discloses a trapped vortex combustion chamber and an aeroengine, wherein the trapped vortex combustion chamber comprises a first airflow channel, and is positioned between an outer casing of an outer duct and an inner casing of the outer duct; the first combustion chamber is positioned at one side of the outer duct inner casing, which is close to the outer casing of the outer duct, and is communicated with the first air flow channel, and the outer duct air flow forms a standing vortex in the first combustion chamber when passing through the first air flow channel; the second airflow channel is positioned between the inner duct outer casing and the inner duct inner casing; the second combustion chamber is positioned at one side of the inner duct outer casing, which is close to the inner duct inner casing, and is communicated with the second airflow channel, and the inner duct airflow forms a standing vortex in the second combustion chamber when passing through the second airflow channel; the first combustion chamber and the second combustion chamber are both connected with a fuel nozzle and a corresponding ignition electric nozzle. The application has the effects of obtaining larger thrust and not increasing the outer axial dimension.

Description

Trapped vortex combustion chamber and aeroengine

Technical Field

The application relates to the technical field of turbine engines, in particular to a trapped vortex combustion chamber and an aeroengine.

Background

The main turbofan engine currently mainly comprises the following three types: the thrust-increasing turbofan engine for high-speed fixed wing aircrafts (such as military fighters), the middle-large bypass ratio turbofan engine for subsonic fixed wings and the turbofan engine with a conventional turbine interstage combustion chamber can be divided into a non-thrust-increasing state and a thrust-increasing state, wherein the thrust generated by the thrust-increasing state of the engine is increased by about 60 percent compared with the non-thrust-increasing state, but the fuel consumption is increased by 150 to 200 percent, so that the thrust-increasing state is generally used briefly only in special conditions such as air fights and the like, and the service time is extremely limited; the middle-large bypass ratio turbofan engine has the remarkable advantages that the unit thrust oil consumption is low, but the bypass ratio is generally larger than 3, the unit thrust is smaller, an afterburner cannot be arranged, and the flight Mach number cannot exceed 1.0; compared with a turbofan engine with an afterburner, the turbofan engine with the conventional turbine interstage combustion chamber has the advantages that only the inner duct is heated, and the outer duct gas is not heated, so that the unit thrust lifting amplitude of the engine is limited, and is generally about 20% -30%.

Therefore, the existing turbofan engine has limited thrust, and if the forced turbofan engine is adopted, the axial dimension of the engine is oversized because of adding the forced-air combustion chamber, which is unfavorable for the design and layout of the aircraft.

Disclosure of Invention

The application provides a trapped vortex combustion chamber and an aeroengine, which are used for solving the technical problem that the axial size of a high-thrust turbine engine is overlarge.

According to one aspect of the present application, there is provided a trapped vortex combustor comprising a first airflow passage between an outer casing of an outer duct and an inner casing of the outer duct; the first combustion chamber is positioned at one side of the outer duct inner casing, which is close to the outer casing of the outer duct, and is communicated with the first air flow channel, and the outer duct air flow forms a standing vortex in the first combustion chamber when passing through the first air flow channel; the second airflow channel is positioned between the inner duct outer casing and the inner duct inner casing; the second combustion chamber is positioned at one side of the inner duct outer casing, which is close to the inner duct inner casing, and is communicated with the second airflow channel, and the inner duct airflow forms a standing vortex in the second combustion chamber when passing through the second airflow channel; the first combustion chamber and the second combustion chamber are both connected with a fuel nozzle and a corresponding ignition electric nozzle.

By adopting the technical scheme, the external air flow forms the standing vortex in the first combustion chamber, the internal air flow forms the standing vortex in the second combustion chamber, the air flow in the standing vortex is reflux and is positioned in the concave cavity, the influence of the main flow is less, and the performance is stable, so that flame formed after the fuel injection and ignition in the first combustion chamber and the second combustion chamber is stable; conventional turbofan engines employ a brayton cycle, and increasing the pre-turbine temperature is an effective means of increasing the thrust-to-weight ratio of the engine, however limited by technology and cost, and limited in the magnitude of the pre-turbine temperature increase. The conventional turbofan engine is often heated at a single point in the main combustion chamber, and the inner duct and the outer duct are additionally provided with the first combustion chamber and the second combustion chamber which are matched with the main combustion chamber for heating, so that the temperature in front of the turbine can be effectively increased, and the thrust of the engine is further improved; compared with the traditional forced turbofan engine, the forced-air turbofan engine has the advantages that the forced-air combustion chamber behind the turbine is not required to be arranged, and the concave cavity form is changed, so that the axial length of the engine is obviously shortened, and the scheme can obtain larger thrust without increasing additional axial dimension.

Compared with the conventional turbine interstage combustion chamber which is only heated in the engine connotation, the first combustion chamber and the second combustion chamber provided by the application can realize simultaneous heating of the inner and outer connotations by simultaneous oil injection and ignition, and can also heat one of the connotation and the outer connotation by only oil injection and telephone in one of the chambers, thereby increasing the exhaust speed of the engine, improving the thrust, reducing the geometrical regulating mechanism of the engine connotation and reducing the complexity and weight of the engine.

Optionally, an external duct flow stabilizer is arranged in the first air flow channel, the external duct air flow enters the first combustion chamber after passing through the external duct flow stabilizer, an internal duct flow stabilizer is arranged in the second air flow channel, and the internal duct air flow enters the second combustion chamber after passing through the internal duct flow stabilizer.

Through adopting above-mentioned technical scheme, interior channel flow stabilizer and outer channel flow stabilizer can play the effect of reducing the velocity of flow, and interior channel flow stabilizer and second burning chamber, outer channel flow stabilizer and first burning chamber form the backward flow district respectively, make first burning chamber and second burning intracavity form the trapped vortex better to promote combustion efficiency.

Optionally, the first combustion chamber and the second combustion chamber are both annular surrounding the inner casing of the inner duct.

Through adopting above-mentioned technical scheme, annular first combustion chamber and the second combustion chamber that sets up can carry out even heating to the air current of all flowing through interior ducts and outer duct, and then guarantee that the velocity of flow of each direction air current is unanimous on the circumference to obtain stable thrust.

Optionally, the first combustion chamber is closely attached to the second combustion chamber, and the second combustion chamber ignition power nozzle of the first combustion chamber is commonly connected with an ignition cable.

By adopting the technical scheme, the internal structure of the engine can be more compact, and the wiring length is shortened.

According to another aspect of the present application, there is further provided an aeroengine, including an outer casing of an outer duct, an inner casing of an inner duct, and the trapped vortex combustion chamber described above, wherein an outer duct channel through which an outer duct airflow passes is formed between the outer casing of the outer duct and the inner casing of the outer duct, the outer duct channel is communicated with the first airflow channel, an inner duct channel through which an inner duct airflow passes is formed between the inner casing of the outer duct and the inner casing of the inner duct, and the inner duct channel is communicated with the second airflow channel.

Optionally, the device further comprises an air inlet guide, a fan, a high-pressure compressor, a combustion chamber, a high-pressure turbine working blade and a low-pressure turbine which are sequentially arranged from an air inlet end to an air outlet end, and the trapped vortex combustion chamber is positioned between the high-pressure turbine working blade and the low-pressure turbine.

Through adopting above-mentioned technical scheme, this scheme first combustion chamber and second combustion chamber are located behind the high-pressure turbine, compare with afterburner fan engine setting afterburner behind the turbine, and the heating ambient pressure of this scheme is high, and the speed is lower, and this is favorable to promoting combustion efficiency and circulation efficiency, and the oil consumption is lower, is favorable to increasing the dead time.

Optionally, a high-pressure turbine guide vane is arranged between the combustion chamber and the high-pressure turbine working vane, a third combustion chamber is formed on the high-pressure turbine guide vane, a standing vortex is formed in the third combustion chamber when air flow in the combustion chamber passes through the high-pressure turbine guide vane, and a fuel nozzle and an ignition electric nozzle are arranged in the third combustion chamber.

Through adopting above-mentioned technical scheme, on the one hand, the high-pressure turbine stator is more played the effect of water conservancy diversion, and on the other hand is through setting up the third combustion chamber on the high-pressure turbine stator, and the heating ambient pressure in third combustion chamber is high, and air velocity is low, is favorable to burning, through the mode of newly-increased combustion point, can effectively promote the temperature before the engine whirlpool, and then promotes engine thrust.

Optionally, an adjustable tail nozzle is arranged at the air outlet end of the outer casing of the outer duct.

By adopting the technical scheme, the area of the tail nozzle channel can be adjusted when the tail nozzle is adjusted, so that the distribution of the expansion ratio of air flow in the turbine and the tail nozzle can be changed, and the control of the working state of the whole engine is realized.

Optionally, the high-pressure turbine guide vane is provided with a plurality of fuel nozzles on all the high-pressure turbine guide vanes at equal intervals around the circumference of the inner casing of the inner channel.

By adopting the technical scheme, setting a plurality of high-pressure turbine guide vanes means setting the same number of combustion points, the combustion points are distributed along the circumferential direction, so that heating is more uniform, and the purpose of synchronous oil injection of the fuel nozzles on all the high-pressure turbine guide vanes is also to promote the consistency of combustion of a plurality of combustion points so as to realize uniform heating.

Optionally, the fuel nozzle of the high-pressure turbine guide vane is a hollow cone nozzle, and the included angle between the hollow cone nozzle and the inlet of the third combustion chamber is 60 degrees.

Through adopting above-mentioned technical scheme, from after the fuel of blowout in the hollow cone nozzle gets into the third combustion chamber, form the vortex in the third combustion chamber under the effect of third combustion chamber entry inflow, can mix with the air current better, promote combustion efficiency.

In summary, the present application includes at least one of the following beneficial technical effects:

1. by newly adding heating points in the inner duct and the outer duct, the traditional single-point heating is changed into a multi-point heating mode, so that the temperature before the turbine can be effectively increased, and the thrust of the engine is further improved;

2. the afterburner behind the turbine is eliminated, and the concave cavity is changed, so that the axial length of the engine is obviously shortened.

In addition to the objects, features and advantages described above, the present application has other objects, features and advantages. The present application will be described in further detail with reference to the drawings.

Drawings

The accompanying drawings, which are included to provide a further understanding of the application and are incorporated in and constitute a part of this specification, illustrate embodiments of the application and together with the description serve to explain the application. In the drawings:

FIG. 1 is a schematic diagram of a forced turbofan engine;

FIG. 2 is a schematic structural view of a trapped vortex combustor in accordance with a preferred embodiment of the present application;

FIG. 3 is a schematic view of the airflow direction of the trapped vortex of the present application;

FIG. 4 is a schematic structural view of an aircraft engine according to a preferred embodiment of the application;

FIG. 5 illustrates a side view of a high pressure turbine vane of a preferred embodiment of the present application.

Legend description:

1. an outer duct outer casing; 2. an outer duct flow stabilizer; 3. an ignition cable; 4. an outer duct inner casing; 5. a second combustion chamber fuel nozzle; 6. a first combustion chamber fuel nozzle; 7. an inner duct outer casing; 8. an inner casing of the inner duct; 9. an inner bypass flow stabilizer; 10. a nozzle is started; 11. an air intake guide; 12. a fan; 13. a high pressure compressor; 14. a main combustion chamber; 15. high pressure turbine vanes; 16. high pressure turbine rotor blades; 17. a trapped vortex combustion chamber; 18. a low pressure turbine; 19. an adjustable tail nozzle; 20. afterburner; 21. a first combustion chamber; 22. a second combustion chamber; 23. and a third combustion chamber.

Detailed Description

Embodiments of the application are described in detail below with reference to the attached drawing figures, but the application can be practiced in a number of different ways, as defined and covered below.

The application is described in further detail below with reference to fig. 1-4.

The existing turbofan aero-engine mainly comprises the following three types: a forced turbofan engine, a medium-large bypass ratio turbofan engine, and a turbofan engine with a conventional turbine interstage combustion chamber.

Referring to FIG. 1, a conventional forced induction turbofan engine is typically used in a high-speed fixed wing aircraft (e.g., a military fighter aircraft) and is divided into inner and outer ducts, consisting of fans 12, compressors, combustors, turbines, afterburners 20, nozzles, and the like.

The working principle of the forced turbofan engine is as follows: after entering the engine, the air is divided into an inner duct air flow and an outer duct air flow. The content gas flows through the content high-pressure compressor 13 to be compressed, enters the content combustion chamber to be combusted, becomes high-temperature and high-pressure gas, drives the content high-pressure turbine 18 and the low-pressure turbine 18 (part of internal energy of the gas is converted into mechanical energy of the turbine to drive the content high-pressure compressor 13 and the fan 12 to be in and out of the turbine), and then enters the spray pipe to be mixed with the content gas flow. The outer duct air flows through the fan 12 to be compressed by the outer duct, enters the spray pipe, is mixed with the inner duct air flow, and then is sprayed out of the engine at a high speed through the spray nozzle, and the exhaust gas mixed by the inner duct and the outer duct generates reaction force, namely the thrust of the engine, on the engine. When the aircraft needs the engine to generate larger thrust, the fuel can be supplemented and injected into the afterburner 20 behind the turbine to further burn because the fuel gas behind the turbine and the external duct airflow are not burnt out, so that the internal energy of the fuel gas is further increased, and the thrust of the engine is increased.

The working state of the forced turbofan engine can be divided into a non-forced state and a forced state. The thrust force generated by the stress state of the engine is increased by about 60% compared with the stress-free state, but the fuel consumption rate is increased by 150% -200%. Therefore, the stress state is generally used for a short time only in special cases such as air combat, and the use time is extremely limited. The data for a typical boost turbofan engine abroad is given in the following table.

On the other hand, to ensure adequate combustion of the high velocity gas within afterburner 20, afterburner 20 is typically longer, accounting for more than 30% of the overall engine length, resulting in a longer overall engine length. In summary, the conventional structure of the afterburner engine has an excessive fuel consumption in the afterburner 20 and an excessive length in the afterburner state.

The middle-large bypass ratio turbofan engine is often used for subsonic fixed wing aircrafts, and the aircrafts mainly emphasize the endurance time, such as civil airliners, public service machines, investigation unmanned aerial vehicles and the like, and do not emphasize the maneuverability and the supersonic speed, so that the middle-large bypass ratio turbofan engine is mostly adopted. The engine structure is shown in the following figures. The engine is also divided into an inner duct and an outer duct, and consists of a fan 12, a gas compressor, a combustion chamber, a turbine, an afterburner 20, a nozzle and the like. The working principle is similar to that of a forced turbofan engine, and the main difference is that the power device fan 12 has larger diameter, the expansion ratio of a tail nozzle is low, the forced-air combustor 20 cannot be installed, the exhaust speed is low, and the unit thrust is small. The remarkable advantage of such engines is low fuel consumption per thrust, but the bypass ratio is generally greater than 3, the unit thrust is small, the afterburner 20 cannot be installed, and the flight Mach number cannot exceed 1.0.

Turbofan engines with conventional turbine interstage combustors, i.e., combustion of the added fuel in the transition between high and low pressure turbine or in the vane passages of the low pressure turbine 18, have been calculated to increase the thrust per unit of such engines by over 20% with existing total pressure ratios and pre-turbine temperatures. Because the combustion chamber is added in front of the low-pressure turbine 18, the exhaust speed of the inner culvert of the engine is increased, and thus the exhaust speed of the inner culvert and the outer culvert of the engine cannot be well matched, the engine is provided with a conventional tail nozzle area adjustable mechanism, and a geometric adjustable device is generally required to be arranged on the section of the inner culvert outlet of the engine. In addition, compared to turbofan engines with afterburners 20, the internal duct heating is only provided, and the external duct gas is not provided, so that the thrust enhancement range per unit of the engine is limited, generally about 20% -30%.

The application discloses a trapped vortex combustion chamber 17 and an aeroengine, which are used for solving the technical problem that the axial size of a high-thrust turbine engine is overlarge.

Referring to fig. 2, the trapped vortex combustor 17 includes a first airflow passage and a second airflow passage, the first airflow passage being located between the outer casing 1 and the inner casing 4; an inner duct outer casing 7 is arranged between the outer duct inner casing 4 and the inner duct inner casing 8, and a second airflow channel is formed between the inner duct outer casing 7 and the inner duct inner casing. A first combustion chamber 21 is formed on one side of the outer culvert inner casing 4, which is close to the outer culvert outer casing 1, the first combustion chamber 21 is communicated with a first air flow channel, and a standing vortex is formed in the first combustion chamber 21 when the outer culvert air flow passes through the first air flow channel; the inner duct outer casing 7 is recessed towards one side far away from the inner duct channel to form a second combustion chamber 22 connected with the second air flow channel, and the inner duct air flow forms a standing vortex in the second combustion chamber 22 when passing through the second air flow channel; the first combustion chamber 21 and the second combustion chamber 22 are each connected to a fuel nozzle and a corresponding ignition nozzle 10.

The trapped vortex is formed under the interaction of the air inlet of the front and rear walls of the first combustion chamber 21 and the second combustion chamber 22 and the main flow, and plays roles of flowing down at a high speed, stabilizing flame and improving combustion efficiency. Referring to fig. 3, a cavity is designed in the flow conduit to create a recirculation flow field therein as the air flows through the cavity. The generation of this recirculation zone and flow field characteristics are largely dependent on two factors: the radial pressure gradient is on the one hand large, and on the other hand, when the air flow is involved in the concave cavity, the air flow is turned around to form backflow due to the blocking of the rear wall surface. The vortex in the cavity is thus between the pressure gradient vortex and the streamline vortex. Meanwhile, due to the protection of the concave cavity, the vortex is less influenced by the main flow, has stable performance and is very suitable for flame stabilization.

An outer duct flow stabilizer 2 is arranged in the first air flow channel, the outer duct flow passes through the outer duct flow stabilizer 2 and then enters the first combustion chamber 21, an inner duct flow stabilizer 9 is arranged in the second air flow channel, and the inner duct flow passes through the inner duct flow stabilizer 9 and then enters the second combustion chamber 22. The connotation stabilizer and the connotation stabilizer are both V-shaped groove support plate type stabilizers, and combustion flame stabilization of the first combustion chamber 21 and the second combustion chamber 22 can be realized.

The first combustion chamber 21 and the second combustion chamber 22 are both annular rings surrounding the inner casing 8. The first combustion chamber 21 and the second combustion chamber 22 which are annularly arranged can uniformly heat all the airflows flowing through the inner duct and the outer duct, so that the flow velocity of the airflows in all directions on the circumference is ensured to be consistent, and stable thrust is obtained. The first combustion chamber 21 is closely attached to the second combustion chamber 22, and the ignition electric nozzles 10 of the second combustion chamber 22 of the first combustion chamber 21 are commonly connected with an ignition cable 3, so that the internal structure of the engine can be more compact, and the wiring length can be shortened.

Referring to fig. 4, an aeroengine includes: the outer casing 1 of the outer duct, the inner casing 4 of the outer duct, the inner casing 8 of the inner duct and the standing vortex combustion chamber 17 are arranged, an outer duct channel for the passage of outer duct airflow is formed between the outer casing 1 of the outer duct and the inner casing 4 of the outer duct, the outer duct channel is communicated with the first airflow channel, an inner duct channel for the passage of inner duct airflow is formed between the inner casing 4 of the outer duct and the inner casing 8 of the inner duct, and the inner duct channel is communicated with the second airflow channel.

An air inlet guide 11, a fan 12, a high-pressure compressor 13, a main combustion chamber 14, a high-pressure turbine working blade 16 and a low-pressure turbine 18 are sequentially arranged from an air inlet end to an air outlet end in the engine, and a trapped vortex combustion chamber 17 is positioned between the high-pressure turbine working blade 16 and the low-pressure turbine 18. A high-pressure turbine guide vane 15 is arranged between the main combustion chamber 14 and the high-pressure turbine working vane 16, a third combustion chamber is formed on the high-pressure turbine guide vane 15, a standing vortex is formed in the third combustion chamber 23 when the air flow in the main combustion chamber 14 passes through the high-pressure turbine guide vane 15, and a fuel nozzle and an ignition electric nozzle 10 are arranged in the third combustion chamber 23.

The high-pressure turbine guide vanes 15 are uniformly arranged at intervals around the circumference of the inner casing 8 of the inner duct, the fuel nozzles on all the high-pressure turbine guide vanes 15 synchronously spray fuel, the fuel nozzles of the high-pressure turbine guide vanes 15 are hollow cone nozzles, and the included angle between the hollow cone nozzles and the inlet of the third combustion chamber 23 is 60 degrees. After the fuel oil sprayed from the hollow cone nozzle enters the third combustion chamber 23, vortex is formed in the third combustion chamber 23 under the action of incoming flow at the inlet of the third combustion chamber 23, so that the fuel oil can be better mixed with air flow, and the combustion efficiency is improved.

The air outlet end of the outer casing 1 of the outer culvert is provided with an adjustable tail nozzle 19, the area of the tail nozzle channel can be adjusted when the tail nozzle is adjusted, and then the distribution of the expansion ratio of air flow in the turbine and the tail nozzle can be changed, so that the control of the working state of the whole engine is realized. The adjustable tail jet is typically comprised of a housing, a flow controller, and an actuator. The housing has a plurality of flow channels designed therein to provide a specific flow field configuration for the passage of gas therethrough. The flow controller is the core of the spout and is typically comprised of several movable vanes, shutters or other similar components. These components can be turned on or off as needed to properly adjust the gas flow and velocity. The actuator is a component for controlling the on-off state of the flow controller, and is usually composed of an electric motor, a hydraulic motor, a pneumatic piston or the like. The actuator receives instructions from the electronic control unit and then pushes or pulls the components of the flow controller in accordance with the instructions to effect the adjustment of the orifice size.

The reason why this configuration can improve performance is as follows: conventional turbofan engines employ a brayton cycle, and increasing the pre-turbine temperature is an effective means of increasing the thrust-to-weight ratio of the engine, however subject to technical and cost-limited cycling, and limited range of pre-turbine temperature increases. Changing the single point heating of the main combustion chamber 14 of a conventional turbofan engine to a multi-point heating mode is an effective way to boost engine thrust under existing material and cooling technology limitations. Multiple thermal node warming includes both afterburner 20 and inter-stage combustion modes.

Afterburner 20 has problems of large size and weight, high fuel consumption, inability to cruise for a long time, and significant degradation in the matching efficiency of the hair. Compared with the adoption of the turbine interstage combustion chamber, the scheme can improve the comprehensive efficiency of the unmanned aerial vehicle while improving the thrust and thrust-weight ratio of the engine.

Compared with the conventional turbine interstage combustion chamber which is heated only in the engine content, the application utilizes one part to realize simultaneous heating of the inner and outer content, thereby increasing the exhaust speed of the engine, improving the thrust, reducing the geometrical regulating mechanism of the engine content and reducing the complexity and weight of the engine.

Regarding the reason for the length reduction of the present configuration: compared with the traditional forced turbofan engine, the engine of the present configuration eliminates the forced-air combustion chamber 20 and changes the form of concave cavity, thereby obviously shortening the axial length of the engine; compared with an engine of the forced turbofan 12, the heating node of the high-pressure turbine guide vane 15 and the trapped vortex combustion chamber 17 is positioned in the middle of a high-pressure turbine with higher pressure and behind the high-pressure turbine, and compared with the forced turbofan engine, the forced turbofan engine has the advantages that the heating environment pressure of the heating node of the high-pressure turbine guide vane 15 and the trapped vortex combustion chamber 17 is high, the speed is lower, the combustion efficiency is improved, the circulation efficiency is improved, the oil consumption is lower, and the aviation time is increased; the length is shorter to facilitate mating with the aircraft.

Compared with a conventional turbofan engine with a medium-large bypass ratio, the scheme can obviously improve the power-weight ratio of the engine by utilizing the supplemental heating of the outer bypass of the engine, so that both supersonic speed and long-endurance are considered, and the turbofan engine can be used for a supersonic aircraft. Compared with a conventional interstage combustion turbofan engine, the scheme adopts ideas of external culvert heating, internal heating of the high-pressure turbine guide vanes 15 and the like, has higher heat efficiency, and can cancel an internal culvert area adjusting mechanism because the external culvert is introduced to heat, so that the pressure of the internal culvert and the external culvert of the mixer inlet is easier to balance.

The engine working process is as follows:

in the starting stage of the engine, the main combustion chamber 14 is ignited, and the high-pressure turbine does work to drive the air compressor to increase the rotating speed so as to achieve slow running;

in the take-off stage, heating nodes of the high-pressure turbine guide vane 15 and the standing vortex combustion chamber 17 are sequentially opened, so that the take-off thrust of the engine is improved, and short-distance take-off and landing are realized;

climbing stage: the oil supply quantity of the heating nodes of the high-pressure turbine guide vane 15 and the standing vortex combustion chamber 17 is gradually reduced, and the engines are closed one by one after approaching the cruising altitude;

cruising phase: when the subsonic cruising device is in subsonic cruising, only the first heating point of the main combustion chamber 14 works, and when the subsonic cruising device is in supersonic speed, the heating nodes of the high-pressure turbine guide vane 15 and the trapped vortex combustion chamber 17 heat different gears according to the thrust requirement, so that the supersonic speed flight capability is realized.

Computational simulation analysis based on engineering thermodynamics shows that after the inner and outer cavities of the two sides of the trapped vortex combustor 17 are introduced to heat the inner and outer cavities simultaneously, the unit thrust of the engine is increased by more than 40%, and compared with a turbofan engine with the afterburner 20, the maximum state fuel consumption is reduced by about 30%, and the unit thrust can be increased by more than 20% compared with a scheme of using the interstage combustor only. The outer duct is closed by closing, and only the heating function of the inner duct is reserved, so that the oil consumption of the engine can be continuously reduced, and the long-endurance can be more favorably maintained. The engine performance benefit simulation is analyzed as follows.

The above description is only of the preferred embodiments of the present application and is not intended to limit the present application, but various modifications and variations can be made to the present application by those skilled in the art. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present application should be included in the protection scope of the present application.

Claims (10)

1. A trapped vortex combustor, comprising:

the first airflow channel is positioned between the outer casing (1) of the outer duct and the inner casing (4) of the outer duct;

the first combustion cavity (21) is positioned on one side of the outer duct inner casing (4) close to the outer duct outer casing (1), and is formed by the outer duct inner casing (4) being recessed inwards in the radial direction, the first combustion cavity (21) is communicated with the first air flow channel, and trapped vortex is formed in the first combustion cavity (21) when the outer duct air flow passes through the first air flow channel;

the second air flow channel is positioned between the inner duct outer casing (7) and the inner duct inner casing (8);

the second combustion chamber (22) is positioned on one side of the inner duct outer casing (7) close to the inner duct inner casing (8) and is formed by the inner duct outer casing (7) being recessed outwards along the radial direction, the second combustion chamber (22) is communicated with the second airflow channel, and the inner duct airflow forms a standing vortex in the second combustion chamber (22) when passing through the second airflow channel;

the first combustion chamber (21) and the second combustion chamber (22) are both connected with a fuel nozzle and a corresponding ignition electric nozzle (10).

2. The trapped vortex combustor of claim 1, wherein,

the novel combustion engine is characterized in that an outer culvert flow stabilizer (2) is arranged in the first air flow channel, and the outer culvert air flow enters the first combustion cavity (21) after passing through the outer culvert flow stabilizer (2), an inner culvert flow stabilizer (9) is arranged in the second air flow channel, and the inner culvert air flow enters the second combustion cavity (22) after passing through the inner culvert flow stabilizer (9).

3. The trapped vortex combustor of claim 2, wherein:

the first combustion chamber (21) and the second combustion chamber (22) are annular surrounding the inner casing (8) of the inner duct.

4. The trapped vortex combustor and aeroengine of claim 1, wherein,

the first combustion chamber (21) is closely attached to the second combustion chamber (22), and the first combustion chamber (21) and the second combustion chamber (22) are connected with an ignition cable (3) together through the ignition electric nozzle (10).

5. An aeroengine, which is characterized in that,

the novel combustion chamber comprises an outer casing (1), an inner casing (4), an inner casing (8) and the standing vortex combustion chamber (17) as claimed in any one of claims 1-4, wherein an outer casing channel for passing outer air flow is formed between the outer casing (1) and the inner casing (4), the outer casing channel is communicated with a first air flow channel, an inner casing channel for passing inner air flow is formed between the inner casing (7) and the inner casing (8), and the inner casing channel is communicated with a second air flow channel.

6. The aircraft engine of claim 5, wherein the aircraft engine is configured to control the engine speed,

the aeroengine further comprises an air inlet guide (11), a fan (12), a high-pressure air compressor (13), a main combustion chamber (14), a high-pressure turbine working blade (16) and a low-pressure turbine (18) which are sequentially arranged from an air inlet end to an air outlet end, and the trapped vortex combustion chamber (17) is located between the high-pressure turbine working blade (16) and the low-pressure turbine (18).

7. The aircraft engine of claim 6, wherein the aircraft engine is configured to control the engine speed,

the high-pressure turbine guide vane (15) is arranged between the main combustion chamber (14) and the high-pressure turbine working vane (16), a third combustion chamber (23) is formed in the high-pressure turbine guide vane (15), and a standing vortex is formed in the third combustion chamber (23) when the air flow in the main combustion chamber (14) passes through the high-pressure turbine guide vane (15), and a fuel nozzle and an ignition electric nozzle (10) are arranged in the third combustion chamber (23).

8. The aircraft engine of claim 6, wherein the aircraft engine is configured to control the engine speed,

an adjustable tail nozzle (19) is arranged at the air outlet end of the outer casing (1) of the outer culvert.

9. The aircraft engine of claim 7, wherein the aircraft engine is configured to control the engine speed,

the high-pressure turbine guide vanes (15) are uniformly arranged at intervals around the circumference of the inner casing (8) of the inner culvert, and the fuel nozzles on all the high-pressure turbine guide vanes (15) synchronously spray fuel.

10. The aircraft engine of claim 7, wherein the aircraft engine is configured to control the engine speed,

the fuel nozzle of the high-pressure turbine guide vane (15) is a hollow cone nozzle, and the incoming flow included angle between the hollow cone nozzle and the inlet of the third combustion chamber (23) is 60 degrees.