Trapped vortex combustion chamber and aircraft engine

Abstract

The present application discloses a trapped vortex combustion chamber and an aircraft engine, wherein the trapped vortex combustion chamber includes a first airflow channel located between an outer casing of an outer duct and an inner casing of an outer duct; a first combustion chamber located on a side of the inner casing of the outer duct close to the outer casing of the outer duct, the first combustion chamber being connected to the first airflow channel, and a trapped vortex is formed in the first combustion chamber when the outer duct airflow passes through the first airflow channel; a second airflow channel located between the outer casing of an inner duct and an inner casing of an inner duct; a second combustion chamber located on a side of the outer casing of an inner duct close to the inner casing of the inner duct, the second combustion chamber being connected to the second airflow channel, and a trapped vortex is formed in the second combustion chamber when the inner duct airflow passes through the second airflow channel; the first combustion chamber and the second combustion chamber are both connected to a fuel nozzle and a corresponding ignition nozzle. The present application has the effect of obtaining greater thrust without increasing the axial size of the outer duct.

Description

Technical Field

[0001] The present invention relates to the technical field of turbine engines, and in particular to a trapped vortex combustion chamber and an aircraft engine. Background Art

[0002] Currently, mainstream turbofan engines fall into three categories: afterburning turbofan engines for high-speed fixed-wing aircraft (such as military fighter jets), medium-to-high bypass ratio turbofan engines for subsonic fixed-wing aircraft, and turbofan engines with conventional interstage combustion chambers. Afterburning turbofan engines can operate in either non-afterburning or afterburning modes. The afterburning mode generates approximately 60% more thrust than the non-afterburning mode, but fuel consumption increases by 150% to 200%. Therefore, afterburning is generally used only briefly in special situations such as dogfights, and its operating time is extremely limited. The significant advantage of medium-to-high bypass ratio turbofan engines is their low fuel consumption per unit thrust. However, their bypass ratio is generally greater than 3, resulting in low unit thrust and the inability to install an afterburner. Consequently, their flight Mach number cannot exceed 1.0. Compared to turbofan engines with afterburners, turbofan engines with conventional interstage combustion chambers heat only the inner duct, leaving the outer bypass gas unheated. This results in a limited increase in unit thrust, typically around 20% to 30%.

[0003] Therefore, the thrust of existing turbofan engines is limited. If an afterburner turbofan engine is used, the axial size of the engine will be too large due to the need to add an afterburner combustion chamber, which is not conducive to the design and layout of the aircraft. Summary of the Invention

[0004] The present invention provides a trapped vortex combustion chamber and an aero-engine, so as to solve the technical problem that the axial dimension of a large-thrust turbine engine is too large.

[0005] According to one aspect of the present invention, there is provided a trapped vortex combustion chamber, comprising a first airflow channel, located between an outer duct outer casing and an outer duct inner casing; a first combustion chamber, located on a side of the outer duct inner casing close to the outer duct outer casing, the first combustion chamber being communicated with the first airflow channel, and a trapped vortex is formed in the first combustion chamber when the outer duct airflow passes through the first airflow channel; a second airflow channel, located between the inner duct outer casing and the inner duct inner casing; a second combustion chamber, located on a side of the inner duct outer casing close to the inner duct inner casing, the second combustion chamber being communicated with the second airflow channel, and a trapped vortex is formed in the second combustion chamber when the inner duct airflow passes through the second airflow channel; both the first combustion chamber and the second combustion chamber are connected to a fuel nozzle and a corresponding ignition nozzle.

[0006] By adopting the above technical solution, the outer bypass airflow forms a trapped vortex in the first combustion chamber, and the inner bypass airflow forms a trapped vortex in the second combustion chamber. The airflow in the trapped vortex is a backflow and is located in a concave cavity. It is less affected by the mainstream and has stable performance. Therefore, the flame formed after fuel injection and ignition in the first and second combustion chambers is relatively stable. Conventional turbofan engines use the Brayton cycle. Raising the temperature before the turbine is an effective means to improve the engine's thrust-to-weight ratio. However, due to technical and cost limitations, the increase in the temperature before the turbine is limited. Conventional turbofan engines often heat at a single point in the main combustion chamber. However, this application adds a first and second combustion chamber to the inner and outer bypass ducts, which work together with the main combustion chamber for heating, effectively raising the temperature before the turbine and thereby increasing the thrust of the engine. Compared with traditional afterburner turbofan engines, this configuration does not require an afterburner after the turbine, but instead uses a concave cavity, which significantly shortens the axial length of the engine. Therefore, this solution can achieve greater thrust without increasing the additional axial dimension.

[0007] Compared with the conventional turbine interstage combustion chamber which only heats the engine interior, the first combustion chamber and the second combustion chamber provided in the present invention can not only simultaneously spray fuel and ignite to achieve simultaneous heating of the inner and outer casings, but also spray fuel and ignite in only one of the cavities to heat one of the inner and outer casings, thereby increasing the engine exhaust velocity and thus improving the thrust, while reducing the engine interior geometric adjustment mechanism and reducing the engine complexity and weight.

[0008] Optionally, an outer duct flow stabilizer is provided in the first airflow channel, and the outer duct airflow enters the first combustion chamber after passing through the outer duct flow stabilizer. An inner duct flow stabilizer is provided in the second airflow channel, and the inner duct airflow enters the second combustion chamber after passing through the inner duct flow stabilizer.

[0009] By adopting the above technical solution, the inner duct flow stabilizer and the outer duct flow stabilizer can reduce the flow velocity. The inner duct flow stabilizer and the second combustion chamber, and the outer duct flow stabilizer and the first combustion chamber respectively form recirculation zones, so that trapped vortices are better formed in the first combustion chamber and the second combustion chamber to improve combustion efficiency.

[0010] Optionally, the first combustion chamber and the second combustion chamber are both ring-shaped and surround the inner casing.

[0011] By adopting the above technical solution, the annularly arranged first combustion chamber and the second combustion chamber can uniformly heat all airflows flowing through the inner duct and the outer duct, thereby ensuring that the flow rate of the airflow in all directions on the circumference is consistent to obtain stable thrust.

[0012] Optionally, the first combustion chamber is close to the second combustion chamber, and the ignition nozzles of the first combustion chamber and the second combustion chamber are connected to an ignition cable.

[0013] By adopting the above technical solution, the internal structure of the engine can be made more compact and the wiring length can be shortened.

[0014] According to another aspect of the present invention, an aircraft engine is also provided, which includes an outer duct outer casing, an outer duct inner casing, an inner duct inner casing and the above-mentioned trapped vortex combustion chamber, an outer duct channel for the outer duct outer casing and the outer duct inner casing for the outer duct airflow to pass through is formed between the outer duct outer casing and the outer duct inner casing, and the outer duct channel is connected to the first airflow channel, and an inner channel for the inner duct airflow to pass through is formed between the outer duct inner casing and the inner duct inner casing, and the inner channel is connected to the second airflow channel.

[0015] Optionally, it also includes an air intake guide, a fan, a high-pressure compressor, a combustion chamber, high-pressure turbine working blades and a low-pressure turbine arranged in sequence from the air intake end to the air outlet end, and the trapped vortex combustion chamber is located between the high-pressure turbine working blades and the low-pressure turbine.

[0016] By adopting the above technical solution, the first combustion chamber and the second combustion chamber of this solution are located after the high-pressure turbine. Compared with the afterburner turbofan engine with the afterburner combustion chamber set after the turbine, the heating environment pressure of this solution is high and the speed is lower, which is conducive to improving combustion efficiency and cycle efficiency, lowering fuel consumption, and increasing flight time.

[0017] Optionally, a high-pressure turbine guide vane is arranged between the combustion chamber and the high-pressure turbine working blades, and a third combustion chamber is opened on the high-pressure turbine guide vane. When the airflow in the combustion chamber passes through the high-pressure turbine guide vane, a trapped vortex is formed in the third combustion chamber, and a fuel nozzle and an ignition nozzle are provided in the third combustion chamber.

[0018] By adopting the above technical solution, on the one hand, the high-pressure turbine guide vanes play a better role in guiding the flow. On the other hand, by arranging a third combustion chamber on the high-pressure turbine guide vanes, the heating environment pressure of the third combustion chamber is high and the air flow velocity is low, which is conducive to combustion. By adding new combustion points, the engine turbine pre-temperature can be effectively increased, thereby increasing the engine thrust.

[0019] Optionally, an adjustable tail nozzle is provided at the air outlet end of the outer casing of the duct.

[0020] By adopting the above technical solution, the tail nozzle pipe area can be adjusted when the tail nozzle is adjusted, thereby changing the distribution of the expansion ratio of the airflow in the turbine and the tail nozzle, thereby realizing the control of the entire engine working state.

[0021] Optionally, a plurality of high-pressure turbine guide vanes are evenly spaced circumferentially around the inner casing of the inner duct, and the fuel nozzles on all the high-pressure turbine guide vanes spray fuel synchronously.

[0022] By adopting the above technical solution, setting up multiple high-pressure turbine guide vanes means setting up the same number of combustion points. The distribution of the combustion points along the circumferential direction can make the heating more uniform. The purpose of synchronous fuel injection by the fuel nozzles on all high-pressure turbine guide vanes is also to improve the consistency of combustion at multiple combustion points to achieve uniform heating.

[0023] Optionally, the fuel nozzle of the high-pressure turbine guide vane is a hollow cone nozzle, and the angle between the hollow cone nozzle and the inlet flow of the third combustion chamber is 60°.

[0024] By adopting the above technical solution, after the fuel sprayed from the hollow cone nozzle enters the third combustion chamber, a vortex is formed in the third combustion chamber under the action of the flow from the inlet of the third combustion chamber, which can better mix with the airflow and improve combustion efficiency.

[0025] In summary, this application includes at least one of the following beneficial technical effects:

[0026] 1. By adding heating points in the inner and outer ducts, the traditional single-point heating mode is changed to a multi-point heating mode, which can effectively increase the temperature before the turbine and thus improve the engine thrust;

[0027] 2. The afterburner chamber behind the turbine has been eliminated and replaced with a concave cavity, which significantly shortens the axial length of the engine.

[0028] In addition to the above-described objects, features and advantages, the present invention has other objects, features and advantages. The present invention will be further described in detail below with reference to the accompanying drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0029] The accompanying drawings, which constitute part of this application, are intended to provide a further understanding of the present invention. The exemplary embodiments of the present invention and their descriptions are intended to explain the present invention and do not constitute an undue limitation of the present invention. In the accompanying drawings:

[0030] Figure 1 It is a schematic diagram of the structure of an afterburning turbofan engine;

[0031] Figure 2 2. It is a schematic structural diagram of a trapped vortex combustion chamber according to a preferred embodiment of the present invention;

[0032] Figure 3 Schematic diagram of the airflow direction of the trapped vortex of the present invention;

[0033] Figure 4 1 is a schematic structural diagram of an aircraft engine according to a preferred embodiment of the present invention;

[0034] Figure 5 A side view of a high-pressure turbine guide vane according to a preferred embodiment of the present invention.

[0035] Legend:

[0036] 1. Outer casing of the outer duct; 2. Flow stabilizer of the outer duct; 3. Ignition cable; 4. Inner casing of the outer duct; 5. Fuel nozzle of the second combustion chamber; 6. Fuel nozzle of the first combustion chamber; 7. Outer casing of the inner duct; 8. Inner casing of the inner duct; 9. Flow stabilizer of the inner duct; 10. Ignition nozzle; 11. Inlet guide vane; 12. Fan; 13. High-pressure compressor; 14. Main combustion chamber; 15. High-pressure turbine guide vanes; 16. High-pressure turbine working blades; 17. Trapped vortex combustion chamber; 18. Low-pressure turbine; 19. Adjustable tail nozzle; 20. Afterburner; 21. First combustion chamber; 22. Second combustion chamber; 23. Third combustion chamber. DETAILED DESCRIPTION

[0037] The embodiments of the present invention are described in detail below with reference to the accompanying drawings. However, the present invention can be implemented in many different ways as defined and covered below.

[0038] The following is combined with Figure 1 —4 provides further details of this application.

[0039] Existing turbofan aircraft engines mainly include the following three categories: afterburning turbofan engines, medium and large bypass ratio turbofan engines, and turbofan engines with conventional turbine interstage combustion chambers.

[0040] Reference Figure 1 The existing afterburner turbofan engine is generally used for high-speed fixed-wing aircraft (such as military fighter jets). It is divided into inner and outer ducts, and consists of a fan 12, a compressor, a combustion chamber, a turbine, an afterburner combustion chamber 20, a nozzle, etc.

[0041] The operating principle of an afterburning turbofan engine is as follows: After entering the engine, air is divided into an inner duct flow and an outer duct flow. The inner duct flow is compressed by the inner high-pressure compressor 13 and then enters the inner combustion chamber for combustion, converting it into high-temperature, high-pressure combustion gas. This combustion gas drives the inner high-pressure and low-pressure turbines 18 (part of the internal energy of the combustion gas is converted into mechanical energy for the turbines to drive the inner high-pressure compressor 13 and the inner and outer ducts of the fan 12). The combustion gas then enters the nozzle and mixes with the outer duct flow. The outer duct flow is compressed by the outer duct of the fan 12 and then enters the nozzle. After mixing with the inner duct flow, it is ejected out of the engine at high speed through the nozzle. The mixed exhaust gas from the inner and outer ducts generates a reaction force on the engine, which is the engine thrust. When the aircraft requires greater engine thrust, because the combustion gas after the turbine and the outer duct flow still have residual oxygen that has not been completely burned, additional fuel can be injected into the afterburner 20 after the turbine for further combustion, further increasing the internal energy of the combustion gas and increasing the engine thrust.

[0042] Afterburning turbofan engines can operate in either non-afterburning or afterburning modes. The afterburning mode produces approximately 60% more thrust than the non-afterburning mode, but fuel consumption increases by 150% to 200%. Therefore, afterburning is typically used only briefly in special situations such as dogfights, and its use is extremely limited. The following table lists data for typical foreign afterburning turbofan engines.

[0043]

[0044] On the other hand, to ensure sufficient combustion of the high-speed combustion gases within the afterburner 20, the afterburner 20 is typically long, occupying over 30% of the total engine length, thus increasing the overall engine length. In summary, existing afterburner turbofan engines have excessive fuel consumption in the afterburner mode, and the afterburner 20 occupies an excessively long length.

[0045] Medium-to-high bypass ratio turbofan engines are often used in subsonic fixed-wing aircraft. These aircraft, such as civil airliners, business jets, and reconnaissance drones, prioritize endurance over maneuverability and supersonic speed. Therefore, medium-to-high bypass ratio turbofan engines are often used. The engine structure is shown in the figure below. The engine also has internal and external bypass sections, consisting of a fan 12, compressor, combustion chamber, turbine, afterburner 20, and nozzle. The operating principle is similar to that of an afterburning turbofan engine. The main difference is that the fan 12 of this power unit is larger in diameter, the exhaust expansion ratio is lower, and the afterburner 20 cannot be installed. The exhaust velocity is low, resulting in low specific thrust. The significant advantage of this type of engine is low fuel consumption per unit thrust. However, the bypass ratio is generally greater than 3, resulting in low specific thrust, making it impossible to install an afterburner 20, and the flight Mach number cannot exceed 1.0.

[0046] Turbofan engines with conventional inter-turbine combustion chambers, where fuel combustion is added within the transition zone between the high- and low-pressure turbines or within the blade passages of the low-pressure turbine 18, have been shown to increase specific thrust by over 20% under current overall pressure ratios and pre-turbine temperatures. The addition of a combustion chamber before the low-pressure turbine 18 increases the exhaust velocity in the engine's inner duct, leading to poor matching between the exhaust velocities inside and outside the engine. Therefore, in addition to conventional nozzle area adjustment mechanisms, the engine typically requires a geometrically adjustable device for the engine's inner duct outlet cross-section. Furthermore, compared to turbofan engines with afterburners 20, only the inner duct is heated, while the outer duct remains unheated, resulting in a limited increase in specific thrust, typically around 20%-30%.

[0047] The present application discloses a trapped vortex combustion chamber 17 and an aircraft engine to solve the technical problem of excessive axial size of a large thrust turbine engine.

[0048] Reference Figure 2The trapped vortex combustion chamber 17 includes a first airflow channel and a second airflow channel. The first airflow channel is located between the outer duct outer casing 1 and the outer duct inner casing 4. An inner duct outer casing 7 is provided between the outer duct inner casing 4 and the inner duct inner casing 8. 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 the side of the outer duct inner casing 4 close to the outer duct outer casing 1. The first combustion chamber 21 is connected to the first airflow channel. When the outer duct airflow passes through the first airflow channel, a trapped vortex is formed in the first combustion chamber 21. The inner duct outer casing 7 is recessed toward the side away from the inner channel to form a second combustion chamber 22 connected to the second airflow channel. When the inner duct airflow passes through the second airflow channel, a trapped vortex is formed in the second combustion chamber 22. The first combustion chamber 21 and the second combustion chamber 22 are both connected to a fuel nozzle and a corresponding ignition nozzle 10.

[0049] The interaction between the air intake and the main flow at the front and rear walls of the first combustion chamber 21 and the second combustion chamber 22 forms a trapped vortex, which stabilizes the flame and improves the combustion efficiency under high-speed incoming flow. Figure 3 A cavity is designed in the flow duct. When air flows through this cavity, a recirculation flow field is generated within it. The formation and characteristics of this recirculation zone are primarily determined by two factors: the magnitude of the radial pressure gradient, and the obstruction of the rear wall when the airflow enters the cavity, causing it to swirl and form a recirculation flow. The vortex within the cavity is thus somewhere between a pressure gradient vortex and a streamline vortex. Furthermore, due to the protection of the cavity, this vortex is less affected by the main flow, resulting in stable performance and being well-suited for flame stabilization.

[0050] The first airflow channel is provided with an outer duct stabilizer 2, through which the outer duct airflow enters the first combustion chamber 21. The second airflow channel is provided with an inner duct stabilizer 9, through which the inner duct airflow enters the second combustion chamber 22. Both the inner and outer duct stabilizers are V-grooved support plate stabilizers, which can achieve flame stabilization in the first and second combustion chambers 21 and 22.

[0051] Both the first and second combustion chambers 21, 22 are annular, surrounding the inner casing 8. This annular arrangement uniformly heats all airflow through the inner and outer ducts, ensuring consistent airflow velocities in all directions around the circumference for stable thrust. The first combustion chamber 21 is closely adjacent to the second combustion chamber 22, and the ignition nozzles 10 of the first and second combustion chambers 21, 22 are connected to the ignition cable 3, making the engine's internal structure more compact and shortening wiring lengths.

[0052] Reference Figure 4The aircraft engine includes: an outer duct outer casing 1, an outer duct inner casing 4, an inner duct inner casing 8 and the above-mentioned trapped vortex combustion chamber 17. An outer duct channel for the outer duct outer casing 1 and the outer duct inner casing 4 is formed for the outer duct airflow to pass through, and the outer duct channel is connected to the first airflow channel. An inner channel for the inner duct airflow to pass through is formed between the outer duct inner casing 4 and the inner duct inner casing 8, and the inner channel is connected to the second airflow channel.

[0053] The engine comprises an air intake guide 11, a fan 12, a high-pressure compressor 13, a main combustion chamber 14, high-pressure turbine blades 16, and a low-pressure turbine 18, arranged in sequence from the intake end to the exhaust end. A trapped vortex combustion chamber 17 is located between the high-pressure turbine blades 16 and the low-pressure turbine 18. High-pressure turbine guide vanes 15 are located between the main combustion chamber 14 and the high-pressure turbine blades 16. These vanes define a third combustion chamber. When airflow in the main combustion chamber 14 passes through the high-pressure turbine guide vanes 15, a trapped vortex forms within the third combustion chamber 23. The third combustion chamber 23 houses a fuel nozzle and an ignition nozzle 10.

[0054] Multiple high-pressure turbine guide vanes 15 are evenly spaced circumferentially around the inner casing 8. The fuel nozzles on all of these high-pressure turbine guide vanes 15 spray fuel synchronously. These hollow-cone nozzles form a 60° angle with the inlet flow of the third combustion chamber 23. After the fuel ejected from the hollow-cone nozzles enters the third combustion chamber 23, it is stimulated by the inlet flow of the third combustion chamber 23 to form a vortex within the third combustion chamber 23, enabling better mixing with the airflow and improving combustion efficiency.

[0055] The outlet end of the outer casing 1 of the duct is provided with an adjustable tail nozzle 19. When the tail nozzle is adjusted, the area of ​​the tail nozzle duct can be adjusted, thereby changing the distribution of the expansion ratio of the airflow in the turbine and the tail nozzle, thereby achieving control over the operating state of the entire engine. The adjustable tail nozzle is usually composed of components such as a shell, a flow controller and an actuator. Multiple flow channels are designed inside the shell so that a specific flow field structure can be formed when the gas passes through. The flow controller is the core part of the nozzle and is usually composed of several movable blades, gates or other similar components. These components can be opened or closed as needed to appropriately adjust the gas flow and speed. The actuator is the component that controls the on / off state of the flow controller and is usually composed of an electric motor, hydraulic motor or pneumatic piston. The actuator receives instructions from the electronic control unit and then pushes or pulls the components of the flow controller according to the instructions to adjust the size of the nozzle.

[0056] The reason this configuration improves performance: Conventional turbofan engines use the Brayton cycle. Raising the turbine inlet temperature is an effective means of improving the engine's thrust-to-weight ratio. However, due to technical and cost constraints, the extent of this increase is limited. Changing the conventional turbofan engine's main combustion chamber 14 from single-point heating to a multi-point heating mode is an effective way to increase engine thrust within the limitations of existing materials and cooling technologies. Multi-hot-node heating includes both afterburner 20 and interstage combustion modes.

[0057] The afterburner 20 suffers from issues such as large size and weight, high fuel consumption, and the inability to cruise for extended periods, significantly reducing the efficiency of the aircraft's engine matching. Compared to using a turbine interstage combustion chamber, this solution can increase the overall performance of the UAV while boosting engine thrust and thrust-to-weight ratio.

[0058] Compared with conventional turbine interstage combustion chambers that only heat the engine interior, the present invention uses one component to achieve simultaneous heating of the inner and outer chambers, increasing the engine exhaust velocity and thus improving thrust, while reducing the engine's internal geometric adjustment mechanism and reducing the engine's complexity and weight.

[0059] Regarding the reasons for the shortened length of this configuration: compared with the traditional afterburning turbofan engine, the engine of this configuration cancels the afterburner combustion chamber 20 and uses a concave cavity form, which significantly shortens the axial length of the engine; compared with the afterburner turbofan 12 engine, the high-pressure turbine guide vanes 15 and the trapped vortex combustion chamber 17 heating nodes of this configuration are located in the middle and behind the high-pressure turbine with higher pressure. Compared with the afterburning turbofan engine, the heating environment pressure of the high-pressure turbine guide vanes 15 and the trapped vortex combustion chamber 17 heating nodes are high and the speed is lower, which is conducive to improving combustion efficiency and cycle efficiency, lower fuel consumption, and is conducive to increasing flight time; the length is shorter, so as to facilitate matching with the aircraft.

[0060] Compared to conventional medium-to-high bypass ratio turbofan engines, this solution utilizes supplemental heating in the engine's external duct to significantly improve the engine's power-to-weight ratio, thereby achieving both supersonic speed and long endurance, making it suitable for supersonic aircraft. Compared to conventional interstage combustion turbofan engines, this solution utilizes external duct heating and internal heating of the high-pressure turbine guide vanes 15, resulting in higher thermal efficiency. Furthermore, the introduction of heating in the external duct facilitates pressure balance between the internal and external ducts at the mixer inlet, eliminating the internal duct area adjustment mechanism.

[0061] The engine working process is as follows:

[0062] During the engine startup phase, the main combustion chamber 14 is ignited, and the high-pressure turbine works to drive the compressor to increase the speed to reach idle speed;

[0063] During the takeoff phase, the high-pressure turbine guide vanes 15 and the heating nodes of the trapped vortex combustion chamber 17 are opened in sequence to increase the engine's takeoff thrust and achieve short-range takeoff and landing;

[0064] Climbing phase: The fuel supply to the heating nodes of the high-pressure turbine guide vanes 15 and the trapped vortex combustion chamber 17 is gradually reduced, and when the engine approaches the cruising altitude, they are shut down one by one;

[0065] Cruise phase: During subsonic cruise, only the first heating point of the main combustion chamber 14 works. At supersonic speed, the heating nodes of the high-pressure turbine guide vanes 15 and the trapped vortex combustion chamber 17 are heated at different gears according to the thrust requirements to achieve supersonic flight capability.

[0066] Computational simulations based on engineering thermodynamics show that the introduction of dual concave cavities in the trapped vortex combustor (17), which heats both the inner and outer ducts simultaneously, increases engine specific thrust by over 40%. Compared to a turbofan engine with an afterburner (20), peak fuel consumption is reduced by approximately 30%, and compared to a solution using only an interstage combustor, specific thrust is increased by over 20%. By shutting down the outer duct and retaining only the inner duct heating function, engine fuel consumption is further reduced, further facilitating long-range flight. The simulated analysis of engine performance benefits is shown in the table below.

[0067]

[0068] The foregoing description is merely a preferred embodiment of the present invention and is not intended to limit the present invention. Those skilled in the art will readily appreciate that various modifications and variations of the present invention are possible. Any modifications, equivalent substitutions, or improvements made within the spirit and principles of the present invention shall be included within the scope of protection of the present invention.

Claims

1. A trapped vortex combustion chamber, characterized in that: include: A first airflow channel is located between the outer duct outer casing (1) and the inner duct inner casing (4); A first combustion chamber (21) is located on a side of the duct inner casing (4) close to the duct outer casing (1) and is formed by the duct inner casing (4) being recessed radially inward. The first combustion chamber (21) is communicated with the first airflow channel. When the duct airflow passes through the first airflow channel, a trapped vortex is formed in the first combustion chamber (21). The second air flow channel is located between the outer casing (7) of the inner channel and the inner casing (8) of the inner channel; The second combustion chamber (22) is located on a side of the inner channel outer casing (7) close to the inner channel inner casing (8) and is formed by the inner channel outer casing (7) being recessed radially outward. The second combustion chamber (22) is communicated with the second air flow channel. When the inner flow passes through the second air flow channel, a trapped vortex is formed in the second combustion chamber (22). The first combustion chamber (21) and the second combustion chamber (22) are both connected to a fuel nozzle and a corresponding ignition nozzle (10); An outer duct flow stabilizer (2) is provided in the first airflow channel, and the outer duct airflow enters the first combustion chamber (21) after passing through the outer duct flow stabilizer (2). An inner duct flow stabilizer (9) is provided in the second airflow channel, and the inner duct airflow enters the second combustion chamber (22) after passing through the inner duct flow stabilizer (9).

2. The trapped vortex combustion chamber according to claim 1, characterized in that: The first combustion chamber (21) and the second combustion chamber (22) are both ring-shaped and surround the inner casing (8).

3. The trapped vortex combustion chamber according to claim 1, characterized in that: The first combustion chamber (21) is closely attached to the second combustion chamber (22), and the ignition nozzles (10) of the first combustion chamber (21) and the second combustion chamber (22) are connected to an ignition cable (3) in common.

4. An aircraft engine, characterized in that: It comprises an outer duct outer casing (1), an outer duct inner casing (4), an inner duct inner casing (8) and a trapped vortex combustion chamber (17) as described in any one of claims 1 to 3, wherein an outer duct channel for the outer duct outer casing (1) and the outer duct inner casing (4) is formed between the outer duct outer casing (1) and the outer duct inner casing (4) for the outer duct airflow to pass through, and the outer duct channel is connected to a first airflow channel, and an inner duct channel for the inner duct airflow to pass through is formed between the inner duct outer casing (7) and the inner duct inner casing (8), and the inner duct channel is connected to a second airflow channel.

5. The aircraft engine according to claim 4, characterized in that: The aircraft engine further comprises an air inlet guide (11), a fan (12), a high-pressure compressor (13), a main combustion chamber (14), high-pressure turbine working blades (16) and a low-pressure turbine (18) which are arranged in sequence from the air inlet end to the air outlet end. The trapped vortex combustion chamber (17) is located between the high-pressure turbine working blades (16) and the low-pressure turbine (18).

6. The aircraft engine according to claim 5, characterized in that: A high-pressure turbine guide vane (15) is provided between the main combustion chamber (14) and the high-pressure turbine working blade (16). A third combustion chamber (23) is provided on the high-pressure turbine guide vane (15). When the airflow in the main combustion chamber (14) passes through the high-pressure turbine guide vane (15), a trapped vortex is formed in the third combustion chamber (23). A fuel nozzle and an ignition nozzle (10) are provided in the third combustion chamber (23).

7. The aircraft engine according to claim 5, characterized in that: An adjustable tail nozzle (19) is provided at the air outlet end of the outer casing (1) of the duct.

8. The aircraft engine according to claim 6, characterized in that: A plurality of high-pressure turbine guide vanes (15) are evenly spaced around the inner casing (8) in the inner channel, and the fuel nozzles on all the high-pressure turbine guide vanes (15) spray oil synchronously.

9. The aircraft engine according to claim 6, characterized in that: The fuel nozzle of the high-pressure turbine guide vane (15) is a hollow cone nozzle, and the angle between the hollow cone nozzle and the inlet flow of the third combustion chamber (23) is 60 degrees.

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