Multi-engine integrated afterburner system and control method

Abstract

The invention relates to the technical field of aero-engines, in particular to a multi-engine integrated afterburner system and a control method. The system comprises a turbojet engine array, a flow field regulation and control device and an afterburner which are sequentially arranged in the airflow direction, and the flow field regulation and control device is detachably connected in an inlet area of the afterburner and comprises a flow field homogenization assembly and a flow distribution assembly. A combustion runner and a cooling runner are arranged in the afterburner; in the running process of the aircraft, exhaust airflow of the turbojet engine array is gathered in an inlet area of the afterburner, is mixed and homogenized through the flow field homogenization assembly, and then is distributed to the combustion flow channel and the cooling flow channel according to the controllable airflow distribution proportion through the flow distribution assembly. The thermal protection problem during long-time working of the afterburner of the high-Mach-number wide-speed-range engine can be solved, exhaust flow field distortion of multiple engines is improved, and the performance and reliability of the engines are improved.

Description

Technical Field

[0001] This disclosure relates to the field of aero-engine technology, and in particular to a multi-engine integrated afterburner system and control method. Background Technology

[0002] The design of afterburners in traditional military aircraft engines is mainly for short-term operation. The single use time of an afterburner is usually only a few minutes. In this case, the afterburner can meet the thermal protection requirements by relying on heat shields.

[0003] However, with the development of high Mach number wide-range aircraft, afterburners often need to operate continuously for tens of minutes to several hours. Therefore, when multiple engines are connected in parallel to share an afterburner, problems such as insufficient heat sink, flow field distortion of multiple engines, and high control complexity may occur. Summary of the Invention

[0004] This disclosure is made in view of the above-mentioned problems. This disclosure provides a multi-engine integrated afterburner system and its control method.

[0005] According to one aspect of this disclosure, a multi-engine integrated afterburner system is provided, comprising a turbojet engine array, a flow field control device, and an afterburner arranged sequentially along the airflow direction. The flow field control device is detachably connected to the inlet region of the afterburner. The flow field control device includes a flow field homogenization component and a flow distribution component. The flow distribution component is movably connected to one end of the flow field homogenization component away from the inlet region. The interior of the afterburner has a combustion channel and a cooling channel. During the operation of the aircraft, the exhaust airflow of the turbojet engine array converges in the inlet region of the afterburner. After being mixed and homogenized by the flow field homogenization component, the exhaust airflow is distributed to the combustion channel and the cooling channel by the flow distribution component according to a controllable airflow distribution ratio. The controllable airflow distribution ratio is adapted to the operating state of the turbojet engine array and the flight state of the aircraft.

[0006] According to another aspect of this disclosure, a control method is provided for a multi-engine integrated afterburner system provided in embodiments of this disclosure, comprising: During the operation of the aircraft, the exhaust airflow of the turbojet engine array that converges in the inlet region of the afterburner is mixed and homogenized using flow field homogenization components. Using the flow distribution component, the exhaust airflow after mixing and homogenization is distributed to the combustion channel and cooling channel of the afterburner according to a controllable airflow distribution ratio; wherein, the controllable airflow distribution ratio is adapted to the operating state of the turbojet engine array and the flight state of the aircraft.

[0007] As will be described in detail below, in the multi-engine integrated afterburner system according to embodiments of the present disclosure, the exhaust gas from the turbojet engine array converges in the inlet region of the afterburner, receiving the exhaust gas from the turbojet engine array in a shared afterburner configuration. This significantly reduces the wall area requiring active thermal protection and substantially decreases the cooling area of ​​the afterburner, structurally reducing the thermal load on the afterburner during long-term operation. This solves the thermal protection problem of high Mach number, wide-speed-range engine afterburners under long-term operation, ensuring the engine's operational stability requirements. Simultaneously, it eliminates the need for an additional fuel heat sink, simplifying the engine's thermal management system and improving overall engine efficiency. Furthermore, a flow field control device is detachably connected to the inlet region of the afterburner. The flow field homogenization component in this device can be used to mix and homogenize the exhaust gas from the turbojet engine array, improving combustion instability caused by multi-engine exhaust flow field distortion, ensuring the performance and reliability of the turbojet engine array, and creating favorable conditions for stable combustion in the afterburner. The afterburner has combustion channels and cooling channels inside. The flow field control device may also include a flow distribution component, which is movably connected to the end of the flow field homogenization component away from the inlet region. This allows for the segmentation of the exhaust airflow from the turbojet engine array and the distribution of the mixed and homogenized exhaust airflow to the combustion channels and cooling channels according to a controllable airflow distribution ratio. This controllable airflow distribution ratio is adapted to the operating state of the turbojet engine array and the flight state of the aircraft, enabling flexible matching of the power system in a wide speed range and a wide airspace, thereby improving the engine's performance and reliability.

[0008] It should be understood that both the foregoing general description and the following detailed description are exemplary and intended to provide further illustration of the claimed technology. Attached Figure Description

[0009] The above and other objects, features, and advantages of this disclosure will become more apparent from the more detailed description of the embodiments thereof in conjunction with the accompanying drawings. The drawings are provided to further illustrate the embodiments of this disclosure and form part of the specification. They are used together with the embodiments of this disclosure to explain the disclosure and do not constitute a limitation thereof. In the drawings, the same reference numerals generally represent the same components or steps.

[0010] Figure 1A schematic diagram of the structure of a multi-engine integrated afterburner system provided in an exemplary embodiment of this disclosure is shown. Figure 2 A schematic diagram of the structure of a multi-engine integrated afterburner system for removing turbojet engine arrays provided in an exemplary embodiment of this disclosure is shown. Figure 3 This illustration shows a structural schematic diagram of a multi-engine integrated afterburner system that removes the turbojet engine array and outer casing, provided by an exemplary embodiment of this disclosure. Figure 4 A control principle diagram of an engine redundancy control system provided in an exemplary embodiment of this disclosure is shown. Figure 5 A control principle diagram of a multi-engine blockchain provided in an exemplary embodiment of this disclosure is shown; Figure 6 A flowchart illustrating a control method for a multi-engine integrated afterburner system provided in an exemplary embodiment of this disclosure is shown.

[0011] Figure label: 110-Turbojet engine array, 111-Turbojet engine, 120-Flow field control device, 121-Flow field homogenization component, 122-Flow distribution component, 1211-Vortex generator array, 1212-Guide rib, 130-Afterburner chamber, 131-Outer casing and 132-Cylindrical baffle. Detailed Implementation

[0012] To make the objectives, technical solutions, and advantages of this disclosure more apparent, exemplary embodiments according to this disclosure will now be described in detail with reference to the accompanying drawings. Obviously, the described embodiments are merely some embodiments of this disclosure, and not all embodiments of this disclosure. It should be understood that this disclosure is not limited to the exemplary embodiments described herein.

[0013] It should be understood that the steps described in the method embodiments of this disclosure may be performed in different orders and / or in parallel. Furthermore, the method embodiments may include additional steps and / or omit the steps shown. The scope of this disclosure is not limited in this respect.

[0014] The term "comprising" and its variations as used herein are open-ended, meaning "including but not limited to". The term "based on" means "at least partially based on". The term "one embodiment" means "at least one embodiment"; the term "another embodiment" means "at least one additional embodiment"; the term "some embodiments" means "at least some embodiments". Definitions of other terms will be given in the description below. It should be noted that the concepts of "first", "second", etc., used in this disclosure are only used to distinguish different devices, modules, or units, and are not intended to limit the order of functions performed by these devices, modules, or units or their interdependencies.

[0015] It should be noted that the terms "a" and "a plurality of" used in this disclosure are illustrative rather than restrictive, and those skilled in the art should understand that, unless otherwise expressly indicated in the context, they should be understood as "one or more".

[0016] The names of messages or information exchanged between multiple devices in the embodiments of this disclosure are for illustrative purposes only and are not intended to limit the scope of such messages or information.

[0017] The following is a detailed analysis of the potential problems associated with multiple engines sharing an afterburner: insufficient heat sink, flow field distortion across multiple engines, and high control complexity. (1) Analysis of insufficient heat sink: Thermal shielding is primarily designed for short-term use and cannot be actively adjusted to adapt to the variable environments across wide speed ranges and large airspaces. Ramjet engines typically employ fuel regenerative cooling technology, relying on fuel within microchannels in the combustion chamber walls to provide active cooling. However, for high-Mach-number, wide-speed-range engines, fuel regenerative cooling requires covering a large area of ​​the combustion chamber walls. Simultaneously, the lack of a cooling film protection results in extremely high heat flux density (potentially exceeding 1 MW / m²), leading to fuel heat sink requirements far exceeding supply capacity. Traditional fuels (such as RP-3 aviation kerosene) have a heat sink of less than 1 MJ / kg, insufficient for prolonged cooling requirements, significantly increasing the risk of cooling failure.

[0018] (2) Analysis of flow field distortion in multi-engine systems: When multiple engines are connected in parallel and share an afterburner, although the cooling area can be effectively reduced, new challenges arise: due to differences in manufacturing tolerances, operating conditions, and exhaust parameters after the turbine (flow rate deviation of ±15%, temperature fluctuation of ±200℃), strong shear layers and vortices appear at the inlet, causing pressure oscillations (amplitude ≥5% of static pressure), which in turn leads to unstable combustion and reduced efficiency (the reduction can reach 8%), seriously affecting the performance and reliability of the engine.

[0019] However, traditional mixing devices (such as louvers and turbulence columns) are not adaptable to wide speed range conditions. Overmixing at low speeds leads to increased pressure loss, while insufficient mixing at high speeds causes combustion instability.

[0020] (3) Analysis for high control complexity: Traditional centralized control systems have high delays in responding to multi-engine coordinated actions, making it difficult to achieve automatic balancing in the event of a single engine failure, thrust vector control, and real-time balancing of torque disturbances.

[0021] To address the aforementioned issues, this disclosure provides a multi-engine integrated afterburner system and control method to solve the thermal protection problem of afterburners in high Mach number wide-speed-range engines during long-term operation, improve combustion instability caused by exhaust flow field distortion of multiple engines, achieve flexible matching of the power system in a wide speed and airspace range, and improve engine performance and reliability.

[0022] This disclosure provides a multi-engine integrated afterburner system. Figure 1 A schematic diagram of the structure of a multi-engine integrated afterburner system provided in an exemplary embodiment of this disclosure is shown. Figure 2 A schematic diagram of a multi-engine integrated afterburner system for removing turbojet engine arrays, provided by an exemplary embodiment of this disclosure, is shown. Figure 1 As shown, the multi-engine integrated afterburner system includes a turbojet engine array 110 arranged sequentially along the airflow direction. Figure 1 The image only shows the engine nozzles of each turbojet engine 111 in the turbojet engine array 110, the flow field control device 120, and the afterburner 130. The flow field control device 120 is detachably connected to the inlet region of the afterburner 130. Here, the afterburner 130 includes an outer casing 131, and the flow field control device 120 is located within the outer casing 131. Figure 1 and Figure 2 To demonstrate the structure of the flow field control device 120, the outer casing 131 around the flow field control device 120 is omitted.

[0023] Since the flow field control device 120 is detachably connected to the inlet area of ​​the afterburner 130, it can be disassembled and replaced when a malfunction is detected, thus reducing maintenance costs.

[0024] like Figure 1As shown, the turbojet engine array 110, taking four turbojet engines as an example, adopts a 2×2 rectangular array arrangement. The diameter of each turbojet engine is D=300mm, and they are horizontally symmetrically distributed. The turbojet engine array 110 and the flow field control device 120 can be detachably connected or fixed by an external fixing device to ensure that the axis of the turbojet engine array 110 coincides with the axis of the flow field control device 120.

[0025] For high Mach number wide-range engines, fuel regeneration cooling requires covering a large area of ​​the combustion chamber walls. Simultaneously, the lack of a cooling film protection on these walls results in extremely high heat flux density, leading to a fuel heat sink demand far exceeding supply capacity. To address this technical problem, this disclosure draws upon the following embodiments for aero-engines: single-tube main combustion chamber (each main combustion chamber has an independent flame tube and casing), annular main combustion chamber (each main combustion chamber has an independent flame tube, shared by the casing and mixed cooling gas), and ring-shaped main combustion chamber (the main combustion chamber has an integrated annular cavity, shared by the flame tube and casing). Following the historical development approach, this engine utilizes multiple turbojet engines connected in parallel to form a core engine array. A shared afterburner and nozzle (connected to the afterburner 130) receives the turbine exhaust from the turbojet engine array. This significantly reduces the wall area requiring active thermal protection (e.g., when four engines share a combustion chamber, the cooling perimeter is only twice that of a single combustion chamber, while the total perimeter of a single independent combustion chamber is four times). This significantly reduces the cooling area of ​​the afterburner, structurally lowering the thermal load on the afterburner during prolonged operation. Simultaneously, it eliminates the need for an additional fuel heat sink, simplifying the engine's thermal management system and improving overall engine efficiency.

[0026] During the operation of the aircraft, the exhaust gas from the turbojet engine array 110 converges at the inlet region of the afterburner 130. Due to differences in manufacturing tolerances and operating conditions among the individual turbojet engines in the turbojet engine array 110, there are differences in the exhaust parameters after the turbine (flow velocity deviation ±15%, temperature fluctuation ±200℃). This leads to the formation of a strong shear layer and vortices in the inlet region of the afterburner 130, causing pressure oscillations (amplitude ≥5% of static pressure). Consequently, combustion instability and efficiency reduction (up to 8%) occur, severely affecting the performance and reliability of the turbojet engine array 110.

[0027] Based on this, such as Figure 1 and Figure 2As shown, the flow field control device 120 may include a flow field homogenization component 121. The flow field homogenization component 121 can be used to mix and homogenize the exhaust gas flow of the turbojet engine array 110, which can improve the flow field distortion problem caused by the difference in exhaust velocity and temperature of multiple engines, ensure the performance and reliability of the turbojet engine array 110, and create good conditions for stable combustion in the afterburner 130.

[0028] like Figure 1 and Figure 2 As shown, the afterburner 130 has a combustion channel and a cooling channel inside. The flow field control device 120 may also include a flow distribution component 122. The flow distribution component 122 is movably connected to the end of the flow field homogenization component 121 away from the inlet region. This movable connection can be used to adjust the relative positional relationship between the flow distribution component 122 and the flow field homogenization component 121 to achieve the division of the exhaust airflow of the turbojet engine array 110 and the distribution of the mixed and homogenized exhaust airflow to the combustion channel and the cooling channel according to a controllable airflow distribution ratio. This controllable airflow distribution ratio is adapted to the working state of the turbojet engine array and the flight state of the aircraft, realizing flexible matching of the power system in a wide speed range and a wide airspace, and improving the performance and reliability of the engine.

[0029] Here, the controllable airflow distribution ratio can be determined based on the operating status of each pair of turbojet engines in the turbojet engine array 110 and the flight status of the aircraft. For example, when the aircraft is in the takeoff acceleration phase, the controllable airflow distribution ratio is 80%:20% to introduce 80% of the airflow into the combustion channel to meet the engine's high thrust requirements; when the aircraft is in the high-speed cruise phase, the controllable airflow distribution ratio is 50%:50% to introduce 50% of the airflow into the combustion channel, ensuring normal engine operation while reducing the thermal load on the afterburner, expanding the lean fuel shutdown boundary, and improving engine economy.

[0030] Based on this, the multi-engine shared afterburner system provided in this embodiment of the present disclosure converges the exhaust gas from the turbojet engine array into the inlet region of the afterburner, receiving the exhaust gas from the turbojet engine array in a shared afterburner configuration. This significantly reduces the wall area requiring active thermal protection and substantially decreases the cooling area of ​​the afterburner, structurally reducing the thermal load on the afterburner during long-term operation. This solves the thermal protection problem of high Mach number, wide-speed-range engine afterburners under long-term operation, ensuring the engine's operational stability requirements. Simultaneously, it eliminates the need for an additional fuel heat sink, simplifying the engine's thermal management system and improving overall engine efficiency. Furthermore, a flow field control device is detachably connected to the inlet region of the afterburner. The flow field homogenization component in this device can be used to mix and homogenize the exhaust gas from the turbojet engine array, improving combustion instability caused by multi-engine exhaust flow field distortion, ensuring the performance and reliability of the turbojet engine array, and creating favorable conditions for stable combustion in the afterburner. The afterburner has combustion channels and cooling channels inside. The flow field control device may also include a flow distribution component, which is movably connected to the end of the flow field homogenization component away from the inlet region. This allows for the segmentation of the exhaust airflow from the turbojet engine array and the distribution of the mixed and homogenized exhaust airflow to the combustion channels and cooling channels according to a controllable airflow distribution ratio. This controllable airflow distribution ratio is adapted to the operating state of the turbojet engine array and the flight state of the aircraft, enabling flexible matching of the power system in a wide speed range and a wide airspace, thereby improving the engine's performance and reliability.

[0031] In some embodiments, such as Figure 2 As shown, the flow field homogenization component 121 includes multiple guide ribs 1212 that intersect around the axis of the afterburner 130, and the multiple guide ribs 1212 divide the inlet region into multiple independent sub-channels.

[0032] Specifically, the inlet of the afterburner 130 is used to receive the exhaust airflow from four different turbojet engines in the turbojet engine array 110. The exhaust airflow from these four different turbojet engines has different velocity, temperature and pressure distributions. Strong shear layers and large-scale vortices will be formed at the inlet at the confluence, resulting in flow field distortion and instability. A simple confluence chamber cannot effectively solve this problem.

[0033] Based on this, multiple guide ribs 1212 are provided at the inlet of the afterburner 130. These guide ribs 1212 converge around the axis of the afterburner 130, forming an X-shape. For example... Figure 1 and Figure 2As shown, the flow guide assembly consists of four flow guide ribs at 90° to each other. The flow guide ribs are 10 mm thick and have a height of H = 800 mm (flush with the diameter of the afterburner 130). Multiple flow guide ribs 1212 divide the four exhaust airflows from different turbojet engines into eight sub-flows.

[0034] The converging guide ribs act like "structural walls" in the flow field, physically dividing and isolating the originally freely developing and chaotic mixing boundary layer into several independent sub-channels. This avoids direct collisions between core airflows from different sources at the inlet, transforming macroscopic flow field distortion into multiple more controllable microscopic flow problems constrained within the sub-channels.

[0035] Furthermore, the guide ribs provide a clear and stable flow path for each airflow, which significantly reduces the instantaneous global impact of fluctuations in incoming flow parameters (such as sudden changes in the operating conditions of an engine) on the entire afterburner inlet flow field. The flow is "regularized," establishing a predetermined flow field basis with known geometry for downstream combustion, greatly improving the predictability and stability of the combustion process.

[0036] In some embodiments, such as Figure 2 As shown, the flow field homogenization component 121 also includes multiple vortex generator arrays 1211, each of which is fixedly disposed on one end of each guide rib 1212 near the inlet region, for inducing the generation of vortex structures that enhance mixing in the sub-channel.

[0037] Specifically, each vortex generator array 1211 may include multiple triangular vortex generators, which are staggered and arranged on the downstream surface of each guide rib 1212 (the height of the vortex generator is h=50mm and the spacing is s=100mm), inducing the generation of flow vortices and enhancing the exchange of momentum and energy between subcurrents.

[0038] The guide rib 1212 creates a stable and orderly sub-flow channel for the exhaust airflow of the turbojet engine array 110. Multiple vortex generators are fixed on the rib to force the generation of counter-rotating vortex pairs. Through forced mixing, the mixing uniformity between the various airflows is significantly improved, effectively improving the flow field distortion problem caused by the difference in engine exhaust velocity and temperature, and creating favorable conditions for stable combustion in the afterburner.

[0039] Based on this, the embodiments of this disclosure can utilize the flow field homogenization component 121 to achieve refined flow field control of the exhaust airflow of the turbojet engine array 110 by "first guiding the flow to shape it, and then generating vortexes for mixing". The guide ribs solve the problem of macroscopic flow field regularization, and the vortex generator array solves the problem of microscopic mixing.

[0040] In some embodiments, such as Figure 2 As shown, the flow distribution component 122 includes an adjustable guide vane array. Each vane in the adjustable guide vane array is rotatably disposed at the end of each guide rib 1212 away from the inlet region. The controllable airflow distribution ratio is controlled by adjusting the actuation parameters of each vane.

[0041] Specifically, an adjustable guide vane array (also known as a split guide vane array) is set at the outlet of the "X-shaped guide rib". This adjustable guide vane array can dynamically adjust the controllable airflow distribution ratio entering the combustion channel and the cooling channel according to the different operating states of the turbojet engine array and the flight state of the aircraft.

[0042] like Figure 2 As shown, four guide ribs at 90° angles to each other divide the exhaust airflow from four different turbojet engines into eight sub-flows. An adjustable guide vane array is movably connected to the tail of two adjacent guide ribs to distribute each sub-flow to the combustion channel and cooling channel according to a controllable airflow distribution ratio.

[0043] Each adjustable guide vane array includes one set of vanes, with three pairs of vanes in series in each set. The chord length of each vane is c=300mm and the span is b=200mm. Each pair of vanes consists of two vanes. The end of one vane away from the other is rotatably connected to the tail of the guide rib. The ends of the two vanes close to each other are not fixedly connected and can be intersecting or adjacent in space, depending on whether they can achieve flow splitting.

[0044] The actuation parameter of the blade can be the deflection angle of the blade, which has a dynamic adjustment range of ±20°. Here, deflection towards the central axis is defined as "-", deflection outward is defined as "+", and the perpendicularity of the blade to the X-shaped guide rib is defined as 0°. The installation angle of each blade in the adjustable guide blade array can be adjusted independently or in conjunction, depending on actual needs. This disclosure does not impose specific limitations on this.

[0045] The engine's demands differ significantly across flight phases: During takeoff acceleration (Ma=0.8-1.5), more airflow is needed for combustion to meet high thrust requirements. Therefore, the blade angle is set to α=+20°, and the controllable airflow distribution ratio is set to 80%:20%, directing 80% of the airflow into the combustion chamber to meet the engine's high thrust requirements. During high-speed cruise, enhanced cooling and control of combustion temperature are necessary. Therefore, the blade angle is set to α=-10°, and the controllable airflow distribution ratio is set to 50%:50%, directing 50% of the airflow into the combustion chamber while the remaining airflow enters the cooling zone. This cooling effect reduces the combustion chamber wall temperature, extends combustion chamber life, and while ensuring normal engine operation, reduces the thermal load on the afterburner and expands the lean fuel kill zone to meet the requirements of good thermal protection and fuel economy.

[0046] It is evident that the adjustable guide vane array can act as an adjustable aerodynamic valve to "distribute" the total exhaust airflow "on demand" and dynamically and continuously adjust the controllable airflow distribution ratio between the core combustion channel and the peripheral cooling channel according to the aircraft's flight status (such as Mach number and altitude) and engine operating status (such as thrust requirements corresponding to the operating conditions), thereby achieving a real-time optimal balance between thrust and thermal protection.

[0047] In some embodiments, Figure 3 A schematic diagram of a multi-engine integrated afterburner system with the turbojet engine array and outer casing removed, provided by an exemplary embodiment of this disclosure, is shown, as follows: Figure 3 As shown, the interior of the afterburner 130 has a cylindrical baffle 132 arranged in the axial direction, the combustion channel is the internal region of the cylindrical baffle 132, and the cooling channel is the annular region surrounding the outside of the cylindrical baffle 132 inside the afterburner 130. The cylindrical baffle 132 is made of high-temperature resistant composite material. The inner wall surface of the cylindrical baffle 132 facing the combustion channel is made of silicon carbide fiber-reinforced silicon carbide ceramic matrix composite material, and the outer wall surface of the cylindrical baffle 132 facing the cooling channel is made of high-temperature alloy or carbon fiber-reinforced silicon carbide composite material.

[0048] Specifically, such as Figure 1 and Figure 2 As shown, the afterburner 130 has the following shape: the front section of the outer casing 131 (including the part of the outer casing 131 located around the flow field control device 120) is a straight cylindrical section, and the rear section gradually tapers to the nozzle inlet. The diameter of the straight cylindrical section is Φ=800mm, the length is L1=500mm, and the contraction angle of the rear section is θ=15°.

[0049] The afterburner 130 adopts a partitioned design, dividing it into a combustion channel and a cooling channel by a cylindrical baffle 132 arranged along the axial direction. The combustion channel is the core combustion zone located inside the cylindrical baffle 132, with a radius of R1 = 400 mm; the cooling channel is an annular cooling area located inside the afterburner 130 surrounding the cylindrical baffle 132, with a width of ΔR = 200 mm.

[0050] The cylindrical baffle 132 is made of high-temperature resistant composite materials, such as ceramic matrix composites (CMC). CMC materials have excellent high-temperature performance and corrosion resistance, which can effectively withstand the high-temperature environment in the combustion chamber, while effectively isolating the combustion flow channel and cooling flow channel, thereby improving the overall performance and reliability of the combustion chamber.

[0051] The inner wall of the cylindrical baffle 132 facing the combustion channel is made of silicon carbide fiber-reinforced silicon carbide ceramic matrix composite material SiC / SiC, which can withstand high-temperature combustion gas of 1800℃.

[0052] The outer wall of the cylindrical baffle 132 facing the cooling channel is made of high-temperature alloy or carbon fiber reinforced silicon carbide composite material, such as high-temperature alloy GH4169, which has a temperature resistance of over 1000℃; or it can be made of high-temperature alloy or carbon fiber reinforced silicon carbide composite material C / C-SiC. Because of the cold air isolation, the wall temperature of the afterburner can be controlled below 800℃ without the need for fuel intervention.

[0053] In some embodiments, the system may further include: The distributed control system is configured to acquire the operating status of the turbojet engine array and the flight status of the aircraft; generate cooperative control commands based on the operating and flight status; and use the cooperative control commands to synchronously adjust the operating status of the turbojet engine array and the actuation parameters of the flow distribution components to control the controllable airflow distribution ratio.

[0054] Specifically, the distributed control system employs blockchain-based redundant control technology to achieve high-precision control of the characteristics of the four engines in the turbojet engine array, as well as the adjustable guide vane array. The decentralized ledger, automatic execution of smart contracts, and tamper-proof data storage of blockchain technology make it a solution for enhancing multi-engine collaborative control and redundancy management.

[0055] When the operating state of the turbojet engine array and the flight state of the aircraft are different, the required controllable airflow distribution ratio is different. Based on this, the embodiments of this disclosure can add a distributed control system. The distributed control system can be used to acquire the operating state of the turbojet engine array and the flight state of the aircraft in real time; generate cooperative control commands based on the operating state and flight state; and use the cooperative control commands to synchronously adjust the operating state of the turbojet engine array and the operating parameters of the flow distribution component to control the controllable airflow distribution ratio.

[0056] This distributed control system acquires global state information (the operating status of the turbojet engine array and the flight status of the aircraft) and generates a single, coordinated control command. It adjusts the engine's operating parameters and the opening of each blade in the adjustable guide vane array in real time, achieving flexible matching of the power system across a wide speed and airspace range. This includes real-time automatic compensation for torque disturbances, rapid switching between operating modes, thrust vector generation, and automatic reconfiguration of thrust distribution strategies to balance power output in the event of a single engine failure. This ensures a high degree of synchronization between engine thrust adjustment and airflow distribution adjustment. This distributed control system integrates the engine (thrust source) and flow distribution components (energy management valves) into a unified "power energy management system," achieving coordinated control of the turbojet engine array and flow distribution components. It achieves global optimization from the system's top level, rather than local optimization.

[0057] In some embodiments, Figure 4 A control principle diagram of an engine redundancy control system provided in an exemplary embodiment of this disclosure is shown. Figure 5 A control principle diagram of a multi-engine blockchain provided in an exemplary embodiment of this disclosure is shown, such as... Figure 4 and Figure 5 As shown, a distributed control system may include: The edge execution layer includes sensors, controllers, and actuators deployed on each turbojet engine in the turbojet engine array. The sensors are configured to acquire the operating status of each turbojet engine in real time, and the controllers are configured to preprocess the operating status of each turbojet engine. The airborne blockchain layer includes a blockchain network consisting of blockchain nodes deployed on each turbojet engine. The blockchain nodes are configured to generate control commands based on their operating and flight status. The blockchain network is configured to process the control commands from multiple blockchain nodes through a blockchain consensus protocol to generate coordinated control commands. The controller is also configured to adjust the operating state of the turbojet engine array using cooperative control commands; and to send control command signals generated based on the cooperative control commands to the actuators. The actuator is configured to receive control command signals and adjust the operating state of the flow distribution component based on the control command signals.

[0058] Specifically, each engine in the edge execution layer is equipped with a blockchain light node, integrating three modules: a data acquisition module, which acquires raw sensor data (including engine speed, fuel flow, vibration signals, etc.) via the CAN bus; a preprocessing module, which can perform edge computing such as wavelet denoising and Kalman filtering; and a communication module, which achieves microsecond-level synchronization. The airborne blockchain layer mainly realizes trusted collaboration between engines, aiming to achieve data consistency among multiple nodes, i.e., engine control nodes, while also enabling rapid detection and recovery of faulty nodes.

[0059] like Figure 4 As shown, each turbojet engine is equipped with multiple control nodes, which are interconnected to form a blockchain structure. After executing Aero-PBF consensus, if consensus is reached, the blockchain nodes release control command signals based on cooperative control instructions. These control command signals are then transmitted to the actuators, achieving highly reliable data storage and rapid transmission. This technology enables real-time and precise monitoring and control of the characteristic parameters (such as speed, exhaust temperature, and pressure) of the four turbojet engines, while also allowing for high-precision adjustment of the blade angle and opening degree of the adjustable guide vane array. If consensus times out, a local security strategy is triggered, implementing degraded mode control.

[0060] In terms of hardware architecture, each turbojet engine is equipped with an independent sensor node (including pressure, temperature, and speed sensors), and the blade drive mechanism integrates an angle encoder and a servo motor; a blockchain network is built using distributed computing nodes (such as FPGA clusters), and data is synchronized between nodes via optical fiber (delay ≤10μs).

[0061] From a control algorithm perspective, a real-time trim module is configured. If a torque disturbance occurs in one turbojet engine, the distributed control system can quickly transmit the disturbance information to other turbojet engines and corresponding control components via a blockchain network. This allows for real-time adjustment of the thrust and adjustable guide vane array status of each turbojet engine, achieving precise torque trimming and ensuring stable operation of the engine system. Based on the extended Kalman filter algorithm, the blade angle is dynamically adjusted according to differences in engine exhaust parameters (ΔT≤±50K, ΔP≤±0.1MPa), ensuring that the uniformity error of the inlet airflow in the combustion zone is ≤±3%.

[0062] It is also equipped with a thrust vectoring module, which generates a ±5° yaw moment by asymmetrically adjusting the blade angles on the left and right sides (e.g., left side α=+10°, right side α=-10°), with a vector efficiency ≥0.8.

[0063] When the operating mode switches from cruise mode to afterburner mode, the distributed control system coordinates the actions of various components based on the preset control strategy and real-time monitoring data to quickly complete the mode switch and meet the power requirements of different flight stages.

[0064] In some embodiments, the blockchain network is also configured to, if a turbojet engine in the turbojet engine array fails, process control commands from other blockchain nodes through a blockchain consensus protocol to generate collaborative control commands.

[0065] Specifically, from the control algorithm perspective, a redundant fault-tolerant module is configured: if one of the turbojet engines suddenly fails, the distributed control system can immediately and automatically balance the thrust and operating status of the remaining turbojet engines based on the fault information and backup control strategy stored in the blockchain, ensuring the aircraft continues to fly safely and greatly improving the reliability and adaptability of the entire engine system. When a turbojet engine fails, the guide vane angle of the corresponding subflow of the failed turbojet engine is reduced within 300ms to decrease the proportion of airflow in the combustion zone; the vane angles of other subflow zones are adjusted by +15° to increase the proportion of airflow in the combustion zone, maintaining the total thrust ≥85% of the rated value.

[0066] From a control principle perspective, the implementation of the redundancy strategy is mainly based on dynamically adjusting the weight of turbojet engine nodes according to their real-time health status. When a turbojet engine fails to transmit data, its weight is reduced to zero, and it is prohibited from participating in consensus.

[0067] In summary, the multi-engine integrated afterburner system provided in this disclosure, through innovative structural design and advanced control technology, effectively solves the thermal protection problems of afterburners during long-term operation and the combustion instability issues caused by multiple engines operating in parallel, providing efficient and reliable propulsion for high Mach number, wide-speed-range aircraft. The specific technical effects it brings can include: (1) Reduced heat load: Multiple turbojet engines share a single afterburner, which significantly reduces the cooling area of ​​the afterburner and thus reduces the heat load on the afterburner during long-term operation. At the same time, there is no need to introduce an additional fuel heat sink, which simplifies the engine's thermal management system and improves the overall efficiency of the engine.

[0068] (2) Improved airflow mixing: The flow guide-mixing composite structure enhances the efficient mixing of exhaust airflow from multiple turbojet engines through the synergistic effect of the "X-shaped flow guide ribs" and the cross vortex generator. It effectively solves the problem of flow field distortion caused by the difference in engine exhaust velocity and temperature, improves the combustion stability of the combustion chamber, reduces vibration and noise caused by combustion instability, and extends the service life of the engine.

[0069] (3) Increased operating range: The adjustable guide vane array can dynamically adjust the controllable airflow distribution ratio into the combustion and cooling channels according to different operating states of the turbojet engine, increasing the operating range of the afterburner. Combined with precise control of multiple turbojet engines, it achieves flexible matching of the power system in a wide speed and airspace range, enabling the engine to adapt to different flight mission requirements and improving the engine's combat effectiveness and adaptability. For example, it provides high thrust during takeoff acceleration and reduces heat load and fuel consumption during high-speed cruise; it can balance torque disturbances in real time to improve flight stability; it can achieve rapid switching of operating modes to meet the needs of different combat scenarios; it generates thrust vectors to enhance aircraft maneuverability; and it automatically balances power output in the event of single-engine failure to ensure flight safety.

[0070] Based on this, the present disclosure also provides a control method for the multi-engine integrated afterburner system provided in the present disclosure. Figure 6 A flowchart illustrating a control method for a multi-engine integrated afterburner system provided in an exemplary embodiment of this disclosure is shown. Figure 6 As shown, the control method for this multi-engine integrated afterburner system includes: S601, during the operation of the aircraft, uses a flow field homogenization component to mix and homogenize the exhaust airflow of the turbojet engine array that converges in the inlet region of the afterburner. S602 utilizes a flow distribution component to distribute the mixed and homogenized exhaust airflow to the combustion and cooling channels of the afterburner according to a controllable airflow distribution ratio; wherein, the controllable airflow distribution ratio is adapted to the operating state of the turbojet engine array and the flight state of the aircraft.

[0071] Specifically, as mentioned above, in a multi-engine integrated afterburner system, the exhaust gas from the turbojet engine array converges in the inlet area of ​​the afterburner, receiving the exhaust gas from the turbojet engine array in a shared afterburner configuration. This significantly reduces the wall area requiring active thermal protection and the cooling area of ​​the afterburner, structurally lowering the thermal load on the afterburner during long-term operation. This solves the thermal protection problem of afterburners in high-Mach number, wide-speed-range engines under long-term operation, ensuring the engine's operational stability requirements. At the same time, it eliminates the need for an additional fuel heat sink, simplifying the engine's thermal management system and improving the overall efficiency of the engine.

[0072] When the exhaust gas from the turbojet engine array enters the flow field control device in the inlet region of the afterburner, the exhaust gas from the turbojet engine array is mixed and homogenized by the flow field homogenization component. This can improve the combustion instability caused by the distortion of the exhaust flow field of multiple engines, ensure the performance and reliability of the turbojet engine array, and create favorable conditions for stable combustion in the afterburner.

[0073] The flow distribution component distributes the exhaust gas from the homogenization treatment to the combustion channel and cooling channel according to a controllable flow distribution ratio. This controllable flow distribution ratio is adapted to the operating state of the turbojet engine array and the flight state of the aircraft, realizing flexible matching of the power system in a wide speed range and a wide airspace, and improving the performance and reliability of the engine.

[0074] According to the technical solution provided by the exemplary embodiments of this disclosure, during the operation of the aircraft, the exhaust airflow of the turbojet engine array converging in the inlet region of the afterburner is mixed and homogenized using a flow field homogenization component; and the mixed and homogenized exhaust airflow is distributed to the combustion channel and cooling channel of the afterburner according to a controllable airflow distribution ratio using a flow distribution component. The controllable airflow distribution ratio is adapted to the operating state of the turbojet engine array and the flight state of the aircraft, effectively solving the thermal protection problem of the afterburner during long-term operation and the combustion instability problem caused by multiple engines in parallel, providing efficient and reliable propulsion power for high Mach number wide-speed range aircraft.

[0075] In some embodiments, the method may further include: The distributed control system is used to acquire the operating status of the turbojet engine array and the flight status of the aircraft. Generate collaborative control commands based on operational and flight status; By using coordinated control commands, the operating status of the turbojet engine array and the operating parameters of the flow distribution components are adjusted synchronously to control the controllable airflow distribution ratio.

[0076] Specifically, the distributed control system employs blockchain-based redundant control technology to achieve high-precision control of the characteristics of the four engines in the turbojet engine array, as well as the adjustable guide vane array. The decentralized ledger, automatic execution of smart contracts, and tamper-proof data storage of blockchain technology make it a solution for enhancing multi-engine collaborative control and redundancy management.

[0077] When the operating state of the turbojet engine array and the flight state of the aircraft are different, the required controllable airflow distribution ratio is different. Based on this, the embodiments of this disclosure can add a distributed control system. The distributed control system can be used to acquire the operating state of the turbojet engine array and the flight state of the aircraft in real time; generate cooperative control commands based on the operating state and flight state; and use the cooperative control commands to synchronously adjust the operating state of the turbojet engine array and the operating parameters of the flow distribution component to control the controllable airflow distribution ratio.

[0078] This distributed control system acquires global state information (the operating status of the turbojet engine array and the flight status of the aircraft) and generates a single, coordinated control command. It adjusts the engine's operating parameters and the opening of each blade in the adjustable guide vane array in real time, achieving flexible matching of the power system across a wide speed and airspace range. This includes real-time automatic compensation for torque disturbances, rapid switching between operating modes, thrust vector generation, and automatic reconfiguration of thrust distribution strategies to balance power output in the event of a single engine failure. This ensures a high degree of synchronization between engine thrust adjustment and airflow distribution adjustment. This distributed control system integrates the engine (thrust source) and flow distribution components (energy management valves) into a unified "power energy management system," achieving coordinated control of the turbojet engine array and flow distribution components. It achieves global optimization from the system's top level, rather than local optimization.

[0079] The above description is merely an illustration of some embodiments of this disclosure and the technical principles employed. Those skilled in the art should understand that the scope of this disclosure is not limited to technical solutions formed by specific combinations of the above-described technical features, but should also cover other technical solutions formed by arbitrary combinations of the above-described technical features or their equivalents without departing from the above-described concept. For example, technical solutions formed by substituting the above features with (but not limited to) technical features disclosed in this disclosure that have similar functions.

[0080] While specific embodiments of this disclosure have been described in detail by way of example, those skilled in the art should understand that the examples are for illustrative purposes only and not intended to limit the scope of this disclosure. Those skilled in the art should understand that modifications can be made to the above embodiments without departing from the scope and spirit of this disclosure. The scope of this disclosure is defined by the appended claims.

Claims

1. A multi-engine integrated afterburner system, characterized in that, The system includes a turbojet engine array, a flow field control device, and an afterburner arranged sequentially along the airflow direction. The flow field control device is detachably connected to the inlet region of the afterburner. The flow field control device includes a flow field homogenization component and a flow distribution component. The flow distribution component is movably connected to the end of the flow field homogenization component away from the inlet region. The interior of the afterburner has a combustion channel and a cooling channel. During the operation of the aircraft, the exhaust airflow of the turbojet engine array converges in the inlet region of the afterburner. After being mixed and homogenized by the flow field homogenization component, the exhaust airflow is distributed to the combustion channel and the cooling channel by the flow distribution component according to a controllable airflow distribution ratio. The controllable airflow distribution ratio is adapted to the operating state of the turbojet engine array and the flight state of the aircraft.

2. The system as described in claim 1, characterized in that, The flow field homogenization component includes multiple guide ribs that intersect around the axis of the afterburner, which divide the inlet region into multiple independent sub-channels.

3. The system as described in claim 2, characterized in that, The flow field homogenization component also includes multiple vortex generator arrays, each of which is fixedly disposed on one end of each of the guide ribs near the inlet region, for inducing the generation of enhanced mixing vortex structures in the sub-channel.

4. The system as described in claim 2, characterized in that, The flow distribution component includes an adjustable guide vane array, wherein each vane in the adjustable guide vane array is rotatably disposed at the end of each guide rib away from the inlet region, and the controllable airflow distribution ratio is controlled by adjusting the actuation parameters of each vane.

5. The system as described in claim 1, characterized in that, The interior of the afterburner has a cylindrical baffle arranged along the axial direction, the combustion channel is the internal region of the cylindrical baffle, and the cooling channel is the annular region inside the afterburner surrounding the outside of the cylindrical baffle. The cylindrical baffle is made of high-temperature resistant composite material. The inner wall surface of the cylindrical baffle facing the combustion channel is made of silicon carbide fiber-reinforced silicon carbide ceramic matrix composite material, and the outer wall surface of the cylindrical baffle facing the cooling channel is made of high-temperature alloy or carbon fiber-reinforced silicon carbide composite material.

6. The system as described in any one of claims 1 to 5, characterized in that, The system also includes: A distributed control system is configured to acquire the operating status of the turbojet engine array and the flight status of the aircraft; generate cooperative control commands based on the operating status and the flight status; and use the cooperative control commands to synchronously adjust the operating status of the turbojet engine array and the actuation parameters of the flow distribution component to control the controllable airflow distribution ratio.

7. The system as described in claim 6, characterized in that, The distributed control system is a redundant and fault-tolerant control system based on a blockchain architecture, including: An edge execution layer includes sensors, controllers, and actuators deployed on each of the turbojet engines in the turbojet engine array. The sensors are configured to acquire the operating status of each turbojet engine in real time, and the controllers are configured to preprocess the operating status of each turbojet engine. The airborne blockchain layer includes a blockchain network consisting of blockchain nodes deployed on each of the turbojet engines. The blockchain nodes are configured to generate control commands based on the operating state and the flight state. The blockchain network is configured to perform consensus processing on the control commands of multiple blockchain nodes through a blockchain consensus protocol to generate the collaborative control commands. The controller is also configured to adjust the operating state of the turbojet engine array using the cooperative control commands; and to send control command signals generated based on the cooperative control commands to the actuators. The actuator is configured to receive the control command signal and adjust the operating state of the flow distribution component based on the control command signal.

8. The system as described in claim 7, characterized in that, The blockchain network is also configured to, if one of the turbojet engines in the turbojet engine array fails, process the control commands of the other blockchain nodes through a blockchain consensus protocol to generate the collaborative control commands.

9. A control method for a multi-engine integrated afterburner system as described in any one of claims 1 to 8, characterized in that, include: During the operation of the aircraft, the exhaust airflow of the turbojet engine array that converges in the inlet region of the afterburner is mixed and homogenized using flow field homogenization components. Using the flow distribution component, the exhaust airflow after mixing and homogenization is distributed to the combustion channel and cooling channel of the afterburner according to a controllable airflow distribution ratio; wherein, the controllable airflow distribution ratio is adapted to the operating state of the turbojet engine array and the flight state of the aircraft.

10. The method as described in claim 9, characterized in that, The method further includes: The operating status of the turbojet engine array and the flight status of the aircraft are obtained using a distributed control system. Generate cooperative control commands based on the operating state and the flight state; The operating state of the turbojet engine array and the operating parameters of the flow distribution component are synchronously adjusted using the coordinated control commands to control the controllable airflow distribution ratio.

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