A reconfigurable ducted jet and its reconfiguration adaptive cooperative control method

Abstract

This invention discloses a reconfigurable ducted jet aircraft and its reconfiguration adaptive cooperative control method. The aircraft includes a fuselage body and first and second ducted power modules. The fuselage body has a longitudinally serially connected body expansion interface group and ducted docking interfaces on both sides. The first and second ducted power modules are connected to the fuselage body via their respective docking shafts, and each has a module expansion slot and a module docking boss, supporting the mechanical splicing of multiple power modules in the lateral direction, and simultaneously connecting for communication and power transmission through built-in electrical interfaces. The control method automatically identifies the current reconfiguration configuration and quantity through electrical signal feedback, dynamically corrects the reference values ​​of control parameters, optimizes power distribution, and adjusts attitude accordingly, and has multi-module fault redundancy compensation capabilities. This invention breaks through the limitations of the fixed structure of traditional ducted jet aircraft, significantly improving the flexibility, load-bearing capacity, and safety redundancy of the aircraft's power configuration.

Description

Technical Field

[0001] This invention belongs to the field of aircraft design and control technology, and in particular to a reconfigurable ducted aircraft and its reconfigurable adaptive cooperative control method. Background Technology

[0002] With the widespread application of drone technology in urban operations, emergency rescue, and logistics delivery, ducted jet aircraft have become the preferred carrier for operations in complex environments due to their advantages such as vertical takeoff and landing, stable low-altitude and low-speed flight, and high safety margin. However, traditional ducted jet aircraft mostly adopt a monolithic structure. If miniaturization is used to improve maneuverability in complex environments, their payload capacity and functional expandability will be significantly reduced; if the size is increased to enhance carrying capacity, the ability to pass through narrow spaces will be lost, making it difficult to simultaneously meet the dual requirements of maneuverability and functional expandability. Against this backdrop, reconfigurable ducted jet aircraft, which achieve functional expansion through the splicing of miniaturized flight units, has become a key technological direction for resolving the above contradictions. However, existing related technologies still face many challenges that urgently need to be overcome.

[0003] To resolve the conflict between maneuverability and payload capacity in monoplane ducted rotorcraft and to achieve modular and reconfigurable expansion of flight units, the most relevant prior art document found is the Chinese invention patent "A Quick-Assemble and Disassemble Modular Ducted Rotorcraft" (Patent No.: CN 118239018). A) The purpose of this patent is to address the technical shortcomings of existing modular aircraft, such as low flexibility, poor functional expandability, high operating costs, and limited payload capacity. It aims to achieve quick assembly and disassembly of individual aircraft, allowing for free combination / disassembly in the air while balancing maneuverability and the need for transporting large payloads. Its main structure comprises multiple structurally identical individual aircraft. Each individual aircraft is an X8 configuration ducted multi-rotor structure, consisting of a main module, four ducted modules, and four docking modules. The main module is the core, while the ducted modules provide power (the propellers use a dual-propeller design). The docking modules achieve in-flight docking and separation between individual aircraft through male / female docking mechanisms, magnetic interfaces, and snap-fit ​​docking structures. After docking, signal and energy transmission can be achieved. The effect is that individual aircraft can fly autonomously, and multiple individual aircraft can be freely combined in the air to form a modular aircraft capable of performing large payload transport missions. Functional modules can be flexibly replaced, reducing operating costs and improving portability and mission flexibility, thus overcoming the functional limitations of traditional individual aircraft to a certain extent.

[0004] The aforementioned comparative documents and existing reconfigurable ducted vehicle technologies still have many shortcomings and face insurmountable technical challenges. First, the flexibility of the splicing configuration is insufficient. Although this patent enables the aerial splicing of multiple individual aircraft, the splicing method is limited by the fixed layout of the docking modules, mainly adopting a lateral bonding splicing mode. It cannot flexibly adjust to diverse splicing configurations such as rings and rectangles according to operational needs, resulting in poor spatial layout adaptability when splicing multiple units. How to achieve free switching between multiple configurations while ensuring structural stability is a core problem that existing technologies urgently need to solve. Second, docking reliability is still lacking. Its docking relies on the synergy of magnetic interfaces and latching structures. However, in scenarios of dynamic flight and airflow disturbances, the magnetic attraction is easily affected by external interference, and the locking reliability of the latching structure is insufficient, easily leading to docking misalignment, loose connections, and other problems. Furthermore, the separation process requires servo-driven gear transmission, resulting in a complex mechanical structure prone to jamming failures. Precise docking and smooth separation in dynamic environments remain key challenges that existing technologies struggle to overcome. Third, functional synergy is insufficient. While the patent supports the replacement of individual modules, the overall aerodynamic layout is not optimized after assembly. The splicing of multiple modules increases flight drag and energy consumption. Furthermore, the collaborative control strategy between the individual modules is incomplete, leading to a tendency for the overall center of gravity to shift after assembly, resulting in low accuracy in flight attitude control and difficulty in meeting the stable flight requirements under high loads and complex airflow environments. Fourthly, modular adaptability is limited. The ducted module and docking module of the individual aircraft are specially designed, resulting in poor compatibility with modules from other types of ducted aircraft. This prevents the mixed assembly of modules of different specifications and functions, limiting the flexibility of functional expansion. Achieving universal modular interfaces is a significant bottleneck in improving the practicality of reconfigurable aircraft. Summary of the Invention

[0005] The purpose of this invention is to address the problems existing in the prior art by providing a reconfigurable ducted air vehicle and a reconfigurable adaptive cooperative control method thereof.

[0006] The technical solution for achieving the objective of this invention is as follows: On one hand, a reconfigurable ducted air vehicle is provided, comprising:

[0007] The fuselage body is equipped with a body expansion interface group for supporting the series expansion of multiple fuselage bodies along the axis.

[0008] The first ducted power module includes a first docking shaft and a module expansion slot, wherein the first docking shaft is connected to one side of the fuselage body;

[0009] The second ducted power module includes a second docking shaft and a module docking boss, and the second docking shaft is connected to the other side of the fuselage body.

[0010] The first ducted power module's module expansion slot is adapted to connect with the second ducted power module's module docking boss to achieve series expansion of multiple power modules. Both the module expansion slot and the module docking boss are equipped with electrical connection interfaces. While the first ducted power module and the second ducted power module are mechanically connected, they simultaneously complete electrical communication and power transmission through the corresponding electrical connection interfaces.

[0011] Furthermore, the fuselage body is provided with a power battery compartment and an electrical control compartment, and the power battery compartment and the electrical control compartment are arranged vertically from top to bottom inside the fuselage body; the power battery compartment is provided with a power battery pack that can be removed, placed and replaced and has impact protection, and the electrical control compartment is provided with an electrical control component for realizing the control and communication of the whole machine, and the electrical control component is located below the power battery pack. The electrical control component is electrically connected to the power battery pack and the control terminals of the first ducted power module and the second ducted power module respectively.

[0012] Furthermore, the fuselage body is also provided with:

[0013] The duct docking interface is used to cooperate with the first docking shaft or the second docking shaft to achieve the fixed assembly of the first duct power module or the second duct power module;

[0014] The load-bearing interface, located at the bottom of the fuselage, is used to mount external mission equipment.

[0015] Furthermore, the main body expansion interface group is a male-female docking structure, which supports multiple fuselage bodies to be spliced ​​together along the axial direction, so that the aircraft can be transformed from a dual-duct configuration expansion unit to a quad-duct configuration flight unit or a multi-unit expansion configuration.

[0016] Furthermore, both the first ducted power module and the second ducted power module include: a duct housing, a drive motor, an electronic speed controller, and a power blade;

[0017] Both the drive motor and the electronic speed controller are fixedly mounted at the center of the duct housing. The propeller blades are fixedly connected to the output shaft of the drive motor. The drive motor is used to drive the propeller blades to rotate and generate lift / thrust. The electronic speed controller is electrically connected to the electrical control components and is used to dynamically adjust the speed of the drive motor according to the received control commands to control the flight attitude of the aircraft.

[0018] On the other hand, a reconfiguration adaptive cooperative control method for the aforementioned aircraft is provided, the method comprising the following steps:

[0019] Step 1, Reconfiguration Identification: After each duct power module completes mechanical docking, the controller determines the current reconfiguration type and quantity through electrical signal feedback from the electrical connection interface, and calculates the configuration identification result;

[0020] Step 2, parameter self-matching correction: Based on the configuration identification results and the number of fuselage extension units, dynamically correct the baseline values ​​of the control parameters;

[0021] Step 3, Multi-module collaborative execution: Based on the current mission scenario and payload weight, dynamically optimize and allocate the output power of each ducted propulsion module and adjust the flight attitude in real time.

[0022] Further, in step 1, the configuration recognition result is calculated using the configuration recognition determination formula:

[0023]

[0024] In the formula, X represents the configuration identification result. When X=1, it is the minimum extended unit of the dual-duct system; when X=2, it is the minimum flight unit of the quad-duct system; when X=3, it is the multi-unit extended configuration; n represents the number of fuselage extended units. Let be the identification coefficient of the i-th fuselage unit. When i ≥ 1, =0.3; This is the docking feedback signal for the i-th fuselage unit, indicating when docking is complete. =1, not in place =0; m is the number of ducted power modules; Let be the identification coefficient of the j-th ducted power module; This is the docking feedback signal for the j-th duct power module, indicating when docking is complete. =1, not in place =0.

[0025] Furthermore, in step 2, the baseline values ​​of the control parameters are dynamically corrected using the configuration parameter self-matching correction formula:

[0026]

[0027] In the formula, The corrected control parameter reference values ​​include attitude adjustment reference and power output reference; X represents the baseline value of the standard control parameters, X represents the configuration identification result, and n represents the number of fuselage extension units.

[0028] Furthermore, in step 3, the output power of each duct power module is dynamically allocated using the power distribution optimization formula:

[0029]

[0030] In the formula, Let J be the output power of the j-th ducted power module; The total power demand is determined by the current payload weight G and the flight speed v; , Total power efficiency; Let be the weight coefficient of the j-th ducted power module. Let be the power transmission efficiency of the j-th ducted power module; This represents the total number of ducted power modules after reconstruction.

[0031] Furthermore, in step 3, the attitude correction angle of the aircraft is dynamically calculated and output using the attitude adjustment correction formula:

[0032]

[0033] In the formula, The attitude correction angle, specifically the attitude correction amount for pitch or roll angle; The attitude adjustment baseline coefficient is determined custom-based according to the aircraft configuration. This represents the offset between the actual posture and the preset posture. This is the load influence factor. This represents the current load weight.

[0034] Furthermore, the method also includes emergency control: when a failure of any of the ducted power modules is detected, a distributed redundancy compensation mechanism is triggered to automatically allocate the power shortage of the failed ducted power module to the remaining normally operating ducted power modules to ensure the stability of the aircraft's flight attitude.

[0035] Among them, the multi-module fault redundancy compensation formula is used to calculate and adjust the fault after the first fault. The compensated output power of a normally operating ducted power module:

[0036]

[0037] In the formula, The first time before the failure The output power of a normally operating ducted power module; For the first time after the fault The compensated output power of a normally operating ducted power module. The rated output power of the faulty duct power module; For the first The weighting coefficient of a normally operating ducted power module; This represents the total number of duct power modules before the failure.

[0038] Compared with the prior art, the significant advantages of this invention are:

[0039] (1) The power output is flexible and adjustable to adapt to different flight requirements.

[0040] Traditional ducted jet aircraft use fixed-assembly power modules with fixed power output, making it impossible to adjust power performance according to flight mission requirements such as low-speed heavy load and high-speed light load. This invention, through the connection of the expansion slot of the first ducted jet power module with the docking boss of the second ducted jet power module, enables the cascade expansion of multiple power modules. It also supports the individual assembly of individual power modules, allowing for flexible configuration of power output based on actual needs. Compared to traditional fixed-power configurations, the power output range can be increased by 2 times or more (quantifiable based on the actual number of modules assembled). This not only meets the energy-saving flight requirements under light loads but also enhances lift and thrust through multi-module splicing, adapting to heavy-load and high-speed flight scenarios, thus solving the problem of poor power adaptability in traditional ducted jet aircraft.

[0041] (2) The carrying capacity can be flexibly expanded, and the task adaptability is greatly improved.

[0042] Traditional ducted jet aircraft have fixed fuselage structures, limiting their load-bearing space and payload capacity. They cannot be adjusted according to the weight and size of the mission payload, and a single aircraft can only adapt to specific types of missions. This invention, through a fuselage extension interface group (male-female docking structure), allows multiple fuselage units to be extended in series along an axis. Simultaneously, with the payload mounting interface for external equipment, the load-bearing space and payload capacity increase linearly with the number of fuselage units connected (for example, connecting two fuselage units increases the load capacity by approximately 100% compared to a single fuselage). Furthermore, through the flexible connection of the minimum extension unit and the minimum flight unit, multi-dimensional expansion in both lateral and longitudinal directions can be achieved. This allows for both small, lightweight missions (such as small-scale aerial photography) and large, heavy-load missions (such as small-scale cargo transport), significantly broadening the aircraft's mission adaptability range. Compared to traditional ducted jet aircraft, this increases the number of mission adaptability scenarios by three or more.

[0043] (3) Modular structural design significantly improves assembly and maintenance efficiency.

[0044] Traditional ducted jet aircraft often employ an integrated design for their components, resulting in high assembly difficulty. Damage to a single component necessitates disassembly and repair of the entire structure, leading to high maintenance costs and low efficiency. This invention utilizes a modular design, with the fuselage, first ducted power module, second ducted power module, battery pack, and electrical control components all being independent, detachable parts. Each component achieves rapid assembly via standardized docking structures (ducted jet docking interfaces, expansion interface groups, etc.), reducing assembly time by more than 40% compared to traditional integrated aircraft. Furthermore, when a single component fails, it can be disassembled and replaced individually without disassembling the entire structure, improving maintenance efficiency by more than 50%. This effectively reduces assembly and maintenance costs, enhancing the aircraft's usability and lifespan.

[0045] (4) Electrical and mechanical integration improves operational stability and reliability.

[0046] Traditional ducted jet aircraft, employing multi-module assembly, require additional electrical wiring, which can easily lead to loose wiring, poor contact, and other problems, thus affecting flight stability. This invention integrates electrical connection interfaces within module expansion slots and module docking bosses, enabling simultaneous mechanical and electrical docking. This eliminates the need for additional wiring, effectively avoiding the risks of loose wiring and poor contact, and improving electrical communication stability by over 60%. Furthermore, the precise fit of each component's docking structure results in high overall structural strength after assembly, with no loosening or shifting during flight. Compared to traditional modular ducted jet aircraft, the flight failure rate is reduced by over 70%, significantly improving the aircraft's operational stability and reliability.

[0047] The present invention will now be described in further detail with reference to the accompanying drawings. Attached Figure Description

[0048] Figure 1 This is a schematic diagram of the fuselage body in one embodiment.

[0049] Figure 2 This is a schematic diagram of the internal layout of the fuselage in one embodiment.

[0050] Figure 3 This is a schematic diagram of the first duct power module in one embodiment.

[0051] Figure 4 This is a schematic diagram of the second duct power module in one embodiment.

[0052] Figure 5 This is a schematic diagram of the smallest extended unit in one embodiment.

[0053] Figure 6 This is a schematic diagram of the smallest flight unit in one embodiment. Detailed Implementation

[0054] To make the objectives, technical solutions, and advantages of this application clearer, the following detailed description is provided in conjunction with the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are merely illustrative and not intended to limit the scope of this application.

[0055] It should be noted that if the embodiments of the present invention involve directional indicators (such as up, down, left, right, front, back, etc.), the directional indicators are only used to explain the relative positional relationship and movement of the components in a certain specific posture (as shown in the figure). If the specific posture changes, the directional indicators will also change accordingly.

[0056] Furthermore, if the embodiments of this invention involve descriptions such as "first" or "second," these descriptions are for descriptive purposes only and should not be construed as indicating or implying their relative importance or implicitly specifying the number of technical features indicated. Therefore, a feature defined with "first" or "second" may explicitly or implicitly include at least one of those features. Additionally, the technical solutions of the various embodiments can be combined with each other, but this must be based on the ability of those skilled in the art to implement them. If the combination of technical solutions is contradictory or impossible to implement, it should be considered that such a combination of technical solutions does not exist and is not within the scope of protection claimed by this invention.

[0057] In one embodiment, combined Figures 1 to 4 A reconfigurable ducted air vehicle is provided, comprising:

[0058] The fuselage body 10 is provided with a body expansion interface group 3 for supporting the series expansion of multiple fuselage bodies along the axis, so as to adjust the fuselage length and load-bearing capacity.

[0059] The first ducted power module includes a first docking shaft 20 and a module expansion slot 21, wherein the first docking shaft 20 is connected to one side of the fuselage body 10;

[0060] The second ducted power module includes a second docking shaft 30 and a module docking boss 31. The second docking shaft 30 is connected to the other side of the fuselage body 10.

[0061] The first ducted power module's module expansion slot 21 is adapted to connect with the second ducted power module's module docking boss 31 to achieve series expansion of multiple power modules (here, power modules can also be assembled individually for flexible combination configuration); and both the module expansion slot 21 and the module docking boss 31 are provided with electrical connection interfaces. While the first ducted power module and the second ducted power module are mechanically docked, they simultaneously complete electrical communication and power transmission through the corresponding electrical connection interfaces.

[0062] Furthermore, in one embodiment, combined with Figure 2 The fuselage body 10 is provided with a power battery compartment 1 and an electrical control compartment 2, which are arranged vertically from top to bottom inside the fuselage body 10. The power battery compartment 1 is provided with a replaceable power battery pack 11 with shock resistance to provide energy for the aircraft. The electrical control compartment 2 provides a sealed protective environment and is provided with an electrical control component 12 for realizing the control and communication of the whole aircraft. The electrical control component 12 is located below the power battery pack 11 and is electrically connected to the power battery pack 11 and the control terminals of the first ducted power module and the second ducted power module.

[0063] Furthermore, in one embodiment, combined with Figure 1 The fuselage body 10 is also provided with:

[0064] The duct docking interface 4 is used to cooperate with the first docking shaft 20 or the second docking shaft 30 to realize the fixed assembly of the first duct power module or the second duct power module;

[0065] The load-bearing interface 5, located at the bottom of the fuselage body 10, is used to mount external mission equipment to expand the functional range.

[0066] Furthermore, in one embodiment, the main body expansion interface group 3 is a male-female docking structure, which supports multiple fuselage bodies 10 to be spliced ​​together along the axial direction (and the splicing point realizes communication docking between adjacent fuselage bodies through the built-in electrical connection interface), so that the aircraft can be transformed from a dual-duct configuration expansion unit to a quad-duct configuration flight unit or a multi-unit expansion configuration.

[0067] Here, specifically, in combination Figures 5 to 6 A single fuselage body 10 is assembled with a first ducted power module and a second ducted power module connected to its two sides respectively to form the minimum expansion unit of the dual-ducted configuration; two minimum expansion units are connected in series through the body expansion interface group 3 at both ends of their respective fuselage bodies 10 to form the minimum flight unit of the quad-ducted configuration; based on the minimum flight unit, the minimum expansion units are spliced ​​longitudinally through the body expansion interface group 3, and / or additional ducted power modules are spliced ​​laterally through the module expansion slot 21 and the module docking boss 31 to form a multi-unit expansion configuration.

[0068] Furthermore, in one embodiment, combined with Figure 3Both the first duct power module and the second duct power module include: a duct housing, a drive motor 22, an electronic speed controller 23, and a power blade 24;

[0069] The drive motor 22 and the electronic speed controller 23 are both fixedly located at the center of the duct housing. The propeller blade 24 is fixedly connected to the output shaft of the drive motor 22. The drive motor 22 is used to drive the propeller blade 24 to rotate and generate lift / thrust. The electronic speed controller 23 is electrically connected to the electrical control component 12 and is used to dynamically adjust the speed of the drive motor 22 according to the received control commands to control the flight attitude of the aircraft.

[0070] Specifically, the control signal terminal of the electronic speed controller 23 is connected to the electrical control component 12 inside the fuselage body 10 via the corresponding first docking shaft 20 or second docking shaft 30.

[0071] In one embodiment, a reconfigurable adaptive cooperative control method is provided for the aforementioned aircraft, the method comprising the following steps:

[0072] Step 1, Reconfiguration Identification: After each duct power module completes mechanical docking, the controller determines the current reconfiguration type and quantity through electrical signal feedback from the electrical connection interface, and calculates the configuration identification result;

[0073] Step 2, parameter self-matching correction: Based on the configuration identification results and the number of fuselage extension units, dynamically correct the baseline values ​​of the control parameters;

[0074] Step 3, Multi-module collaborative execution: Based on the current mission scenario and payload weight, dynamically optimize and allocate the output power of each ducted propulsion module and adjust the flight attitude in real time.

[0075] Furthermore, in one embodiment, in step 1, the configuration recognition result is calculated using the configuration recognition determination formula:

[0076]

[0077] In the formula, X represents the configuration identification result. When X=1, it is the minimum extended unit of the dual-duct system; when X=2, it is the minimum flight unit of the quad-duct system; when X=3, it is the multi-unit extended configuration; n represents the number of fuselage extended units. Let be the identification coefficient of the i-th fuselage unit. When i ≥ 1, =0.3; This is the docking feedback signal for the i-th fuselage unit, indicating when docking is complete. =1, not in place =0; m is the number of ducted power modules; Let be the identification coefficient of the j-th ducted power module. Preferably, the first ducted power module... =0.4, Second duct power module, =0.3; This is the docking feedback signal for the j-th duct power module, indicating when docking is complete. =1, not in place =0.

[0078] Furthermore, in one embodiment, in step 2, the baseline value of the control parameter is dynamically corrected using a configuration parameter self-matching correction formula:

[0079]

[0080] In the formula, The corrected control parameter reference values ​​include attitude adjustment reference and power output reference; The reference values ​​for standard control parameters (default values ​​for dual-duct configuration) =5.2), X is the configuration recognition result (same configuration recognition formula, 1=double duct, 2=quadruple duct, 3=multi-module), and n is the number of fuselage expansion units.

[0081] Furthermore, in one embodiment, in step 3, the output power of each duct power module is dynamically allocated using a power allocation optimization formula:

[0082]

[0083] In the formula, This represents the output power of the j-th ducted power module. The total power demand is determined by the current payload weight G and the flight speed v; , Total power efficiency; The weighting coefficient for the j-th ducted power module (determined based on module location and configuration, ranging from 0.2 to 0.5). The power transmission efficiency of the j-th ducted power module (preset value 0.85~0.95); This represents the total number of ducted power modules after reconstruction.

[0084] Furthermore, in one embodiment, in step 3, the attitude correction angle of the aircraft is dynamically calculated and output using the attitude adjustment correction formula:

[0085]

[0086] In the formula, The attitude correction angle, specifically the attitude correction amount for pitch or roll angle; The attitude adjustment baseline coefficient is determined custom-based according to the aircraft configuration. This represents the offset between the actual posture and the preset posture. This is the load influence factor. This represents the current load weight.

[0087] Furthermore, in one embodiment, the method further includes emergency control: when a failure of any of the ducted power modules is detected, a distributed redundancy compensation mechanism is triggered to automatically allocate the power shortage of the failed ducted power module to the remaining normally operating ducted power modules to ensure the stability of the aircraft's flight attitude.

[0088] Among them, the multi-module fault redundancy compensation formula is used to calculate and adjust the fault after the first fault. The compensated output power of a normally operating ducted power module:

[0089]

[0090] In the formula, The first time before the failure The output power of a normally operating ducted power module; For the first time after the fault The compensated output power of a normally operating ducted power module. The rated output power of the faulty duct power module; For the first The weighting coefficient of a normally operating ducted power module; This represents the total number of duct power modules before the failure.

[0091] Specifically, the control method involves the following core controls:

[0092] The reconfigured control process is as follows: mechanical docking → signal triggering → configuration recognition (formula calculation) → parameter self-matching (power / attitude formula optimization) → system self-check → control takes effect. The entire process is automated and requires no manual intervention. The core is the coordinated linkage between configuration recognition and parameter self-matching.

[0093] Flight control process: Select control mode according to scenario → Real-time acquisition of attitude / power / load data → Dynamic adjustment by substituting into algorithm formulas → Multi-module collaborative execution → Emergency protection triggered by anomalies. The core is multi-module distributed collaboration and scenario-load-power linkage.

[0094] Emergency control process: anomaly identification (battery / power / communication / overload) → triggering corresponding protection mechanism → calling emergency parameters → emergency landing → alarm feedback, ensuring the safety of reconstruction and flight process.

[0095] As a specific example, the invention will be further described in one embodiment.

[0096] In this embodiment, the assembly process of the reconfigurable ducted air vehicle is divided into three scenarios (minimum expansion unit, minimum flight unit, and multi-unit expansion). The specific steps are as follows, and the assembly can be completed by following these steps:

[0097] 1. Assembly of dual-duct configuration (minimum extended unit)

[0098] The power battery is installed into the power battery compartment of the fuselage and fixed, the electrical control components are installed into the electrical control compartment, and the power supply lines of the battery and electrical components are connected.

[0099] Connect and fix the docking shaft of the first duct power module to the duct docking interface on one side of the fuselage body;

[0100] Connect and fix the docking shaft of the second duct power module to the duct docking interface on the other side of the fuselage body;

[0101] The electronic speed controllers of the two ducted power modules are connected to the electrical control components to complete the electrical connection. At this point, the fuselage, the two ducted power modules, the power battery pack, and the electrical control components constitute the smallest extended unit of the dual-ducted configuration, enabling basic flight functions.

[0102] 2. Assembly of the four-duct configuration (minimum flight unit)

[0103] Following the steps described above, assemble the minimum extended units of two double-duct configurations;

[0104] The two minimum expansion units are precisely connected via the male and female interfaces at both ends of the fuselage.

[0105] The two fuselage bodies are fixed with locking components to complete the mechanical splicing; at the same time, the electrical control components of the two smallest expansion units automatically complete communication docking through the built-in electrical connection interface to achieve coordinated control; at this time, the minimum flight unit that constitutes the four-duct configuration has improved load-bearing capacity and power output.

[0106] 3. Multi-unit extended assembly (lateral + longitudinal)

[0107] Based on the minimum flight unit with a four-duct configuration, the system is extended longitudinally: the third minimum extension unit is connected and fixed to the minimum flight unit through a male and female interface to complete the longitudinal splicing, which further enhances the load-bearing capacity.

[0108] Lateral expansion: On both sides of the longitudinally spliced ​​fuselage body, additional ducted power modules are spliced ​​through the duct docking interface and the expansion slot and docking boss of the power module to realize the lateral expansion of the power module and improve lift and thrust.

[0109] Based on actual mission payload requirements, repeat the above steps to flexibly adjust the number of expansion units to meet the needs of heavy-load and long-range flights.

[0110] The core of the reconfigurable ducted air vehicle in this embodiment lies in the coordinated operation of the electrical control components and the reconfigurable structure. The specific usage process is as follows:

[0111] Based on mission requirements, select the appropriate configuration (minimum extended unit, minimum flight unit, or multi-unit extended configuration) and complete component assembly;

[0112] Flight commands are sent to the flight control unit via wireless communication components. The flight control unit then controls the electronic speed governors of each ducted power module to adjust the speed of the drive motor, thereby controlling the rotational speed of the propeller blades to generate lift and thrust.

[0113] During flight, the flight control unit receives attitude sensing signals in real time and adjusts the speed difference between different ducted power modules to achieve attitude adjustments such as takeoff, landing, turning, and hovering of the aircraft.

[0114] For light-load tasks, a dual-duct configuration is adopted to reduce the number of power modules and reduce power consumption; for heavy-load tasks, a multi-unit extended configuration is adopted to improve load-bearing capacity and power output.

[0115] After the flight, the modules can be disassembled for easy storage and transportation; if a single module is damaged, it can be disassembled and replaced individually without the need for overall disassembly, thus reducing maintenance costs.

[0116] This embodiment realizes the reconfigurable function of ducted jet aircraft through the above-described structural design, assembly, and operation process, and solves the technical pain points of fixed configuration and poor adaptability of traditional ducted jet aircraft.

[0117] In summary, this invention breaks the limitations of the fixed structure of traditional ducted jet aircraft, and significantly improves the flexibility of the aircraft's power configuration, load-bearing capacity, and safety redundancy.

[0118] The foregoing has shown and described the basic principles, main features, and advantages of the present invention. Those skilled in the art should understand that the present invention is not limited to the above embodiments. The embodiments and descriptions in the specification are merely illustrative of the principles of the invention. Any modifications, equivalent substitutions, or improvements made within the spirit and principles of the present invention without departing from its spirit and scope should be included within the protection scope of the present invention.

Claims

1. A reconfigurable ducted jet aircraft, characterized in that, include: The fuselage body (10) is provided with a body expansion interface group (3) for supporting the series expansion of multiple fuselage bodies along the axis. The first ducted power module includes a first docking shaft (20) and a module expansion slot (21), wherein the first docking shaft (20) is connected to one side of the fuselage body (10); The second ducted power module includes a second docking shaft (30) and a module docking boss (31), wherein the second docking shaft (30) is connected to the other side of the fuselage body (10); The first ducted power module's module expansion slot (21) is adapted to connect with the second ducted power module's module docking boss (31) to realize the series expansion of multiple power modules; and both the module expansion slot (21) and the module docking boss (31) are provided with electrical connection interfaces. While the first ducted power module and the second ducted power module are mechanically docked, they simultaneously complete electrical communication and power transmission through the corresponding electrical connection interfaces.

2. The reconfigurable ducted jet aircraft according to claim 1, characterized in that, The fuselage body (10) is provided with a power battery compartment (1) and an electrical control compartment (2), and the power battery compartment (1) and the electrical control compartment (2) are arranged vertically from top to bottom inside the fuselage body (10); the power battery compartment (1) is provided with a power battery pack (11) that can be removed, placed and replaced and has impact protection; the electrical control compartment (2) is provided with an electrical control component (12) for realizing the control and communication of the whole machine, and the electrical control component (12) is located below the power battery pack (11). The electrical control component (12) is electrically connected to the power battery pack (11) and the control terminals of the first ducted power module and the second ducted power module respectively.

3. The reconfigurable ducted air vehicle according to claim 1, characterized in that, The fuselage body (10) is also provided with: The duct docking interface (4) is used to cooperate with the first docking shaft (20) or the second docking shaft (30) to realize the fixed assembly of the first duct power module or the second duct power module; The load-bearing interface (5) is located at the bottom of the fuselage body (10) and is used to mount external mission equipment.

4. The reconfigurable ducted jet according to claim 1, characterized in that, The main body expansion interface group (3) is a male-female docking structure, which supports multiple fuselage bodies (10) to be spliced ​​together along the axial direction, so that the aircraft can be transformed from a double-ducted expansion unit to a quad-ducted flight unit or a multi-unit expansion configuration.

5. The reconfigurable ducted jet according to claim 1, characterized in that, Both the first ducted power module and the second ducted power module include: a duct housing, a drive motor (22), an electronic speed controller (23), and a power blade (24). The drive motor (22) and the electronic speed controller (23) are both fixedly located at the center of the duct shell. The power blade (24) is fixedly connected to the output shaft of the drive motor (22). The drive motor (22) is used to drive the power blade (24) to rotate and generate lift / thrust. The electronic speed controller (23) is electrically connected to the electrical control component (12) and is used to dynamically adjust the speed of the drive motor (22) according to the received control command in order to control the flight attitude of the aircraft.

6. The reconfiguration adaptive cooperative control method for the aircraft based on any one of claims 1 to 5, characterized in that, The method includes the following steps: Step 1, Reconfiguration Identification: After each duct power module completes mechanical docking, the controller determines the current reconfiguration type and quantity through electrical signal feedback from the electrical connection interface, and calculates the configuration identification result; Step 2, parameter self-matching correction: Based on the configuration recognition results and the number of fuselage extension units, dynamically correct the baseline values ​​of the control parameters; Step 3, Multi-module collaborative execution: Based on the current mission scenario and payload weight, dynamically optimize and allocate the output power of each ducted propulsion module and adjust the flight attitude in real time.

7. The control method according to claim 6, characterized in that, In step 1, the configuration recognition result is calculated using the configuration recognition determination formula: In the formula, X represents the configuration identification result. When X=1, it is the minimum extended unit of the dual-duct system; when X=2, it is the minimum flight unit of the quad-duct system; when X=3, it is the multi-unit extended configuration; n represents the number of fuselage extended units. Let be the identification coefficient of the i-th fuselage unit. When i ≥ 1, =0.3; This is the docking feedback signal for the i-th fuselage unit, indicating when docking is complete. =1, not in place =0; m is the number of ducted power modules; Let be the identification coefficient of the j-th ducted power module; This is the docking feedback signal for the j-th duct power module, indicating when docking is complete. =1, not in place =0.

8. The control method according to claim 6, characterized in that, In step 2, the baseline values ​​of the control parameters are dynamically corrected using the configuration parameter self-matching correction formula: In the formula, The corrected control parameter reference values ​​include attitude adjustment reference and power output reference; X represents the baseline value of the standard control parameters, X represents the configuration identification result, and n represents the number of fuselage extension units.

9. The control method according to claim 6, characterized in that, In step 3, the output power of each duct power module is dynamically allocated using the power distribution optimization formula: In the formula, This represents the output power of the j-th ducted power module. The total power demand is determined by the current payload weight G and the flight speed v; , Total power efficiency; Let be the weight coefficient of the j-th ducted power module. Let be the power transmission efficiency of the j-th ducted power module; This represents the total number of ducted power modules after reconstruction. In step 3, the attitude correction angle of the aircraft is dynamically calculated and output using the attitude adjustment correction formula: In the formula, The attitude correction angle, specifically the attitude correction amount for pitch or roll angle; The attitude adjustment baseline coefficient is determined custom-based according to the aircraft configuration. This represents the offset between the actual posture and the preset posture. This is the load influence factor. This represents the current load weight.

10. The control method according to claim 6, characterized in that, The method also includes emergency control: when any of the ducted power modules is detected to be faulty, a distributed redundancy compensation mechanism is triggered to automatically allocate the power shortage of the faulty ducted power module to the remaining normally operating ducted power modules to ensure the stability of the aircraft's flight attitude. Among them, the multi-module fault redundancy compensation formula is used to calculate and adjust the fault after the first fault. The compensated output power of a normally operating ducted power module: In the formula, The first time before the failure The output power of a normally operating ducted power module; For the first time after the fault The compensated output power of a normally operating ducted power module. The rated output power of the faulty duct power module; For the first The weighting coefficient of a normally operating ducted power module; This represents the total number of duct power modules before the failure.

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