A method and system for augmented stability flight control of an aircraft

Abstract

The application relates to the field of flight control technology and discloses a stability-increasing flight control method and system for an aircraft, which comprises the following steps: establishing an open-loop multi-port Hamilton dynamics model of the aircraft; obtaining real-time state variables and energy conjugate variables through a perception network; performing a passive energy shaping control algorithm based on interconnection and damping distribution to calculate a generalized control input variable group; decoupling the variable group into two-dimensional control instructions for driving variable impedance servo actuators, including a basic deflection angle instruction and a dynamic mechanical damping instruction; and respectively independently changing the deflection position of a micro control surface and the local hinge mechanical damping coefficient of each actuator according to the instructions. The application actively regulates the aeroelastic wave phase through time sequence deflection, directly dissipates high-frequency energy in combination with terminal damping, avoids saturation of a high-frequency instruction of an executing mechanism, and effectively improves the global stability and stability-increasing efficiency of the whole aircraft under complex disturbances.

Description

Technical Field

[0001] This invention relates to the field of flight control technology, specifically to a method and system for stabilizing flight control of an aircraft. Background Technology

[0002] High-aspect-ratio flexible fixed-wing aircraft are widely used in modern aviation due to their physical advantages such as high aerodynamic efficiency and long endurance. However, because of their slender wing structure and low overall stiffness, they are prone to large-scale localized aeroelastic deformation when encountering atmospheric turbulence or gusts during flight. This structural flexibility deformation under multi-physics environment has a strong nonlinear coupling effect with the large-scale rigid body motion of the airframe, which not only seriously interferes with the aircraft's attitude stability but also easily triggers aeroelastic vibration divergence, threatening the structural safety of the entire aircraft.

[0003] Existing aircraft stability augmentation control schemes typically design feedback control laws based on a locally linearized system model. Their core control logic involves using sensors to collect rigid body attitude signals and wing structure vibration signals. After control law calculation, a single deflection position command is output to the servo motor, relying entirely on the aerodynamic torque generated by the control surface deflection to suppress elastic vibrations and maintain flight attitude. However, this traditional control architecture has significant limitations when dealing with the complex nonlinear dynamic characteristics of large, flexible aircraft. Firstly, the single position servo control mechanism requires the control surface actuators to bear the entire control load. When high-frequency, wide-bandgap elastic traveling waves appear on the wing surface, the system must continuously issue high-frequency displacement compensation commands. This easily exceeds the physical response bandwidth limit of conventional mechanical actuators, leading to severe high-frequency command saturation, mechanical wear and overheating, and dynamic signal distortion in the actuators. Secondly, existing technologies lack intervention methods to directly dissipate high-frequency mechanical residual energy at the physical terminal and cannot effectively reshape the propagation phase of aeroelastic waves in the spatiotemporal dimension. In flight conditions with dense aeroelastic modes, it is difficult to achieve rapid convergence of vibration energy by relying solely on the inherent damping of the system and the aerodynamic damping caused by the control surfaces. Instead, the response lag of the actuation system can easily lead to phase mismatch between the control force and the structural elastic restoring force, thereby deteriorating the global stability and safety boundary of the flight control system when dealing with complex disturbances. Summary of the Invention

[0004] To address the shortcomings of existing technologies, this invention provides a flight control method and system for stabilizing aircraft, which solves the problem that traditional single control surface deflection control for large aspect ratio flexible fixed-wing aircraft under complex aerodynamic disturbances easily leads to high-frequency command saturation of the actuator and difficulty in quickly and controllably dissipating aeroelastic vibration energy, resulting in flight attitude instability.

[0005] To achieve the above objectives, the present invention provides a method for stabilizing flight control of an aircraft, executed by a flight control computer, the method comprising the following processing steps: An open-loop multi-port Hamiltonian dynamics model of the aircraft is established. The continuous motion of the aircraft is decomposed into a rigid body domain spatial motion subsystem and a structural elastic domain local deformation subsystem. Based on this, a generalized state vector is constructed, including the generalized coordinates of rigid body attitude, the generalized momentum of rigid body attitude, the generalized coordinates of structural flexible deformation, and the generalized momentum of structural flexible deformation. By combining the rigid body inertia matrix, the structural generalized mass matrix, and the structural generalized stiffness matrix, a total open-loop Hamiltonian energy function is constructed, and its derivative is used to generate an open-loop multi-port Hamiltonian dynamic differential equation that includes the system's internal interconnection matrix and intrinsic dissipation matrix.

[0006] Real-time state variables and energy conjugate variables are acquired through an energy state sensing network deployed on a high aspect ratio flexible fixed-wing aircraft. Three-axis attitude angles and angular velocities are collected using an inertial measurement unit; local structural curvature and strain parameters are collected using a fiber Bragg grating strain sensor array as generalized coordinate inputs to the structural elastic domain; and local transient aerodynamic load parameters are collected using a microelectromechanical dynamic surface pressure sensor array as generalized force inputs to the external aerodynamic domain ports.

[0007] A passive energy shaping control algorithm based on interconnection and damping allocation is implemented. The desired stable equilibrium point of the aircraft for the flight mission is defined. The total open-loop Hamiltonian energy function is superimposed with the actively injected virtual potential energy function to construct a closed-loop desired multi-port Hamiltonian energy function, which is then forced to satisfy the partial derivative constraint condition of the global minimum energy state at the desired stable equilibrium point. Simultaneously, a closed-loop desired dynamic differential equation including the desired interconnection matrix and the desired dissipation matrix is ​​established.

[0008] The generalized control input variable set is analyzed. By simultaneously solving the open-loop multiport Hamiltonian dynamic differential equation and the closed-loop expectation dynamic differential equation, a set of terms including energy gradients is extracted to construct a nonlinear algebraic matching equation. Considering the physical underactuation property where the total number of microcontroller surfaces of the aircraft is less than the total number of system degrees of freedom, the full-rank pseudo-inverse matrix of the input configuration matrix is ​​calculated. This full-rank pseudo-inverse matrix is ​​then used to perform a left product operation on the multidimensional gradient vectors, and the corresponding minimum norm solution is obtained in the control input subspace as the generalized control input variable set.

[0009] The generalized control input variable set is decoupled into two-dimensional control commands. The generalized control input variable set is split into active generalized force component vectors and damping generalized force component vectors. The damping generalized force component vector is used to generate dynamic mechanical damping commands. Specifically, a generalized damping force column vector is constructed using the control injection dissipation matrix and the desired energy gradient column vector. This column vector is then converted into a local desired controlled dissipation force column vector using a damping allocation mapping matrix. This vector is then divided sequentially by the corresponding elements in the physical deflection angular velocity column vectors of each micro-control surface. When the absolute value of the angular velocity is lower than a preset dead zone threshold constant, the division algebraic operation is interrupted, and the system's basic sustaining damping constant is output. The active generalized force component vector is used to generate basic deflection angle commands. Specifically, the physical propagation phase velocity parameters of the dominant aeroelastic deformation wave are extracted. The theoretical time delay constant is calculated by combining the spanwise geometric relative distance of each variable impedance servo actuator. This results in the construction of a spatiotemporal routing allocation matrix, ultimately generating a basic deflection angle command containing a spatial hysteresis time sequence.

[0010] The actuators respond accordingly. The damping adjustment unit inside each variable impedance servo actuator increases the physical mechanical damping coefficient at the local pivot hinge in real time according to the dynamic mechanical damping command; the internal position execution unit drives the micro control surface to generate asynchronous physical deflection actions in sequence according to the basic deflection angle command.

[0011] A second aspect of the present invention provides a flight stability augmentation control system for an aircraft, applied to the flight stability augmentation control method for an aircraft described in the first aspect of the present invention, the system comprising: The multi-port system modeling module is used to establish an open-loop multi-port Hamiltonian dynamics model of the aircraft.

[0012] The energy conjugate variable acquisition module is used to acquire real-time state variables and energy conjugate variables corresponding to the open-loop multi-port Hamiltonian dynamics model through an energy state sensing network deployed on the aircraft.

[0013] The energy shaping solution module is used to execute a passive energy shaping control algorithm based on interconnection and damping allocation, based on the real-time state variables and the energy conjugate variables, to calculate the generalized control input variable set for controlling the aircraft.

[0014] The two-dimensional command decoupling module is used to parse the generalized control input variable group and convert it into two-dimensional control commands that drive each variable impedance servo actuator set on the trailing edge of the aircraft wing. The two-dimensional control commands include a basic deflection angle command and a dynamic mechanical damping command. Each variable impedance servo actuator independently changes the deflection position of the corresponding micro control surface and the local hinge mechanical damping coefficient according to the two-dimensional control commands.

[0015] This invention provides a method and system for augmented flight control of an aircraft. It offers the following advantages: 1. This invention decouples the generalized control input variable group into an active generalized force component vector and a damped generalized force component vector, and uses a variable impedance servo actuator to independently execute the basic deflection angle command and the dynamic mechanical damping command, respectively. This changes the traditional single control approach of aircraft that relies solely on control surface deflection to suppress elastic vibration. This two-dimensional independent control mechanism enables the system to directly dissipate high-frequency aeroelastic residual energy by utilizing the mechanical damping at the local pivot hinge, reducing the risk of physical position command saturation of the actuator under high-frequency alternating loads and ensuring the reliability of the actuation system.

[0016] 2. This invention extracts the physical propagation phase velocity parameters of the dominant aeroelastic deformation wave and constructs a spatiotemporal routing allocation matrix by combining the spanwise geometric relative distances of each micro-control surface. This matrix then generates a basic deflection angle command containing a spatial hysteresis time sequence. This mechanism drives multiple control surfaces to produce asynchronous physical deflection actions, forming transient local aerodynamic forces with a specific phase difference in space. This actively changes the transmission boundary and phase of the wing aeroelastic deformation wave, causing the elastic potential energy accumulated inside the wing structure to be directionally converted into the restoring torque of the fuselage rigid body attitude, thereby improving the stability augmentation flight control efficiency of the aircraft.

[0017] 3. This invention addresses the physical underactuated property where the number of actuators is less than the system's generalized state vector degrees of freedom. In solving the multi-port Hamiltonian matching equation, it utilizes a full-rank pseudo-inverse matrix to obtain the minimum norm solution, maintaining global asymptotic stability of the entire system without truncating the higher-order elastic modes of the structure. Simultaneously, it sets an operating dead zone threshold constant in the generation and calculation of dynamic mechanical damping commands. When the absolute value of the physical deflection angular velocity of the control surface is too low, it directly interrupts the division operation and outputs the system's basic sustaining damping constant. This eliminates the potential for computational divergence from the digital processing level, improving the operational safety of the control system when dealing with extreme gust wind conditions. Attached Figure Description

[0018] Figure 1 This is a schematic diagram of the method flow of the present invention; Figure 2 This is a schematic diagram of the physical-mechanical damping isomorphic mapping process of the present invention; Figure 3 This is a schematic diagram of the aeroelastic wave spatiotemporal routing and phase modulation logic of the present invention; Figure 4 This is a schematic diagram of the system architecture of the present invention; Figure 5 This is a schematic diagram showing the comparison of dynamic response and actuator output curves of the aircraft of the present invention under discrete gust disturbances; in, Figure 5In the diagram, (a) is a comparison diagram of the elastic deformation displacement response of the wingtip, (b) is a comparison diagram of the pitch angle attitude response of the rigid body of the fuselage, and (c) is a comparison diagram of the position servo command output of the actuator. Detailed Implementation

[0019] The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0020] This invention provides a flight stability augmentation control system for an aircraft, which may include: A high aspect ratio flexible fixed-wing aircraft airframe comprises a fuselage and flexible wings connected to both sides of the fuselage. The flexible wings have spanwise distributed spars and aerodynamic outer surfaces.

[0021] Multiple micro-control surfaces are distributed and installed along the trailing edge spanwise of the flexible wing. Each micro-control surface is physically independent of the others and is connected to the trailing edge load-bearing structure of the flexible wing through a corresponding actuation mechanism.

[0022] The actuation mechanism driving the micro-manipulation surface is a variable impedance servo actuator. Each variable impedance servo actuator includes a physical position servo motor and a variable mechanical damping hinge mechanism. The physical position servo motor is used to drive the corresponding micro-manipulation surface to produce a deflection action. The variable mechanical damping hinge mechanism is connected between the output shaft of the physical position servo motor and the micro-manipulation surface, and is used to change the mechanical damping coefficient at the physical connection of the hinge in real time according to the received control command.

[0023] An energy state sensing network is deployed inside the fuselage of a high aspect ratio flexible fixed-wing aircraft. The energy state sensing network includes an inertial measurement unit, a fiber Bragg grating strain sensor array, and a microelectromechanical dynamic surface pressure sensor array.

[0024] The inertial measurement unit (IMU) is fixed at the center of gravity of the fuselage and is used to acquire the rigid body attitude parameters of the high aspect ratio flexible fixed-wing aircraft in real time. The rigid body attitude parameters include three-axis attitude angles and three-axis attitude angular rates.

[0025] A fiber Bragg grating strain sensor array is attached along the spanwise direction to the surface of the wing spars structure inside the flexible wing. The fiber Bragg grating strain sensor array is configured with multiple distributed strain measurement nodes, each of which is used to acquire local structural curvature and structural strain parameters at its respective wing spars location in real time.

[0026] A microelectromechanical dynamic surface pressure sensor array is embedded in the aerodynamic outer surface of a flexible wing. The array is configured with multiple distributed pressure measurement nodes, each corresponding to the strain measurement nodes in the spanwise coordinate system. Each pressure measurement node is used to measure local transient aerodynamic load parameters at its location on the wing surface in real time.

[0027] The aircraft's stability augmentation flight control system also includes a flight control computer. The flight control computer establishes bidirectional data communication connections with the inertial measurement unit, the fiber Bragg grating strain sensor array, the microelectromechanical dynamic surface pressure sensor array, and each variable impedance servo actuator via an onboard digital communication bus.

[0028] The flight control computer receives the local structural strain parameters output by the fiber Bragg grating strain sensor array and uses them as inputs to the generalized coordinate parameters of the structural elastic domain in the Hamiltonian energy model of the aircraft system.

[0029] The flight control computer receives the local transient aerodynamic load parameters output by the microelectromechanical dynamic surface pressure sensor array and uses them as the generalized force parameter input to the aerodynamic domain port in the Hamiltonian energy model of the aircraft system.

[0030] The control processor inside the flight control computer extracts the aeroelastic instantaneous power flow parameter at the current structural node of the aircraft based on the generalized coordinate parameters and generalized force parameters acquired at the same time. This instantaneous power flow parameter is used as the reference feedback variable for the energy shaping algorithm to calculate and output the deflection angle command and physical damping coefficient adjustment command to drive each variable impedance servo actuator.

[0031] See attached document Figure 1 This invention provides a flight control method for stabilizing an aircraft, executed by a flight control computer. The method includes a series of control execution steps that are sequentially and logically related. The flight control computer first establishes an open-loop multi-port Hamiltonian dynamics model based on the physical properties of the high-aspect-ratio flexible fixed-wing aircraft.

[0032] After the model is established, the flight control computer obtains the real-time aircraft state variables and energy conjugate variables through the energy state perception network.

[0033] Based on the acquired state variables and energy conjugate variables, the flight control computer executes a passive energy shaping control algorithm based on interconnection and damping allocation to calculate the desired set of generalized control input variables.

[0034] The flight control computer analyzes the generalized control input variable set and maps it into two-dimensional control commands that drive each variable impedance servo actuator through spatiotemporal allocation rules. The two-dimensional control commands include basic deflection angle commands and dynamic mechanical damping commands.

[0035] Each variable impedance servo actuator executes corresponding physical actions according to the received two-dimensional control commands, changes the deflection position of the corresponding micro-control surface and the local hinge mechanical damping coefficient, and completes the dissipation of aeroelastic potential energy and the correction of the rigid body attitude of the aircraft.

[0036] The technical details of establishing the open-loop multi-port Hamiltonian dynamics model in the above control method are elaborated here. The flight control computer decomposes the continuous motion of the high aspect ratio flexible fixed-wing aircraft into a large-scale spatial motion subsystem in the rigid body domain and a local deformation subsystem in the structural elastic domain.

[0037] To characterize the transient properties of the entire system, a global generalized state vector of the system is constructed. Generalized state vector satisfy .

[0038] In the generalized state vector In the definition, each symbol uses English characters as subscripts. Among them, A generalized coordinate column vector representing the rigid body attitude of an aircraft; The generalized momentum column vector representing the rigid body attitude of an aircraft; A generalized coordinate column vector representing the flexible deformation of the wing structure; This represents the generalized momentum column vector representing the flexible deformation of the wing structure. (Superscript) This represents the transpose operation of a matrix or vector.

[0039] Establish the total open-loop Hamiltonian energy function of the aircraft system in inertial space. Total open-loop Hamiltonian energy function Used to quantize the current state vectors of rigid body domains and structural elastic domains. The total mechanical energy is calculated using the following formula: ; In the formula, The inverse matrix representing the rigid body inertia matrix associated with the generalized coordinates of the rigid body; The inverse matrix representing the generalized mass matrix of the flexible wing structure; The generalized stiffness matrix represents the flexible wing structure; This represents the gravitational potential energy field of the aircraft in Earth's gravitational field, related to the generalized coordinates of a rigid body. The sum of the first and second terms in the formula represents the total kinetic energy of the system, and the sum of the third and fourth terms represents the total potential energy of the system.

[0040] Based on the energy conservation principle and port energy interaction mechanism in analytical mechanics, the flight control computer will use the total open-loop Hamiltonian energy function. By taking the derivative, an open-loop multi-port Hamiltonian dynamic differential equation is constructed to describe the continuous dynamic response of the aircraft: ; In the formula, the left side of the equal sign Represents the generalized state vector The derivative with respect to time. The first term on the right-hand side of the equation is the internal state evolution term of the system, where... The system's internal interconnection matrix satisfies the antisymmetric property, i.e. Interconnection matrix Physically, it represents an internal energy exchange channel that maps the rigid body translation and rotation components inside the aircraft to the bending and torsional components of the wing structure without energy loss.

[0041] Let be the system's intrinsic dissipation matrix, which satisfies the positive semi-definite property, i.e. Dissipation matrix Physically mapping the inherent structural damping characteristics inside the composite flexible wing and the natural aerodynamic damping characteristics of the wing surface under steady flight conditions, this characterizes the physical process by which the system converts mechanical energy into heat energy dissipation.

[0042] This is the column vector of partial derivatives of the system energy function with respect to the state variables, also known as the energy gradient or conjugate variable. The physical essence of the conjugate variable is the system restoring force. Within the wing structure displacement subspace, this partial derivative is directly equivalent to the generalized structural elastic restoring force generated under the current deformation state of the wing.

[0043] The second term on the right-hand side of the differential equation is the control input term. This represents the generalized control input vector jointly output by various variable impedance servo actuators; Input configuration matrix. The local aerodynamic generalized forces generated by the micro-control surfaces distributed at specific geometric coordinate positions on the trailing edge of the wing are mapped to each degree of freedom of the generalized state vector of the whole aircraft, thus establishing the transformation relationship between the physical coordinates of the operation and the generalized coordinates of the system.

[0044] The third term on the right side of the differential equation This represents the energy input port for external aerodynamic disturbances associated with the system's state variables and time. This port characterizes the uncontrollable transient aerodynamic loads exerted on a flexible fixed-wing aircraft by gusts or atmospheric turbulence, which are the factors leading to the generalized state vector. An external energy input source that deviates from the balance point and excites the elastic modes of the wing structure.

[0045] This embodiment further details the setting of the closed-loop expected system model and stability criteria in the control law solution process. Addressing the technical problem of state deviation and elastic vibration divergence in the open-loop multi-port Hamiltonian dynamics model when encountering aerodynamic disturbances, the flight control computer executes a passive control algorithm based on interconnection and damping allocation. First, it constructs the desired dynamic response target of the system through a mathematical model.

[0046] The flight control computer sets the desired stable equilibrium point of the aircraft system based on preset flight mission commands. Desired stable equilibrium point The state vector satisfies .in, Let be the target rigid body attitude angle vector corresponding to the current flight command; in this equilibrium state, the generalized momentum of the rigid body attitude motion and the generalized momentum of the structural flexible deformation are both constrained to zero vectors. Let be the generalized displacement vector of the reference static elastic deformation of the wing structure under the action of aerodynamic loads in the target attitude.

[0047] Establish the expected stable equilibrium point Subsequently, the flight control computer constructs a closed-loop expected multiport Hamiltonian energy function in the control algorithm module. Closed-loop expectation multi-port Hamiltonian energy function Includes the total open-loop Hamiltonian energy function And the virtual potential energy function actively injected by the algorithm The mathematical relationship between the two can be expressed as follows: ; To ensure the global asymptotic stability of the closed-loop system throughout the entire flight envelope, the flight control computer applies the Lyapunov stability criterion to the desired multiport Hamiltonian energy function. In the shape constraints. Specifically, the operation involves setting... The function term structure, forced At the desired stable equilibrium point The state possesses a strictly global minimum energy state. This stability criterion translates into corresponding mathematical partial derivative constraints, i.e., the first-order partial derivatives satisfy... Furthermore, the Hessian matrix, which has second-order partial derivatives with respect to the state vector, satisfies the positive definite condition. This partial derivative constraint establishes that the system state always has a direction in the multidimensional space. The potential energy recovery gradient.

[0048] Based on the above-constructed closed-loop expected multiport Hamiltonian energy function The flight control computer establishes a closed-loop expected dynamic differential equation to constrain the physical path for the system state to converge to the equilibrium point. The equation expression is as follows: ; In the formula, To determine the closed-loop desired multi-port Hamiltonian energy gradient, Defined as the expected interconnection matrix. The expected interconnection matrix is ​​derived from the original system's internal interconnection matrix. Interconnect matrix with control injection Superimposed to satisfy The control-injected interconnect matrix itself satisfies the antisymmetric property. .

[0049] At the level of algorithm execution, the flight control computer adjusts the control injection interconnection matrix through calculation. The matrix element values ​​correspond to the intersection dimension of the rigid body state subspace variables and the structural elastic state subspace variables. Adjusting these matrix elements physically alters the coupling coefficient at the port from the elastic strain energy of the wing structure to the rigid body attitude kinetic energy of the fuselage, providing a calculation benchmark for the energy interconnection relationship to guide the expected phase shift of the wing deformation wave.

[0050] Similarly, in the equation Defined as the expected dissipation matrix. The expected dissipation matrix is ​​derived from the intrinsic dissipation matrix of the original system. With control injection dissipation matrix Superimposed to satisfy The control injection dissipation matrix satisfies the symmetric positive definite property. .

[0051] At the specific execution level of the algorithm, the flight control computer increases the control injection dissipation matrix. The values ​​of specific element components on the main diagonal specify the target energy dissipation rate that the system needs to achieve along the spatial position direction of each micro-control surface after absorbing overload energy due to external gust disturbances. The setting of the target energy dissipation rate directly constitutes the basic data source for the flight control system to issue physical-mechanical damping dissipation commands to the actuators.

[0052] This embodiment details how the flight control computer generates specific control commands based on a multi-port Hamiltonian model. The flight control computer transforms the desired closed-loop dynamic objective of the system into a set of generalized control input variables required by the actuators by analyzing algebraic matching equations.

[0053] The flight control computer extracts the system evolution terms and control input terms from the constructed open-loop multi-port Hamiltonian dynamic differential equations and simultaneously establishes them on both sides of the closed-loop desired dynamic differential equations. During this simultaneous calculation, the algorithm does not directly calculate the external aerodynamic disturbance ports. Instead of relying on the transient integral, it relies on the Lyapunov energy dissipation boundary property inherent in the passive control theory framework to include the unknown gust disturbance energy within the robust stability margin of the system.

[0054] By making the open-loop system dynamics converge to the desired closed-loop dynamics, the flight control computer constructs a nonlinear algebraic matching equation containing control input variables in the digital processing module. The mathematical expression of the matching equation is: ; In the matching equation, the set of terms on the right-hand side of the equation physically represents the total amount of transient generalized energy gradient that needs to be supplemented or extracted into the physical space of the system by the various micro-control surfaces distributed on the trailing edge of the wing in order for the aircraft system to reach the global minimum desired energy state.

[0055] For solving the matching equations, the flight control computer identifies the matrix dimensions based on the physical connectivity attributes of the high-aspect-ratio flexible fixed-wing aircraft. Since the total number of micro-control surfaces distributed along the wing trailing edge is typically less than the total number of rigid and structural elastic degrees of freedom encompassed by the generalized state vector of the entire aircraft system, the input configuration matrix... The rectangular matrix with fewer columns than rows indicates that the controlled object is in a physically underactuated state.

[0056] To address this underactuated property, the flight control computer employs pseudo-inverse matrix algebra. The computer reads the matrix elements corresponding to the current state variables and calculates a full-rank pseudo-inverse input configuration matrix. Subsequently, this pseudo-inverse matrix is ​​used to perform a left product operation on the multidimensional gradient vectors on the right-hand side of the matching equation, thereby obtaining the corresponding minimum norm solution within the control input subspace. This solution is the generalized control input vector required at the current moment. .

[0057] Obtaining the generalized control input vector Subsequently, the flight control computer executes the generalized control law decoupling procedure. Based on the algebraic origin of each physical matrix term on the right-hand side of the matching equation, the computer transforms the single calculation result into a column vector. The decomposition is divided into two independent generalized force component vectors with mutually exclusive execution objectives, and the decomposition relation satisfies... .

[0058] in, This represents the active generalized force component vector. The active generalized force component vector is derived from the control injection interconnect matrix involved in the matching equation. With virtual potential gradient The computational terms are mapped and generated. At the physical execution level, the specific values ​​contained in this component vector are used to instruct the micro-manipulation surfaces at the corresponding spatial locations to change the local aerodynamic boundaries in order to generate active aerodynamic torques for controlling the transfer of elastic waves to the rigid body motion ports.

[0059] in, Let be the damping generalized force component vector. The damping generalized force component vector is derived from the control injection dissipation matrix involved in the matching equation. The computational terms are generated separately. At the physical execution level, the specific values ​​contained in this component vector are used to indicate the additional structural motion damping and dissipation tasks undertaken by the actuator at the corresponding spatial location on the mechanical end, in order to eliminate high-frequency residual energy during aeroelastic deformation. This decoupling step establishes an independent signal source for the subsequent two-dimensional physical command output of the variable impedance servo actuator.

[0060] This embodiment describes in detail the process by which the flight control computer converts the generalized force target at the algorithm level into drive signals for specific physical devices. In traditional rigid position servo systems, flight control commands only include single control surface position requirements. To address the high-frequency saturation problem of actuators caused by large-scale flexible deformation of the aircraft under gust disturbances, this solution relies on variable impedance servo actuators at the physical connections of each micro-control surface to construct a two-dimensional independent command decoupling architecture.

[0061] The variable impedance servo actuator is mechanically composed of two controlled physical units connected in series or parallel: a position execution unit and a damping adjustment unit. The position execution unit changes the absolute deflection angle of the micro-control surfaces relative to the wing's load-bearing structure; the damping adjustment unit, equipped with a magnetorheological damper or a variable cross-section throttle valve, changes the mechanical damping coefficient at the actuation hinge node in real time based on the input current or voltage signal. This physical structure provides the hardware basis for the variable impedance servo actuator to simultaneously receive and execute position commands and mechanical damping commands separately.

[0062] In conjunction with the aforementioned hardware infrastructure, the flight control computer is equipped with a command decoupling and allocation module within the control processing layer. This module receives the active generalized force component vector and the damping generalized force component vector derived analytically from the aforementioned matching equations, and assigns them to the first... One variable impedance servo actuator (of which Discretize the dimensions of the total number of micro-manipulation surfaces.

[0063] In the first data channel, the instruction decoupling allocation module extracts the instruction for the first... The actuator's active generalized force component scalar value. This scalar value characterizes the magnitude of the active restoring torque generated by the change in aerodynamic downwash flow required by the system. The computer module uses a preset aerodynamic conversion operator to inversely map this generalized force component scalar value into a geometric deflection angle, generating a basic deflection angle command for the actuator. And send it to the position execution unit.

[0064] In the second data channel, the instruction decoupling allocation module extracts the instruction for the first... The generalized damping force component of the actuator is a scalar value. This scalar value characterizes the amount of transient mechanical residual energy that the micro-control surface is required to absorb or dissipate in its direction of motion. Based on the current physical deflection angular velocity of the micro-control surface, the computer module converts this generalized damping force component into the required viscous drag coefficient at the actuator hinge node, generating a dynamic mechanical damping command. It is then sent to the damping adjustment unit.

[0065] By decoupling the operations of the allocation module through instructions, for the first The total generalized force required by a variable impedance servo actuator Deconstructed into an algebraic sum of two physical mechanisms, their relationship satisfies the equation: ; In the formula, Extracted from the generalized control input vector The corresponding number in the middle The total driving component of each actuator; Characterizes the positive mapping function of aerodynamic generalized force corresponding to the basic deflection angle command; This is the variable representing the basic deflection angle command output by the computer. The variable representing the dynamic mechanical damping command output by the computer; For the first real-time feedback through sensors The physical deflection angular velocity parameter of a micro control surface.

[0066] Through the aforementioned decoupling architecture and mapping relationship, the system enables the deflection actions of each micro-control surface and the soft and hard physical drag characteristics of the mechanical nodes to be synchronously controlled in the time dimension. This implementation step achieves independent responses at the execution terminal level to the system's active attitude control torque generation operation and passive flutter mechanical damping dissipation operation.

[0067] See attached document Figure 2 This embodiment provides a detailed explanation of the specific computational steps for the command decoupling and allocation module to generate dynamic mechanical damping commands. This mapping step converts the virtual energy dissipation rate set by the control algorithm in the multi-port Hamiltonian state space into the physical viscous resistance parameters of the physical components of each variable impedance servo actuator hinge mechanism.

[0068] The flight control computer extracts the control injection dissipation matrix calculated by the aforementioned control algorithm. Simultaneously extract the desired energy gradient column vector of the current closed-loop system state. By performing matrix multiplication between the two, the computer constructs a column vector representing the generalized damping forces that need to be actively dissipated in each generalized degree of freedom of the entire system. Simultaneously, the computer, through measuring elements connected to the rotation axes of each micro-control surface, collects and constructs a column vector in real time containing the current deflection angular velocity of all micro-control surfaces in their physical states. .

[0069] After obtaining the aforementioned system-level parameters, the computer performs projection calculations to distribute the global generalized force to local physical nodes. The computer then calls the pre-defined diagonalized damping allocation mapping matrix stored in the memory unit. Damping assignment mapping matrix The diagonal elements in the matrix represent the force arm vector distribution properties of the aerodynamic centers of each independent micro-control surface on the trailing edge of the wing from the center of the aircraft's global coordinate system. The computer maps the damping assignment matrix. Multiply by the generalized damping force column vector on the left to reduce its dimension and convert it into the local desired controlled dissipative force column vector required at the physical coordinate origin of each variable impedance actuator.

[0070] Subsequently, the computer performs damping coefficient extraction calculations to obtain the specific hardware instructions for each actuator. Since the physical mechanical viscous damping force is equal to the product of the damping coefficient and the velocity, the computer divides each element in the locally desired controlled dissipative force column vector obtained from the above steps, in turn, by the current deflection angular velocity column vector. The angular velocity element in.

[0071] To ensure the safe operation of the physical execution system, the computer is configured with a division-to-zero saturation protection check process within the division operation stage. When the division operation is detected... When the absolute value of the deflection angular velocity of a micro-control surface is lower than the preset dead zone threshold constant, the computer interrupts the real-time division algebra operation for that node and locks the output value of the command channel corresponding to that actuator to the system's basic holding damping constant. When the first... When the absolute value of the deflection angular velocity of a micro-control surface is greater than or equal to the preset dead zone threshold constant, the computer performs the division operation normally.

[0072] Under normal operating conditions that pass the division-by-zero saturation protection check, the specific mathematical and algebraic operations that generate the complete dynamic mechanical damping instruction set array internally by the computer satisfy the following equation: ; In the formula, the left side of the equal sign represents the arrangement according to the sequence of physical positions in the spanwise direction. A column vector of mechanical damping commands for a variable impedance servo actuator; the symbols on the right side of the equation. This indicates that a Hadamard division operation is performed on each corresponding element of the two column vectors.

[0073] The flight control computer will complete the calculation and output the mechanical damping instruction set array after threshold protection judgment, and distribute and send it to the damping adjustment unit interface of each corresponding variable impedance servo actuator through the communication cable to control the change of physical damping coefficient at its hinge node.

[0074] See attached document Figure 3 This embodiment provides a detailed explanation of the specific calculation and allocation steps for the command decoupling and allocation module to generate the basic deflection angle command set array. This implementation step is used to convert the active generalized force component vector output by the algorithm into a physical deflection timing sequence of various micro-control surfaces distributed at different positions in the wing space, thereby changing the boundary conditions of the elastic traveling wave.

[0075] The flight control computer extracts the generalized coordinate column vector of structural flexible deformation within a continuous time step fed back by the fiber Bragg grating strain sensor array. The processor module within the computer performs differentiation on the spatial position coordinates and difference operations on the time dimension on the generalized coordinate parameters within this continuous time step. Through this operation, the processor module extracts the physical propagation phase velocity parameters and wave propagation direction vector of the dominant aeroelastic deformation wave propagating along the wing spanwise direction at the current moment.

[0076] After acquiring the physical propagation phase velocity parameters, the flight control computer reads the absolute spanwise spatial coordinate data of the wing trailing edge positions of each variable impedance servo actuator, which are pre-stored in the storage unit. Using the origin node of the currently extracted dominant aeroelastic deformation wave as the reference point, the computer calculates the spanwise geometric relative distance between the absolute spanwise spatial coordinates of each actuator and the reference point.

[0077] The computer divides the spanwise geometric relative distance of each actuator by the physical propagation phase velocity parameter to calculate the result for the first actuator. The theoretical time delay constant at the spatial cross-section where each variable impedance servo actuator is located. This constant represents the physical time required for the peak value of the current structural deformation wave to reach each specific maneuvering surface cross-section.

[0078] Based on the theoretical time delay constants corresponding to all microcontroller surface nodes, the flight control computer constructs a system-level spatiotemporal routing assignment matrix. The dimensionality of the spatiotemporal routing assignment matrix matches the number of control surfaces, and the matrix elements on its main diagonal are transformed from the theoretical time delay constants of the corresponding nodes into the corresponding phase modulation transfer operators. This matrix configuration establishes an algebraic mapping constraint between the spacecraft's spatial geometric coordinates and the physical time axis.

[0079] After completing the matrix construction, the flight control computer extracts the active generalized force component vectors obtained from the aforementioned decoupling. The computer then calculates the reference deflection angle magnitude column vector for each microcontroller surface under time-delay conditions using a preset aerodynamic inverse mapping function. Subsequently, the computer multiplies the spatiotemporal routing allocation matrix with this reference deflection angle magnitude column vector to generate the final basic deflection angle instruction set array for each variable impedance servo actuator.

[0080] Through the above calculations, the basic deflection angle commands generated for the micro-control surfaces distributed at various physical locations along the span are not synchronized in the time dimension, but rather present as a spatial hysteresis sequence containing a fixed phase difference. The flight control computer sends this command set array to the position execution unit of each variable impedance servo actuator.

[0081] Each actuator sequentially drives the micro-control surfaces to deflect according to the received spatial hysteresis sequence. This temporally distributed deflection action locally alters the transient downwash flow field and local lift boundary distribution on the wing's aerodynamic outer surface. Multiple local aerodynamic forces with time differences superimpose on the same wing structure, forcing the propagating wing structure elastic deformation wave to produce the expected structural spatial phase shift. At the Hamiltonian system dynamics level, this specific phase shift will adjust the interconnected coupling characteristics of the original system to match the desired closed-loop interconnection matrix, transforming the accumulated structural elastic potential energy inside the wing into rigid body attitude recovery torque applied to the fuselage node ports.

[0082] See attached document Figure 4 This invention also provides a stability-enhancing flight control system for an aircraft. This embodiment provides a detailed breakdown of the control system functional modules running within the flight control computer. This flight control system relies on the underlying processor architecture and storage units to perform digital logic operations, establishing an information interaction path between the software data processing channel and external physical hardware. For underactuated aircraft configurations employing variable impedance actuators, the flight control system includes an energy conjugate variable acquisition module, a multi-port system modeling module, an energy shaping calculation module, and a two-dimensional command decoupling and issuance module.

[0083] The energy conjugate variable acquisition module is equipped with a hardware communication receiving interface. This module receives triaxial attitude angle and angular rate data output from the inertial measurement unit via an airborne digital communication bus, discrete point structural curvature data output from the fiber Bragg grating strain sensor array, and local aerodynamic pressure data output from the microelectromechanical dynamic surface pressure sensor array.

[0084] The digital processor inside the energy conjugate variable acquisition module performs spline interpolation and spatial integration on the received discrete structural curvature data, converting it into a continuous generalized elastic coordinate vector. Subsequently, this module combines rigid body motion data and local aerodynamic pressure data to calculate the first-order partial derivatives of the total Hamiltonian energy function with respect to each generalized state variable under the current physical state of the aircraft, and directly outputs the conjugate variable column vector representing the transient elastic restoring force of each physical domain of the system to the subsequent module.

[0085] The multi-port system modeling module is equipped with a read-only storage area and a parameter configuration interface. This module retrieves the aircraft rigid body inertial matrix parameters, structural generalized mass matrix parameters, and structural generalized stiffness matrix parameters pre-stored in the read-only storage area. Using the retrieved static physical parameters and real-time variables passed from the previous stage, this module constructs the total Hamiltonian energy function and the corresponding state-space evolution differential equation representing the open-loop system in the memory subspace. Simultaneously, based on instructions issued by the external flight mission system, this module sets the target attitude equilibrium point and configures the desired closed-loop multi-port energy model digital structure, including active interconnect adjustment terms and dissipative injection terms, in the memory space.

[0086] The energy shaping solution module establishes internal data bus connections with the energy conjugate variable acquisition module and the multi-port system modeling module. This solution module is used to perform core algebraic optimization calculations. For the state shift problem caused by the introduced external aerodynamic disturbances, the solution module constructs a system of simultaneous numerical equations equivalent to matching equations within the processor register.

[0087] Because the number of micro-control surface channels configured in the system is less than the total system degrees of freedom, the solution module calls a pre-compiled linear algebra library to perform a pseudo-inverse matrix calculation program on the input configuration matrix. During the calculation cycle, the solution module performs a left multiplication operation on the obtained pseudo-inverse matrix and the deviation vector of the equations on both sides, calculates the minimum norm generalized control input column vector that maintains the closed-loop energy gradient, and extracts the active generalized force component containing the aerodynamic target and the damped generalized force component containing the mechanical dissipation target from the column vector.

[0088] The two-dimensional command decoupling and distribution module is configured at the data output end of the flight control system device, and includes a damping mapping calculation unit and a space routing configuration unit.

[0089] The damping mapping calculation unit receives the damping generalized force components. This unit distributes the global dissipation commands in spatial coordinates by invoking the spanwise physical arm distribution weight matrix. Internally, the unit performs a division operation, dividing the distributed local generalized dissipation force by the instantaneous deflection angular velocity of each control surface obtained through sensor feedback. During the division operation, the unit has a built-in threshold protection logic branch that terminates the division operation and outputs a constant when the angular velocity parameter falls below a set safety boundary, ultimately generating a dynamic mechanical damping command array for each actuator.

[0090] The spatial routing configuration unit receives the active generalized force components. The computation unit executes a time-difference and spatial coordinate comparison algorithm in memory to calculate the physical propagation phase velocity of the dominant wave of the current wing structure's elastic deformation. By comparing the fixed installation coordinate system and wave velocity parameters of each variable impedance servo actuator, the computation unit dynamically generates a one-dimensional time delay array internally and uses this time delay array to construct the corresponding spatiotemporal routing allocation matrix. After this matrix mapping calculation, the active generalized force components are transformed into a basic deflection angle command array with a spatial-temporal phase difference.

[0091] The flight control system uses its external output interface module to package the dynamic mechanical damping command array and the basic deflection angle command array into a standard serial communication data frame, which is then sent to the variable impedance servo actuators distributed on the trailing edge of the wing according to a set clock frequency. This drives the physical servo motors and variable mechanical damping hinge mechanisms to respond, thus completing the physical phase modulation and controlled dissipation of elastic wave energy within the aircraft system.

[0092] Specific application examples: In this application embodiment, the high aspect ratio flexible fixed-wing aircraft is in level flight cruise at an altitude of 3000 meters and a Mach number of 0.4. There are 6 micro control surfaces (3 on each side) evenly distributed along the span of the trailing edge of the wing, and each control surface is driven by a variable impedance servo actuator.

[0093] During the simulation experiment At that time, the system encountered a vertically upward 1-cos discrete gust disturbance, the energy of which was transmitted through the external aerodynamic disturbance port. The intake system causes significant bending-torsional coupling first-order elastic deformation of the wing structure and induces deviation of the rigid body pitch attitude of the fuselage.

[0094] The flight control computer collects real-time rigid body attitude parameters and structural strain parameters through an energy state sensing network to construct a global generalized state vector. Then, the total open-loop Hamiltonian energy function in the current state is calculated: ; In response to this disturbance, the flight control computer sets the desired stable equilibrium point to maintain level flight. And construct a closed-loop expected multi-port Hamiltonian energy function. By executing a passive control algorithm based on interconnection and damping allocation, the system internally generates the desired closed-loop dynamic differential equation: ; The computer solves the matching equation using a full-rank pseudo-inverse matrix to derive the generalized control input vector and decouples it. Experiments compare the response data of two control architectures: Reference group (traditional control): employs a traditional linear feedback control law, outputting a single physical position deflection command, relying on wing aerodynamic dissipation of vibration energy. Experimental group (this invention's scheme): extracts the spatiotemporal routing allocation matrix to generate a basic deflection angle command containing a spatial hysteresis time-series sequence. Meanwhile, based on the expected dissipation matrix Generate dynamic mechanical damping commands And send it to each variable impedance servo actuator.

[0095] See attached document Figure 5 The experimental verification results are explained in detail and quantitatively. (Attached) Figure 5 It contains three subgraphs.

[0096] First sub-figure (top): Comparison of elastic deformation displacement at the wingtip. This figure reflects the convergence characteristics of the local deformation subsystem in the structural elastic domain. The solid black line represents the proposed solution, and the dashed black line represents the conventional control solution. In the initial stage of gust action (2 to 3 seconds), the wingtip bends upward due to the energy injection from the gust. The conventional control solution, due to physical hysteresis in actuator response and delayed aerodynamic damping generation, results in large-amplitude reciprocating oscillations at the wingtip, with a peak value of approximately 0.45 meters and a convergence time exceeding 8 seconds. The proposed solution alters the aeroelastic wave phase through spatiotemporal routing rules and directly extracts elastic potential energy at the physical end by adjusting the local hinge mechanical damping coefficient, reducing the maximum wingtip deformation to within 0.3 meters and achieving oscillation convergence in approximately 4.5 seconds. This demonstrates the expected dissipation matrix... The effectiveness of mapping to terminal mechanical damping.

[0097] The second sub-figure (center): Comparison of rigid body pitch angle attitude. This figure reflects the stability of the rigid body domain spatial motion subsystem. The dotted line represents the initial trim reference before the aircraft is disturbed. After being disturbed, the asymmetric elastic deformation of the wing induces additional low-frequency pitch torque. Under the conventional control scheme (dashed line), due to the strong coupling between rigid body motion and elastic deformation not being decoupled, the pitch angle fluctuates significantly (maximum deviation of approximately 2.5 degrees). The scheme of this invention (solid line) changes the coupling channel coefficient for the transfer of elastic potential energy to the rigid body kinetic energy port through the active interconnection injection matrix. The deformation energy accumulated in the wing is directionally converted into attitude recovery torque, thereby suppressing the maximum pitch angle deviation to within 1 degree, and the aircraft quickly recovers to the desired stable equilibrium point.

[0098] The third sub-figure (bottom): Comparison of high-frequency command saturation response of the actuator. This figure reflects the hardware operational safety of the system when dealing with physical underactuation attributes and high-frequency loads. In the traditional scheme (dashed line), facing high-frequency aeroelastic traveling waves, the system forcibly outputs high-frequency, large-amplitude deflection angle commands, causing the actuator motor to frequently reach the physical position response limit within the 3s to 5s range, resulting in severe command saturation and signal distortion. The two-dimensional command decoupling architecture of this invention (solid line) separates the actuation task: it uses a smooth reference deflection angle command to provide active recovery aerodynamic torque, while the high-frequency vibration dissipation task is handled by a passive mechanical damper. Therefore, the actual deflection curve of the actuator position servo motor is smooth and has an extremely low frequency, eliminating the high-frequency command saturation risk of the actuator from a physical perspective.

Claims

1. A method for augmenting flight control of an aircraft, characterized in that, Includes the following steps: An open-loop multi-port Hamiltonian dynamics model of the aircraft is established, wherein the aircraft is a high aspect ratio flexible fixed-wing aircraft. By using the energy state sensing network deployed on the aircraft, real-time state variables and energy conjugate variables corresponding to the open-loop multi-port Hamiltonian dynamics model are obtained. Based on the real-time state variables and the energy conjugate variables, a passive energy shaping control algorithm based on interconnection and damping allocation is executed to calculate the generalized control input variable set for controlling the aircraft. The generalized control input variable group is parsed and converted into two-dimensional control commands that drive the various variable impedance servo actuators set on the trailing edge of the aircraft wing. The two-dimensional control commands include basic deflection angle commands and dynamic mechanical damping commands. Each of the variable impedance servo actuators independently changes the deflection position of the micro manipulation surface connected to the variable impedance servo actuator and the local hinge mechanical damping coefficient according to the two-dimensional control command.

2. The method for augmenting flight control of an aircraft according to claim 1, characterized in that, The steps for establishing the open-loop multi-port Hamiltonian dynamics model of the aircraft include: The continuous motion of the aircraft is decomposed into a rigid body domain spatial motion subsystem and a structural elastic domain local deformation subsystem, and a generalized state vector is constructed that includes rigid body attitude generalized coordinates, rigid body attitude generalized momentum, structural flexible deformation generalized coordinates, and structural flexible deformation generalized momentum. Construct a total open-loop Hamiltonian energy function, which includes a rigid body inertia matrix, a structural generalized mass matrix, and a structural generalized stiffness matrix; By differentiating the total open-loop Hamiltonian energy function, an open-loop multi-port Hamiltonian dynamic differential equation containing the system's internal interconnection matrix and intrinsic dissipation matrix is ​​constructed.

3. The method for augmenting flight control of an aircraft according to claim 2, characterized in that, The steps for obtaining the real-time state variables and energy conjugate variables corresponding to the open-loop multi-port Hamiltonian dynamics model through an energy state sensing network deployed on the aircraft include: The three-axis attitude angles and three-axis attitude angular rates of the aircraft are collected by an inertial measurement unit fixed to the center of gravity of the aircraft fuselage. By using an array of fiber Bragg grating strain sensors attached to the surface of the wing spars of the aircraft, local structural curvature and structural strain parameters are collected and used as generalized coordinate parameter inputs for the local deformation subsystem of the structural elastic domain. A microelectromechanical dynamic surface pressure sensor array embedded in the aerodynamic outer surface of the wing is used to collect local transient aerodynamic load parameters, which are then used as generalized force parameter inputs for the external aerodynamic domain ports.

4. The method for augmenting flight control of an aircraft according to claim 2, characterized in that, The steps of executing the passive energy shaping control algorithm based on interconnection and damping allocation include: Set the desired stable equilibrium point of the aircraft in response to preset flight mission commands; A closed-loop expected multi-port Hamiltonian energy function is constructed. The closed-loop expected multi-port Hamiltonian energy function is composed of the superposition of the total open-loop Hamiltonian energy function and the actively injected virtual potential energy function. The closed-loop expected multi-port Hamiltonian energy function is constrained to satisfy the partial derivative constraint condition of the global minimum energy state at the expected stable equilibrium point. A closed-loop expected dynamic differential equation is established, which includes an expected interconnection matrix formed by the superposition of the internal interconnection matrix and the control injection interconnection matrix of the system, and an expected dissipation matrix formed by the superposition of the intrinsic dissipation matrix and the control injection dissipation matrix.

5. The method for augmenting flight control of an aircraft according to claim 4, characterized in that, The step of calculating the set of generalized control input variables for controlling the aircraft includes: The open-loop multi-port Hamiltonian dynamic differential equation and the closed-loop expectation dynamic differential equation are combined on both sides of the equation, and the set of terms containing the energy gradient column vector of the closed-loop expectation multi-port Hamiltonian energy function is extracted to construct a nonlinear algebraic matching equation. For the physical underactuated property where the total number of microcontroller surfaces is less than the total number of degrees of freedom covered by the generalized state vector, calculate the full-rank pseudo-inverse matrix of the aircraft input configuration matrix; The full-rank pseudo-inverse matrix is ​​used to perform a left product operation on the multidimensional gradient vector on the right side of the nonlinear algebraic matching equation to obtain the corresponding minimum norm solution in the control input subspace, and the minimum norm solution is used as the generalized control input variable set at the current time.

6. The method for augmenting flight control of an aircraft according to claim 5, characterized in that, The steps of parsing the generalized control input variable set and converting it into two-dimensional control commands to drive the various variable impedance servo actuators set on the trailing edge of the aircraft wing include: The generalized control input variable group is split into an active generalized force component vector and a damping generalized force component vector; The active generalized force component vector is generated by mapping the calculation terms involving the control injection interconnect matrix and the gradient of the virtual potential energy function in the nonlinear algebraic matching equation, and is used to generate the basic deflection angle command. The damping generalized force component vector is generated by mapping the calculation terms involving the control injection dissipation matrix in the nonlinear algebraic matching equation separately, and is used to generate the dynamic mechanical damping command.

7. The method for augmenting flight control of an aircraft according to claim 6, characterized in that, The steps for generating the dynamic mechanical damping command include: Matrix multiplication is performed based on the control injection dissipation matrix and the desired energy gradient column vector under the current closed-loop system state to construct a generalized damping force column vector; Call the damping allocation mapping matrix, multiply the damping allocation mapping matrix on the left by the generalized damping force column vector, and convert it into a local desired controlled dissipation force column vector; Divide each element in the column vector of the local desired controlled dissipation force by the corresponding element in the column vector of the physical deflection angular velocity of each of the micro-manipulation surfaces collected and fed back at the current moment. When it is determined that the absolute value of the physical deflection angular velocity corresponding to the micro-control surface is lower than the preset dead zone threshold constant, the division algebra operation for the physical node is interrupted and the system basic sustaining damping constant is output, generating the dynamic mechanical damping command.

8. The method for augmenting flight control of an aircraft according to claim 6, characterized in that, The steps for generating the basic deflection angle command include: Extract the physical propagation phase velocity parameters of the dominant aeroelastic deformation wave; The spanwise geometric relative distance between the wing trailing edge position of each of the variable impedance servo actuators and the dominant aeroelastic deformation wave source node is calculated using the absolute spanwise spatial coordinates. The theoretical time delay constant is calculated by dividing each of the spanwise geometric relative distances by the physical propagation phase velocity parameter. A spatiotemporal routing allocation matrix matching the number of micro maneuvering surfaces is constructed using all the theoretical time delay constants, and the active generalized force component vector is converted into a reference deflection angle magnitude column vector under time delay-free conditions; The spatiotemporal routing allocation matrix is ​​multiplied with the reference deflection angle magnitude column vector to generate the basic deflection angle instruction containing the spatial hysteresis time series.

9. The method for augmenting flight control of an aircraft according to claim 8, characterized in that, The steps for each of the variable impedance servo actuators to independently change the deflection position of the micro-manipulation surface connected to the variable impedance servo actuator and the local hinge mechanical damping coefficient according to the two-dimensional control command include: The damping adjustment unit inside each of the variable impedance servo actuators increases the physical and mechanical damping coefficient at the local pivot hinge of the micro-operation surface in real time according to the received dynamic mechanical damping command. The position execution unit inside each of the variable impedance servo actuators drives the connected micro manipulation surfaces to generate asynchronous physical deflection actions in sequence according to the received basic deflection angle command containing the spatial hysteresis timing sequence.

10. A flight stability augmentation control system for an aircraft, characterized in that, A flight control method for stabilizing an aircraft as described in any one of claims 1-9, comprising: A multi-port system modeling module is used to establish an open-loop multi-port Hamiltonian dynamics model of an aircraft, which is a high aspect ratio flexible fixed-wing aircraft. The energy conjugate variable acquisition module is used to acquire real-time state variables and energy conjugate variables corresponding to the open-loop multi-port Hamiltonian dynamics model through an energy state sensing network deployed on the aircraft. The energy shaping calculation module is used to execute a passive energy shaping control algorithm based on interconnection and damping allocation based on the real-time state variables and the energy conjugate variables, and calculate the generalized control input variable set for controlling the aircraft. The two-dimensional command decoupling module is used to parse the generalized control input variable group and convert it into two-dimensional control commands that drive each variable impedance servo actuator set on the trailing edge of the aircraft wing. The two-dimensional control commands include a basic deflection angle command and a dynamic mechanical damping command. Each variable impedance servo actuator independently changes the deflection position of the corresponding micro control surface and the local hinge mechanical damping coefficient according to the two-dimensional control commands.

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