Integrated design method for aerodynamic, dynamic, and control coupling of a variant aircraft

Abstract

The application provides a method for integrated design of aerodynamics, dynamics and control of a morphing aircraft, and belongs to the technical fields of numerical simulation, flight dynamics and control and wind tunnel test. The method comprises the following steps: determining full-size and morphing working condition parameters of the aircraft; obtaining various kinds of aerodynamic data in the morphing process and establishing a multi-rigid-body dynamics model; discretizing continuous morphing motion into state points and trimming, calculating transfer functions and eigenvalues of each state point, and designing a discrete PID control law; obtaining a continuous state space equation through gain interpolation, and building a time-varying control law; after numerical simulation verification, performing a wind tunnel test, making the model fly freely at a time when trimming is not possible, and intervening in the control law at a time when trimming is possible, observing the convergence time and iteratively optimizing the control law. The method realizes integrated design of aerodynamics, dynamics and control of the morphing aircraft, improves the design efficiency, and improves the flight quality in the morphing process.

Description

Technical Field

[0001] This invention relates to the fields of numerical simulation, flight dynamics and control, and wind tunnel testing, and in particular to an integrated design method for the coupling of aerodynamics, dynamics, and control of variant aircraft. Background Technology

[0002] With the increasing complexity of mission requirements for military and civilian aircraft, variability aircraft capable of flying across airspace and speed ranges and possessing multi-mission adaptability have become an important development direction. Variability aircraft actively change their shape by varying sweep angles, varying wingspans, and folding wings, thereby optimizing aerodynamic performance and expanding their flight envelope under different flight conditions.

[0003] However, the variable configuration process introduces strong unsteady aerodynamic characteristics and time-varying inertial properties, causing the aircraft's dynamic model to exhibit nonlinear and time-varying features, posing a severe challenge to flight control. In the traditional aircraft design process, aerodynamic analysis, dynamic modeling, and control law design are usually completed independently by different departments, and then integrated and verified at the end. This approach is not only time-consuming and inefficient, but also makes it difficult to fully consider the strong coupling effects between aerodynamics, dynamics, and control, easily leading to design iterations and increasing R&D costs and time.

[0004] To address the specific needs of vari-propellant aircraft, there is an urgent need for a method that can integrate aerodynamic data acquisition, multi-rigid-body dynamics modeling, and time-varying control law design. Current technologies lack a complete integrated design process from numerical simulation to wind tunnel testing, making it difficult to achieve real-time cancellation of unsteady aerodynamic and inertial disturbances during variable configurations. Therefore, developing an efficient and universal coupled design method is of great significance for improving the handling stability and flight performance of vari-propellant aircraft. Summary of the Invention

[0005] In view of the above-mentioned technical problems in related technologies, this invention proposes an integrated design method for the coupling of aerodynamics, dynamics and control of variant aircraft, which can overcome the above-mentioned shortcomings of the prior art.

[0006] To achieve the above-mentioned technical objectives, the technical solution of the present invention is implemented as follows: An integrated design method for the coupling of aerodynamics, dynamics, and control in variant aircraft; The integrated design method for the aerodynamics, dynamics, and control coupling of this variant aircraft includes the following steps: Step 1: Determine the length, characteristic dimensions, mass, inertia characteristics, and flight Mach number, altitude, angular rate of motion of the variant components, and flight status of the full-size aircraft. Step 2: Acquire various types of aerodynamic data of the aircraft during the configuration change process, and build a multi-rigid-body kinematics and dynamics model of the aircraft during the configuration change process; Step 3: Discretize the continuous rotation angles of the aircraft during configuration changes into rotation state points with angular intervals, balance the force and torque equations for the rotation state points, and calculate the transfer function of the inherent characteristics of the variant aircraft at each rotation state point based on the balancing results. Step 4: Analyze the short-period modal characteristics at each rotating state point, calculate the characteristic roots of the linearized state equation for each state point, and design the PID control law for each discrete point. Step 5: By using the scheduling gain method, the continuous state space equation during the deformation of the aircraft is interpolated into the state equation at the trim point, thus obtaining the state space equation and output equation during the deformation process. A dynamic control strategy is selected for the small disturbance model after trimming, and the control law of the nonlinear time-varying system is built. Step Six: Based on the acquired aerodynamic data, the constructed multi-rigid-body dynamics model, and the time-varying control law, complete the numerical simulation flight verification; Step 7: Based on the numerical simulation results, determine the parameters of the scaled-down test model, build a wind tunnel test flight trajectory calculation program with the time-varying control law, and calibrate the positioning accuracy and precision of the parallel six-degree-of-freedom mechanism in the wind tunnel. Step 8: After starting the wind tunnel and waiting for the flow field to stabilize, drive the model to move. At the state point where the equilibrium equation cannot be adapted, make the model perform free flight motion without rudder deflection control. When the variant aircraft reaches the moment when it can adapt to the equilibrium equation, intervene the time-varying control law and add rudder deflection control to the test model. Step Nine: After the variable configuration maneuver is completed, observe the convergence time of the aircraft displacement and attitude angle from the intervention of rudder deflection control. If the control law cannot achieve the control target, redesign the control law and repeat Step Eight until the integrated design of the variable aircraft aerodynamics, flight mechanics and control coupling is completed.

[0007] Furthermore, the length, characteristic dimensions, mass, and inertia characteristics of the full-size aircraft include: axial length, time-varying reference length, time-varying reference area, time-varying center of mass position, and time-varying axial, longitudinal, and lateral moments of inertia; the various types of aerodynamic data include: steady aerodynamic data under different deformation state apexes, different angles of attack, and different combinations of rudder deflection angles, unsteady aerodynamic data during deformation, and dynamic derivatives of different state apexes.

[0008] Furthermore, the multi-rigid-body kinematics and dynamics model includes kinematic equations, multi-rigid-body deformation dynamics equations, inertial forces and moments of inertia caused by the change in configuration, and the kinematic geometric relationships of the deformable components during the change in configuration.

[0009] Furthermore, the dynamic control strategy is selected from one or more of the following: LPV control strategy, preset performance control strategy based on disturbance observer, incremental nonlinear dynamic inverse control strategy, or deep reinforcement learning control strategy.

[0010] Furthermore, the specific method for obtaining the state-space matrix at the deformation time through interpolation using the scheduling gain method in step five is as follows: Let the rotation angle of the deformable component be... Define interpolation coefficients ,in and This represents the rotation angle of the component corresponding to two adjacent balancing state points; The state space matrix at the current time step , , , The following interpolation method is used to obtain: .

[0011] Furthermore, the dynamic control strategy selected in step five is a linear variable parameter control strategy, and the constructed time-varying control law adopts a proportional-integral-derivative controller with a filter. Its control law expression is: ; in, For error signals, These are proportional gain, integral gain, and derivative gain, respectively. The error signal is passed through a first-order low-pass filter. The output after that, is the filter time constant.

[0012] Furthermore, the timing of the intervention of the control law is as follows: the aerodynamic torque and aerodynamic force are measured in real time by the built-in six-component balance of the model and transmitted to the flight trajectory calculation program with the control law. When the program determines that the current state can meet the balance conditions of force and torque, the control law is activated.

[0013] Furthermore, the upper limit of the convergence time is 10 seconds. If this time is reached, regardless of whether the control objective is achieved, the parallel six-degree-of-freedom mechanism is commanded to drive the test model back to a safe position, and the current control law design is considered to have failed.

[0014] Furthermore, calibrating the positioning accuracy and precision of the parallel six-degree-of-freedom mechanism includes: obtaining the motor encoder value from the theoretical pose value based on the motion mathematical model of the six-degree-of-freedom parallel mechanism, and driving the mechanism to move; using an absolute measuring device to actually measure multiple sets of pose commands, and correcting the unknown parameters in the motion mathematical model through the nonlinear least squares method until the pose command and the actual measured pose error meet the precision requirements.

[0015] Furthermore, the mathematical model of the motion of the six-degree-of-freedom parallel mechanism is as follows: ; and: ; in, , , For the first The coordinates of the hinge point of the lateral moving ball hinge in the ground coordinate system; , , , , , For the first The slope and intercept of the hinge point coordinates of the branched static ball hinge with respect to the motor encoder value; for The value of the first branch motor encoder; For the first Support rod length; , , For the first The coordinates of the hinge point of the branched moving ball hinge in the model coordinate system; This is the transformation matrix from the model coordinate system to the ground coordinate system.

[0016] The beneficial effects of this invention are as follows: by integrating the acquisition of unsteady aerodynamic data of the variator aircraft, multi-rigid-body dynamics modeling and time-varying control law design into an integrated platform, and by adopting a timely intervention strategy of control law in wind tunnel testing, the three aspects of aerodynamics, dynamics and control can be optimized in a coordinated and iterative manner, avoiding the repetitive work caused by traditional serial design, thereby shortening the design cycle, reducing R&D costs, and improving the attitude stability and trajectory tracking accuracy of the aircraft during the variable configuration process. Attached Figure Description

[0017] To more clearly illustrate the specific embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the specific embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For those skilled in the art, other drawings can be obtained from these drawings without creative effort.

[0018] Figure 1 The response curve of the short-cycle modal analysis of the integrated design method for the coupling of aerodynamics, dynamics and control of the variant aircraft described in this invention; Figure 2 This is a schematic diagram of the PID control law of the integrated design method for coupling aerodynamics, dynamics, and control of the variant aircraft described in this invention. Figure 3 This is a schematic diagram illustrating the process of using deep learning to select controller structure and adjust parameters in the integrated design method of aerodynamics, dynamics, and control coupling of the variant aircraft described in this invention. Figure 4 This is a schematic diagram of the control law initiation process of the integrated design method for the aerodynamic, dynamic, and control coupling of the variant aircraft described in this invention. Figure 5 This is a schematic diagram of the structure of the integrated design method for the aerodynamic, dynamic, and control coupling of the variant aircraft described in this invention, where the variable sweep angle is spaced in ten-degree intervals as a state point. Figure 6 This is a schematic diagram of the pitch angle response curves under different trim states of the integrated design method for the aerodynamic, dynamic, and control coupling of the variant aircraft described in this invention. Figure 7 This is a flowchart of the integrated design method for the aerodynamic, dynamic, and control coupling of the variant aircraft described in this invention. Detailed Implementation

[0019] The technical solution of the present invention will be clearly and completely described below with reference to the embodiments. 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] In the description of this invention, it should be understood that the terms "center," "longitudinal," "lateral," "length," "width," "thickness," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," "outer," "clockwise," and "counterclockwise," etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are used only for the convenience of describing this invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limiting this invention.

[0021] Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of indicated technical features. Thus, a feature defined as "first" or "second" may explicitly or implicitly include one or more of the stated features. In the description of this invention, "a plurality of" means two or more, unless otherwise explicitly specified. Furthermore, the terms "installed," "connected," and "linked" should be interpreted broadly; for example, they may refer to a fixed connection, a detachable connection, or an integral connection; they may refer to a mechanical connection or an electrical connection; they may refer to a direct connection or an indirect connection through an intermediate medium; and they may refer to the internal connection of two components. Those skilled in the art can understand the specific meaning of the above terms in this invention based on the specific circumstances.

[0022] like Figures 1 to 7 As shown, the integrated design method for the aerodynamic, dynamic, and control coupling of a variant aircraft according to an embodiment of the present invention includes the following detailed steps: Step (1) Determine the length, characteristic dimensions, mass, and inertia characteristics of the full-size variable-sweep aircraft, including: the axial length of the full-size variable-sweep aircraft, the reference length that varies with time, the reference area that varies with time, the position of the center of mass that varies with time, and the axial, longitudinal, and lateral moments of inertia that vary with time.

[0023] Step (2) Determine the flight Mach number, flight altitude, and rotational angular rate of the aircraft when performing variable sweep maneuvers, including: selecting the flight Mach number when performing variable sweep based on previous numerical simulations, at which the unsteady aerodynamic effects are smaller compared to other Mach numbers; selecting the mission phase altitude based on the aircraft's flight mission; and selecting a variable sweep angular rate with a reduced frequency within the low-frequency range when performing variable sweep.

[0024] Step (3) Determine the flight state of the aircraft when it is in a variant, including: determining the peak of the variable sweep state when it is folded up to unfolded, as well as the angle of attack, rudder deflection angle, and throttle opening corresponding to the peak.

[0025] Step (4) Determine the types of aerodynamic data to be acquired for the variable-sweep aircraft, including: steady aerodynamic data at different sweep angle apexes, different angles of attack, and different control deflection angles under different combinations of sweep angles at the flight Mach number; unsteady aerodynamic data during variable sweep; and the corresponding dynamic derivatives at different apexes. The dynamic derivative is calculated as follows: Equation (1) In formula (1): Equation (2) in, Angle of attack that varies over time; Initial angle of attack; Angle of attack amplitude; This is the pitch oscillation frequency. This is the pitch moment coefficient; The pitch moment coefficient for the current state; These are the derivatives of the pitch moment coefficient with respect to the angle of attack, the rate of change of the angle of attack, and the pitch angular velocity, respectively. This represents the rate of change of angle of attack. To reduce the frequency; This is the reference length for the aircraft. The velocity of the incoming airflow to the aircraft.

[0026] Substituting equation (2) into equation (1), equation (2) can be expressed in the following form: Equation (3) In equation (3), and These are undetermined coefficients. After calculating the pitch moment coefficient versus time curve for a specific state, the curve is fitted using the least squares method to obtain... and These three fitting coefficients ultimately yield the static pitch moment coefficient and the combined pitch moment derivative, as shown in equation (4): Equation (4).

[0027] Step (5) Build a multi-rigid-body kinematics and dynamics model of the aircraft when it changes configuration, including: the modeling method includes the Newton-Euler method, the Lagrange method or the Kane method; the modeling equations include the kinematic equations, the multi-rigid-body deformation dynamics equations, the inertial forces and inertial moments caused by the change configuration, and the kinematic geometric relationships of the deformable parts when the configuration changes.

[0028] Step (6) determines the continuous flight control strategy during aircraft configuration changes, including: The control algorithms include: PID control algorithm, PIDF control algorithm, fuzzy control algorithm, and LQR control algorithm; The control strategies include: LPV (linear variable parameter) control strategy, preset performance control strategy based on disturbance observer, incremental nonlinear dynamic inverse control measurement, and deep reinforcement learning control strategy.

[0029] Step (7) Use numerical methods to calculate aerodynamic data and analyze characteristics of the variable sweep aircraft.

[0030] Step (8) discretizes the continuous rotation angles of the aircraft during its sweep into rotation state points with certain angular intervals, such as... Figure 5 As shown.

[0031] Step (9) balance the force and moment equations for the discrete variable sweep state points, and give the balanced angle of attack and rudder deflection angle. The balanced state is shown in Table 1.

[0032] Table 1. Trim Status of the Aircraft During Variable Swing

[0033] Step (10) Based on the trimmed angle of attack and rudder deflection, calculate the transfer function of the inherent characteristics of the variant aircraft at each variable sweep state point: Equation (5) In equation (5), For the speed of the aircraft For the mass of the aircraft, For flight time, Let be the slope of the engine thrust as a function of throttle opening. This refers to the engine throttle opening. For the angle of attack of the aircraft, The axial aerodynamic force acting on the aircraft The longitudinal aerodynamic force acting on the aircraft The force of gravity acting on the aircraft The inclination angle of the aircraft's trajectory. This refers to the axial inertial force experienced by the aircraft during variable sweep. This refers to the longitudinal inertial force experienced by the aircraft during its sweepback maneuver. The moment of inertia during sweepback. The pitch angular velocity, The basic pitch moment of the aircraft at zero angle of attack and zero rudder deflection. The slope of the pitch moment acting on the aircraft as a function of the angle of attack. The slope of the pitch moment acting on the aircraft as a function of the pitch angular velocity. The pitch angular velocity in the aircraft body coordinate system Projection of the axis The slope of the pitch moment acting on the aircraft as a function of the pitch control deflection angle. For pitch rudder deflection, This refers to the inertial torque experienced by the aircraft during deformation. The pitch angle.

[0034] Linearize equation (5) at the balancing point: Equation (6) in, For resistance speed The partial derivatives, For resistance angle of attack The partial derivatives, For thrust Throttle opening The partial derivatives, For lift speed The partial derivatives, For lift angle of attack The partial derivatives, For lift pitch rudder deflection The partial derivatives, pitching moment speed The partial derivatives, pitching moment angle of attack The partial derivatives, pitching moment pitch rudder deflection The partial derivatives, and it is important to note the inertial force. , and inertial torque We moved to the left side of the equation and analyzed which state variables the inertial force was related to based on the kinematic geometry of the variant aircraft's variable sweepback. Here, we assume that the inertial force, inertial torque, and the derivatives of the four state variables (i.e., If it relates to, then: Equation (7) Equation (8) Substituting equation (8) into equation (6), we get: Equation (9) Equation (9) can be written in state-space matrix form: Equation (10) Equation (11) Equation (12) Based on equations (10), (11), and (12), the state equations and output equations are as follows: Equation (13) Multiply the left side of equation (13) by This yields the following form: Equation (14) make Then, the final state-space expression is as follows: Equation (15) So, the transfer function of the inherent characteristics of the mutated aircraft The following formula represents: Equation (16) in, It is the identity matrix. For Laplace variables.

[0035] Step (11) Analyze the short-period modal characteristics of the variator at each state point: Taking a certain variator as an example, perform short-period modal response analysis under variable sweep. The variable sweep trim points are 0°, 10°, 20°, and 30°, and its pitch angle response curve is as follows: Figure 6 As shown. By Figure 6 It can be seen that their short-period characteristics are all divergent, and the degree of divergence varies under different balancing states.

[0036] Step (12) Based on the short-period modal characteristics, calculate the characteristic roots of the linearized state equation of the aircraft at each state point; Equation (17) in, It is an eigenvalue. Is with The identity matrix of the same order, det, represents the determinant. Solving this characteristic equation yields all eigenvalues. The real part of the eigenvalues ​​determines stability: if all real parts are negative, the system is stable; if the real parts are positive, the system diverges; if the real parts are zero and the imaginary parts are non-zero, the system is critically stable (oscillating with constant amplitude).

[0037] Step (13) Based on the characteristic roots of the state equation, design the PID control law for each discrete point. The flowchart for designing the PID control law for each discrete point is as follows: Figure 2As shown, when the output of the aircraft's state equation differs from the command target, the difference is fed into the proportional, integral, and derivative components, respectively, and finally summed to form the output signal of the control law. The proportional gain, integral gain, derivative gain, and filter constant are tuned using the ZN tuning method. Simultaneously, the output signal generated by the control law is used as the aircraft's input response and transmitted to the aircraft's inherent characteristic transfer function to observe the output response of the state equation, thus analyzing the aircraft's flight performance with the control law.

[0038] Step (14) uses the scheduling gain method to interpolate the continuous state-space equations during aircraft deformation into state equations at the balancing point, thus obtaining the state-space equations and output equations at a certain deformation moment. Assume the rotation angle of the deforming component is... Now define the interpolation coefficients : Equation (18) in, and For the aircraft at varying sweep angles In the first The and the first Between the variable sweep angles corresponding to each trim point, and the interpolation coefficients at this point... That is, the scheduling gain; therefore, through the scheduling gain For those in the first The and the first The state space matrix corresponding to each balanced state point is interpolated to obtain the state space matrix at that moment. The interpolation process is shown in equation (19): Equation (19).

[0039] Step (15) For the small disturbance model at the discretized balancing point, select the linear variable parameter (LPV) control law. Design the PIDF (proportional-integral-derivative) controller at the balancing point: Equation (20) in, It is proportional gain. It is integral gain. It is the differential gain. It is the filter time constant.

[0040] Now assume The state variable is a first-order low-pass filter. If the output is , then: Equation (21) The differential output with a first-order filter can then be expressed as: Equation (22) Introducing state variables The integral, taken as the error input, yields: Equation (23) Therefore, the total output is: Equation (24) By rearranging equations (21), (23), and (24), we can obtain: Equation (25) The state-space matrix can be obtained from equation (25): Equation (26) Based on the PIDF controller at the trim point designed above, the state space matrix obtained by the scheduling gain and interpolation in step (14) is used to interpolate the controller, and the "controller + aircraft inherent characteristics" are combined into a closed-loop system, as shown in the following equation: Equation (27) Among them, matrix Let be the state space matrix of the controller. As long as the closed-loop system is stable, the interpolated controller has completed the control of the aircraft during deformation.

[0041] Step (16) completes the numerical simulation flight verification based on "multiple types of aerodynamic data + multi-rigid-body dynamics model + time-varying control law". Among them, the multiple types of aerodynamic data include: static aerodynamic data of the variator at each variable sweep time, dynamic aerodynamic data of the variator throughout the variable sweep process, and pitch dynamic derivative data at each variable sweep time; the multi-rigid-body dynamics model includes: kinematic description, dynamic equations, inertial force caused by deformation, inertial torque caused by deformation, and kinematic geometry during deformation; the time-varying control law adopts the LPV (linear variable parameter) control law.

[0042] Step (17) Determine the scale, length, characteristic dimensions, mass, and inertia characteristics of the test model.

[0043] Step (18) Build a flight trajectory calculation program with the time-varying control law for wind tunnel testing. Specifically, the designed control law is packaged into a DLL format and embedded into the existing wind tunnel vehicle attitude calculation program to realize the real-time calling of the control law by the attitude calculation program.

[0044] Step (19) calibrates the positioning accuracy and precision of the axial displacement, lateral displacement, longitudinal displacement, pitch angle, yaw angle, and roll angle of the parallel six-degree-of-freedom mechanism in the wind tunnel. First, given the theoretical pose value, the motor encoder value is obtained by inverse solving according to the mathematical model of the six-degree-of-freedom parallel mechanism (Equation 27). The motor encoder rotates to drive the static ball joints of the six links of the parallel mechanism to perform linear motion, and finally the actual pose motion is completed by the end-stage secondary aircraft model of the parallel mechanism. At the same time, the 100 sets of pose commands are measured using an absolute measurement wall, and the unknown parameters in the mathematical model of motion are obtained by nonlinear least squares method. Then, 20 sets of command attitude angles are given again. If the pose command and the actual measured pose error do not meet the accuracy requirements, 20 sets of command attitude angles need to be measured again until they meet the accuracy requirements. The mathematical model of the six-degree-of-freedom parallel mechanism is: Equation (28) In equation (28), The static coordinate transformation of the moving ball hinge point of the i-th branch of a six-degree-of-freedom parallel mechanism in the ground coordinate system coordinate; Let x be the x-coordinate of the moving ball hinge point of the i-th branch of the six-degree-of-freedom parallel mechanism in the ground coordinate system after static coordinate transformation; Let y be the y-coordinate of the moving ball hinge point of the i-th branch of the six-degree-of-freedom parallel mechanism in the ground coordinate system after static coordinate transformation; Let z be the z-coordinate of the moving ball hinge point of the i-th branch of the six-degree-of-freedom parallel mechanism in the ground coordinate system after the static coordinate transformation. , Let x be the slope and intercept of the x-coordinate of the stationary ball hinge point of the i-th branch of a six-degree-of-freedom parallel mechanism with respect to the value of the motor encoder. , Let y-coordinate of the i-th branch static ball hinge point of a six-DOF parallel mechanism be the slope and intercept of the motor encoder value with respect to the value of the motor encoder. , Let z be the slope and intercept of the z-coordinate of the stationary ball hinge point of the i-th branch of a six-degree-of-freedom parallel mechanism with respect to the value of the motor encoder. The encoder value for the i-th branch motor; Let be the length of the i-th branch.

[0045] In addition, the coordinate transformation relationship is as follows: Equation (29) In equation (29), , , Let be the coordinates of the moving ball hinge point of the i-th branch of the six-degree-of-freedom parallel mechanism in the model coordinate system; This is the transformation matrix from the coordinate system of the secondary aircraft model to the ground coordinate system.

[0046] Step (20) Start the wind tunnel.

[0047] Step (21) After the flow field stabilizes, start the parallel six-degree-of-freedom mechanism to drive the variable sweep model motion.

[0048] Step (22) At the state point where the equilibrium equation cannot be adapted, the model undergoes free flight motion without rudder deflection control under the influence of unsteady aerodynamic forces.

[0049] Step (23) When the variant aircraft reaches a point where it can adapt to the equilibrium equations, the time-varying control law intervenes, adding rudder deflection control to the test model to control the aircraft's attitude angle and displacement. Taking a variable-sweep aircraft as an example, the control law intervenes when the aircraft is performing a variable-sweep operation. The aerodynamic torque and aerodynamic force information are measured by the six-component balance built into the model and transmitted to the flight trajectory calculation program with the control law described in step (18). The trajectory calculation program uses numerical methods to balance the aerodynamic force and aerodynamic torque on the aircraft with other external forces and torques. If the balance cannot be achieved, the control law is not activated and the rudder deflection is not activated. If the balance condition is achieved, the control law DLL encapsulation program is activated, and the rudder deflection is activated in real time. The flowchart is shown below. Figure 4 As shown.

[0050] After the variator completes the sweepback maneuver in step (24), observe the convergence time of the variator's displacement and attitude angle from the intervention of the rudder deflection control to meet the flight quality requirements. If convergence is achieved, time-domain response indicators such as overshoot, steady-state error, delay time, peak time, and attenuation ratio should also be observed.

[0051] Step (25): If the observation time reaches 10 seconds, regardless of whether the control objective is achieved, the parallel six-degree-of-freedom mechanism will drive the test model back to a safe position. This is to ensure the safety of the wind tunnel and the model. Observations will cease after 10 seconds, and if the control law still does not take effect after 10 seconds, it indicates that the control law and control strategy are poorly designed.

[0052] Step (26) Shut down the wind tunnel.

[0053] If the control law cannot achieve the control target in step (27), the control law needs to be redesigned. Repeat steps (20) to (26) to complete the design of the variable sweep time-varying control law scheme for the mutated aircraft under a certain working condition, and finally complete the integrated design of the aerodynamics, flight mechanics and control coupling of the mutated aircraft.

[0054] To facilitate understanding of the above technical solutions of the present invention, the following detailed description of the above technical solutions of the present invention is provided through specific embodiments and working principles.

[0055] Example: An integrated design scheme for the aerodynamic, dynamic, and control coupling of mutating aircraft is developed. This involves obtaining the aerodynamic characteristics, dynamic response characteristics, and control law schemes of the aircraft during mutating processes in simulation and experimental environments, thus achieving integrated design. The specific process is as follows: Figure 7 As shown.

[0056] Based on the flight mission requirements of a certain type of variant aircraft, we designed the aerodynamic shape, dynamic model, and control law for variable configuration under the required flight conditions. Now, we are carrying out the integrated scheme design in numerical simulation and wind tunnel test environments.

[0057] First, based on the predetermined aerodynamic shape of the variant aircraft, and according to the aerodynamic characteristic dimensions, mass and inertia parameters, and flight conditions of the shape, numerical simulation calculations are performed on the static aerodynamic characteristics, dynamic aerodynamic characteristics, and dynamic derivative characteristics. Static stability analysis, hysteresis effect analysis, and positive and negative dynamic derivative determination are then performed on the three aerodynamic characteristics.

[0058] Then, a multi-rigid-body dynamics model of the aircraft during deformation is constructed to express the inertial forces and moments caused by time-varying effects. The influence of the inputs (static aerodynamic data, unsteady aerodynamic data, dynamic derivatives, inertial forces, and moments) to the dynamics model is analyzed through numerical simulation. Discrete trim points in the continuous time-varying process are selected, and the equilibrium equations are trimmed according to the trim point states. The trim angle of attack and rudder deflection are given, and the short-period modes of the aircraft's inherent characteristic transfer function under each trim state are analyzed.

[0059] Based on the transfer function of the short-period mode of the aircraft, the characteristic roots at each discrete point are calculated. The control law at the discrete point is designed based on the characteristic roots. The state space equation and output equation at each discrete point are interpolated by the scheduling gain method. The state space equation is linearized by small perturbation. The controller equation is interpolated again by the scheduling gain. A suitable control strategy is selected to complete the design of the time-varying control law.

[0060] Having completed the integrated design of the aerodynamics, dynamics, and control scheme for the variator aircraft, wind tunnel tests were then conducted to observe the attitude angle response effect when the time-varying control law was introduced. If the observed control law is effective, the control effect of the time-varying control law was evaluated based on time-domain indicators such as overshoot time, convergence time, and decay rate of the attitude angle response.

[0061] In summary, by utilizing the above-mentioned technical solution of the present invention, the acquisition of unsteady aerodynamic data of the variator aircraft, multi-rigid-body dynamics modeling, and time-varying control law design are integrated into an integrated platform. Furthermore, a timely intervention strategy of the control law is adopted in wind tunnel testing. This enables the aerodynamic, dynamic, and control aspects to be synergistically and iteratively optimized, avoiding the repetitive work caused by traditional serial design. Consequently, the design cycle is shortened, the R&D cost is reduced, and the attitude stability and trajectory tracking accuracy of the aircraft during the variable configuration process are improved.

[0062] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, and not to limit them; although the present invention has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some or all of the technical features; and these modifications or substitutions do not cause the essence of the corresponding technical solutions to deviate from the scope of the technical solutions of the embodiments of the present invention.

Claims

1. A method of integrated design of a variable aircraft aerodynamic, dynamic, control coupling, characterized in that, Includes the following steps: Step 1: Determine the length, characteristic dimensions, mass, inertia characteristics, and flight Mach number, altitude, angular rate of motion of the variant components, and flight status of the full-size aircraft. Step 2: Acquire various types of aerodynamic data of the aircraft during the configuration change process, and build a multi-rigid-body kinematics and dynamics model of the aircraft during the configuration change process; Step 3: Discretize the continuous rotation angles of the aircraft during configuration changes into rotation state points with angular intervals, balance the force and torque equations for the rotation state points, and calculate the transfer function of the inherent characteristics of the variant aircraft at each rotation state point based on the balancing results. Step 4: Analyze the short-period modal characteristics at each rotating state point, calculate the characteristic roots of the linearized state equation for each state point, and design the PID control law for each discrete point. Step 5: By using the scheduling gain method, the continuous state space equation during the deformation of the aircraft is interpolated into the state equation at the trim point, thus obtaining the state space equation and output equation during the deformation process. A dynamic control strategy is selected for the small disturbance model after trimming, and the control law of the nonlinear time-varying system is built. Step Six: Based on the acquired aerodynamic data, the constructed multi-rigid-body dynamics model, and the time-varying control law, complete the numerical simulation flight verification; Step 7: Based on the numerical simulation results, determine the parameters of the scaled-down test model, build a wind tunnel test flight trajectory calculation program with the time-varying control law, and calibrate the positioning accuracy and precision of the parallel six-degree-of-freedom mechanism in the wind tunnel. Step 8: After starting the wind tunnel and waiting for the flow field to stabilize, drive the model to move. At the state point where the equilibrium equation cannot be adapted, make the model perform free flight motion without rudder deflection control. When the variant aircraft reaches the moment when it can adapt to the equilibrium equation, intervene the time-varying control law and add rudder deflection control to the test model. Step Nine: After the variable configuration maneuver is completed, observe the convergence time of the aircraft displacement and attitude angle from the intervention of rudder deflection control. If the control law cannot achieve the control target, redesign the control law and repeat Step Eight until the integrated design of the variable aircraft aerodynamics, flight mechanics and control coupling is completed.

2. The method of integrated design of a variable aircraft aerodynamic, dynamic, control coupling according to claim 1, characterized in that, The length, characteristic dimensions, mass, and inertia characteristics of the full-size aircraft include: axial length, time-varying reference length, time-varying reference area, time-varying center of mass position, and time-varying axial, longitudinal, and lateral moments of inertia; the various types of aerodynamic data include: steady aerodynamic data under different deformation state apexes, different angles of attack, and different combinations of rudder deflection angles, unsteady aerodynamic data during deformation, and dynamic derivatives of different state apexes.

3. The method of claim 1, wherein, The multi-rigid-body kinematics and dynamics model includes kinematic equations, multi-rigid-body deformation dynamics equations, inertial forces and moments of inertia caused by deformation, and the kinematic geometry of deformable components during deformation.

4. The integrated design method for the aerodynamic, dynamic, and control coupling of a variator aircraft according to claim 1, characterized in that, The dynamic control strategy is selected from one or more of the following: LPV control strategy, preset performance control strategy based on interference observer, incremental nonlinear dynamic inverse control strategy, or deep reinforcement learning control strategy.

5. The integrated design method for the aerodynamic, dynamic, and control coupling of a variant aircraft according to claim 1, characterized in that, The specific method for obtaining the state-space matrix at the deformation time through interpolation using the scheduling gain method in step five is as follows: Let the rotation angle of the deformable component be... Define interpolation coefficients ,in and This represents the rotation angle of the component corresponding to two adjacent balancing state points; The state space matrix at the current time step , , , The following interpolation method is used to obtain: 。 6. The integrated design method for the aerodynamic, dynamic, and control coupling of a variator aircraft according to claim 1, characterized in that, The dynamic control strategy selected in step five is a linear variable parameter control strategy, and the time-varying control law constructed uses a proportional-integral-derivative controller with a filter. Its control law expression is as follows: ； in, For error signals, These are proportional gain, integral gain, and derivative gain, respectively. The error signal is passed through a first-order low-pass filter. The output after that, is the filter time constant.

7. The integrated design method for the aerodynamic, dynamic, and control coupling of a variator aircraft according to claim 1, characterized in that, The timing for intervention of the control law is as follows: the aerodynamic torque and aerodynamic force are measured in real time by the built-in six-component balance of the model and transmitted to the flight trajectory calculation program with the control law. When the program determines that the current state can meet the balance conditions of force and torque, the control law is activated.

8. The integrated design method for the aerodynamic, dynamic, and control coupling of a variator aircraft according to claim 1, characterized in that, The upper limit of the convergence time is 10 seconds. If this time is reached, regardless of whether the control objective is achieved, the parallel six-degree-of-freedom mechanism is commanded to drive the test model back to a safe position, and the current control law design is considered to have failed.

9. The integrated design method for the aerodynamic, dynamic, and control coupling of a variator aircraft according to claim 1, characterized in that, The calibration of the positioning accuracy and precision of the parallel six-degree-of-freedom mechanism includes: obtaining the motor encoder value from the theoretical pose value by inverse solving the motion mathematical model of the six-degree-of-freedom parallel mechanism, and driving the mechanism to move; using an absolute measuring device to actually measure multiple sets of pose commands, and correcting the unknown parameters in the motion mathematical model by nonlinear least squares method until the pose command and the actual measured pose error meet the accuracy requirements.

10. The integrated design method for the aerodynamic, dynamic, and control coupling of a variant aircraft according to claim 9, characterized in that, The kinematic mathematical model of the six-degree-of-freedom parallel mechanism is as follows: ； and: ； in, , , For the first The coordinates of the hinge point of the lateral moving ball hinge in the ground coordinate system; , , , , , For the first The slope and intercept of the hinge point coordinates of the branched static ball hinge with respect to the motor encoder value; for The value of the first branch motor encoder; For the first Support rod length; , , For the first The coordinates of the hinge point of the branched moving ball hinge in the model coordinate system; This is the transformation matrix from the model coordinate system to the ground coordinate system.

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