Aeroelastic modeling and analysis method for boosting-energy complementing gliding aircraft

Abstract

The invention relates to the technical field of aeroelastic analysis of aero-engines, in particular to a boosting-energy complementing gliding aircraft thermal / aeroelastic coupling modeling and analysis method considering rigid-elastic coupling. A frame for simultaneously solving rigid body motion and elastic vibration is established, the amplitude and direction disturbance characteristics of thrust pulsation and the rigidity degradation effect caused by pneumatic heat generation-heat conduction are considered, and the problem of prediction deviation caused by neglecting of rigid-elastic coupling and an extreme thermal environment in traditional solving is solved. According to the method, aeroelastic response analysis of the aircraft in different load stages (full load, half load and no load) can be simulated, and hot gas bomb response under the action of thrust disturbance and extreme thermal load is solved, so that the prediction capability of dynamic behaviors of the aircraft under complex working conditions is improved; and a theoretical technical support is provided for aeroelastic response servo suppression of the boosting-energy complementing gliding aircraft in a large impact / extreme thermal environment.

Description

Technical Field

[0001] This invention relates to the field of aeroelasticity analysis technology for aero engines. Background Technology

[0002] As a core direction in the development of precision strike weapons globally, boost-recharge gliders are subject to multiple influences on their dynamic response behavior, including thrust disturbances, rigid-elastic coupling, and extreme thermal environments. This makes aeroelasticity and thermo-aeroelasticity critical issues affecting their design and development. Existing research largely focuses on single physical fields or simplified component-level models, lacking systematic modeling of multi-field coupling effects at the whole-vehicle level. Traditional methods often only consider the interaction between elastic deformation and aerodynamic forces, generally neglecting the dynamic coupling between rigid body motion and elastic vibration, and failing to adequately study the response characteristics of thrust disturbances, a key source of instability. Regarding thermo-coupling, existing models often employ quasi-static assumptions, failing to fully consider the interaction mechanism of aerodynamic heating, heat conduction, and structural stiffness degradation. Consequently, research on the overall dynamic response of boost-recharge gliders under complex conditions (such as combustion instability thrust pulsations, and the coupling of rigid body maneuvers and elastic vibrations) remains in the exploratory stage. A systematic modeling method capable of simultaneously analyzing rigid body motion, elastic vibration, heat conduction, and thrust disturbances needs to be developed. Summary of the Invention

[0003] The purpose of this invention is to overcome the limitations of existing technologies in multiphysics coupling modeling of boost-recharge gliders and to address the shortcomings in existing technologies for studying the dynamic response of such vehicles under complex conditions. It provides an aeroelastic / thermo-aeroelastic modeling and analysis method based on thrust disturbance-rigid-elastic coupling-heat conduction. By establishing a framework for simultaneously solving rigid body motion and elastic vibration, and considering the amplitude and direction disturbance characteristics of thrust pulsations and the stiffness degradation effect caused by aerodynamic heating and heat conduction, this invention solves the prediction bias problem caused by neglecting rigid-elastic coupling and extreme thermal environments in traditional solutions. This enables the analysis of the dynamic response of vehicles under complex conditions and provides a theoretical basis for the servo suppression of aeroelastic response in boost-recharge gliders under large impact / extreme thermal environments.

[0004] The technical solution adopted in this invention is an aeroelastic modeling method for a boost-recharge glider. The aeroelastic modeling includes aeroelastic response solving and modeling under rigid-elastic coupling, thermo-aeroelastic response solving and modeling under rigid-elastic coupling, and aeroelastic response solving and modeling under thrust disturbance. Specifically, it includes the following steps: Step 1: Solving and modeling the aeroelastic response under rigid-elastic coupling. A model for solving the aeroelastic response of an aircraft in a complex external flow field is established by using aeroelastic response modeling techniques that combine the finite element method with computational fluid dynamics. The structural module considers the coupling of rigid body motion and elastic vibration, allowing the aircraft to move freely in all six translational and rotational degrees of freedom, enabling the aircraft to translate and rotate freely under airflow excitation. The flow field module uses computational fluid dynamics to perform unsteady numerical simulations of the initial flow field mesh; The coupling interface between the flow field module and the structure module is modeled by dynamic meshing and data interpolation to obtain the transient displacement response of the aircraft monitoring point, i.e. the aeroelastic response under rigid-elastic coupling. Step 2: Solving and modeling the thermo-aeroelastic response under rigid-elastic coupling. Based on the aeroelastic response solution model established in step one, aeroelastic response solution modeling considering thermal effects is performed. Specifically, this includes: First, a thermal module solution model is established using a loosely coupled time-domain propagation method, which includes aerodynamic heating and heat conduction solutions. After obtaining the temperature field, the temperature change is regarded as an approximately steady-state process. Heat conduction and thermal stress effects are loaded into the aeroelastic response model. At the same time, the change of structural properties with temperature is modeled using a linear degradation model to achieve thermal aeroelastic response solution modeling under the influence of thermal effects, thus obtaining the thermal aeroelastic response of the aircraft under rigid-elastic coupling. Step 3: Modeling and solving the aeroelastic response of rigid-elastic coupling under thrust disturbance. Based on the aeroelastic response solution model established in step one, the aeroelastic response solution model considering thrust disturbance is carried out. First, the dynamic effects of combustion instability and thrust direction deviation on the aircraft structure are simulated, and the thrust disturbance is decomposed into two directions, including axial force and lateral force. Second, the thrust oscillation amplitude, frequency and phase angle are determined, and the thrust disturbance is simulated in the form of harmonics and introduced into the boundary conditions of the aircraft engine nacelle section. Finally, the aeroelastic response solution model under thrust disturbance is realized, and the aeroelastic response of rigid-elastic coupling under thrust disturbance of the aircraft is obtained.

[0005] Furthermore, in step one, the numerical solution of the structural module is based on the structural dynamics equations: , in, These are the structural mass matrix, damping matrix, and stiffness matrix, respectively. For the resultant external force of the structure, This represents its acceleration, velocity, and displacement; Simultaneously, considering rigid-elastic coupling, the rigid body motion degrees of freedom and elastic vibration degrees of freedom are taken into account in the structural governing equations: , , in, and These are the three translational velocities and three rotational angular velocities in the body coordinate system; and The aerodynamic forces and torques generated by rigid body motion; and The aerodynamic forces and torques generated by elastic vibration; m and These are the mass and inertia of the aircraft, respectively. It is the acceleration due to gravity; The transformation matrix from the ground coordinate system to the body coordinate system is given. The degrees of freedom of the rigid body motion are the three-axis translation and rotation around the axis of the aircraft's center of mass. The spatial solution format is constructed using finite element discretization technology and solved by combining it with the time integration algorithm to obtain the rigid-elastic coupled displacement response of the aircraft under aerodynamic loads.

[0006] Furthermore, the calculation of the flow field module is based on the numerical solution of the three-dimensional Reynolds-averaged three-dimensional Navier-Stokes equations: , in, For the solution vector, Let represent density, and u, v, w represent velocities in the three directions, respectively. For fluid domain The boundary, The outward normal vector of the boundary, vector flux From convection vector flux With viscous vector flux Composed of two parts ; In terms of numerical solution, the above fluid control equations are spatially discretized using the finite volume method and temporally discretized using the implicit second-order backward Euler method, employing the SST method. Turbulence models are used to obtain aerodynamic distribution.

[0007] Furthermore, the coupling modeling process includes: (1) Solve the flow field module to determine the nodal pressure distribution and flow field characteristics at the next moment; (2) The obtained pressure distribution is transmitted to the structural module using an interpolation algorithm; (3) Under the newly established structural module boundary conditions, solve the structural dynamic equations to predict the structural deformation at the next moment; (4) Using dynamic mesh technology, the latest structural deformation information is fed back to the flow field module, thereby completing a complete coupled iteration step; The entire coupling process follows this data exchange process, iterating until the preset computation time is reached.

[0008] Furthermore, the heat conduction in step two is described by Fourier's law, and the three-dimensional transient heat conduction equation based on the isotropic assumption is as follows: , in, For heat flux density, The thermal conductivity of the material, This represents the temperature distribution.

[0009] Furthermore, the time-domain propagation of the thermal module in step two adopts a loosely coupled method, and its solution process mainly includes the following steps: (1) Calculation of heat conduction inside the structure; (2) Interpolate the calculated object surface temperature to the flow field grid; (3) Use a computational fluid dynamics solver to perform calculations for aerodynamic heating; (4) Interpolate the surface heat flux density to the heat conduction calculation grid; In the aerodynamic thermal module, the thermal boundary conditions of the coupling boundary adopt a given temperature boundary condition; while in the structural heat conduction calculation, the coupling boundary adopts a given heat flux boundary condition; for the given temperature boundary condition, the following conditions are satisfied on the boundary: For a given heat flux boundary condition, the following holds at the boundary: Where T is temperature, q is heat flux, w represents the wall surface, t represents time, and x, y, z represent position coordinates; In step two, when considering thermal effects, it is assumed that the material properties exhibit a linear trend within the temperature range. Therefore, the material properties C (elastic modulus and coefficient of linear expansion) at a specific temperature T can be expressed in the following form: , in, It is a constant characterizing the rate at which material properties change with temperature.

[0010] Furthermore, the thrust disturbance in step three is modeled using sinusoidal thrust oscillations and lateral disturbance forces to simulate high-frequency disturbance effects that may occur during engine operation, employing harmonic function forms: , in, For steady-state thrust, For the disturbance amplitude, For the perturbation frequency, The phase angle is one factor. On the other hand, a slight deviation in the direction of engine thrust may be caused by factors such as structural vibration, uneven flow in the pipeline, or nozzle sway, which in turn can lead to lateral loads on the aircraft. Lateral disturbance force can be expressed as: , in, The amplitude of the lateral disturbance force. Its dominant frequency, This is the phase angle. This term applies to the location of the aircraft's engine nacelle and affects the overall vibration response.

[0011] This invention also provides an aeroelastic modeling and analysis method for a boost-recharge glider. Specifically, it includes decoupling the rigid-elastic coupling signal from the aeroelastic response solution models under rigid-elastic coupling, the thermo-aeroelastic response solution models under rigid-elastic coupling, and the aeroelastic response solution models under thrust disturbance at monitoring points on the aircraft. The decoupling technique combines empirical mode decomposition (EMD) and filtering decoupling techniques. First, the signal is decomposed into different frequency components using EMD to obtain the cutoff frequency. Then, a high-pass filter is used to extract the high-frequency elastic vibration component using this cutoff frequency, while simultaneously separating the low-frequency rigid body motion component. Finally, effective decoupling of the rigid body motion and elastic vibration of the displacement signal is achieved.

[0012] Furthermore, the empirical mode decomposition mainly includes the following steps for signal decomposition: (1) Identification signal The upper envelope is obtained by fitting all local maxima and local minima using cubic spline interpolation. and lower envelope To cover the entire data range; (2) Calculate the local mean ; (3) Calculate the difference between the signal and the mean: ,like If an eigenmode function satisfies the conditions that the number of zero-crossings and the number of extreme points differ by no more than 1, and the local mean is close to 0, then it is considered a first-order eigenmode function, denoted as . Otherwise, for Continue iterating the same envelope fitting process until the intrinsic mode function condition is met; set the cutoff frequency of the high-pass filter by the frequency of the empirical mode decomposition signal. .

[0013] The beneficial effects of this invention are as follows: This invention provides an aeroelastic modeling and analysis method for boost-recharge gliders. By establishing a framework for simultaneously solving rigid body motion and elastic vibration, it considers the amplitude and direction disturbance characteristics of thrust pulsation and the stiffness degradation effect caused by aerodynamic heat generation and heat conduction. This solves the prediction bias problem caused by neglecting rigid-elastic coupling and extreme thermal environment in traditional solutions. This method can simulate the aeroelastic response analysis of the aircraft under different load stages (full load, half load, and no load) and solve the thermo-aeroelastic response under thrust disturbance and extreme thermal load, thereby improving the prediction ability of the dynamic behavior of the aircraft under complex working conditions and providing theoretical and technical support for the servo suppression of aeroelastic response of boost-recharge gliders under large impact / extreme thermal environment. Attached Figure Description

[0014] Figure 1 is a flowchart of the implementation of the present invention; Figure 2 is a conceptual diagram of the HTV-2 implementation example of the present invention; Figure 3 shows the aeroelastic response of the HTV-2, an embodiment of the present invention, under full load, half load, and no load conditions. Figure 4 shows the aeroelastic response of the HTV-2, an embodiment of the present invention, under full load, half load, and no load conditions under thrust disturbance. Figure 5 shows the thermo-aeroelastic response of the HTV-2 in the embodiment of the present invention under full load, half load, and no load conditions after heating for 200s. Detailed Implementation

[0015] The principles and features of the present invention are described below with reference to the accompanying drawings. The examples given are only for explaining the present invention and are not intended to limit the scope of the present invention.

[0016] To achieve the above objectives, the present invention provides the following specific embodiments: Figure 1 As shown, an aeroelastic modeling method for a boost-recharge glider is characterized by the following steps: aeroelastic response modeling under rigid-elastic coupling, thermo-aeroelastic response modeling under rigid-elastic coupling, and aeroelastic response modeling under thrust disturbance. Step 1: Solving and modeling the aeroelastic response under rigid-elastic coupling. A model for solving the aeroelastic response of an aircraft in a complex external flow field is established by using aeroelastic response modeling techniques that combine the finite element method with computational fluid dynamics. Including structural modules, considering the coupling of rigid body motion and elastic vibration, the aircraft is allowed to move freely in all six translational and rotational degrees of freedom, enabling the aircraft to translate and rotate freely under airflow excitation; The flow field module uses computational fluid dynamics to perform unsteady numerical simulation of the initial flow field mesh. The coupling interface between the flow field module and the structure module is transformed into a coupled model of the structure and flow field by moving the mesh and data interpolation. This results in the transient displacement response of the aircraft monitoring points, i.e., the aeroelastic response under rigid-elastic coupling. Step 2: Solving and modeling the thermo-aeroelastic response under rigid-elastic coupling. Based on the aeroelastic response solution model established in step one, aeroelastic response solution modeling considering thermal effects is performed. Specifically, this includes: First, a thermal module solution model is established using a loosely coupled time-domain propagation method, which includes aerodynamic heating and heat conduction solutions. After obtaining the temperature field, the temperature change is regarded as an approximately steady-state process. Heat conduction and thermal stress effects are loaded into the aeroelastic response model. At the same time, the change of structural properties with temperature is modeled using a linear degradation model to achieve thermal aeroelastic response solution modeling under the influence of thermal effects, thus obtaining the thermal aeroelastic response of the aircraft under rigid-elastic coupling. Step 3: Modeling and solving the aeroelastic response of rigid-elastic coupling under thrust disturbance. Based on the aeroelastic response solution model established in step one, the aeroelastic response solution model considering thrust disturbance is carried out. First, the dynamic effects of combustion instability and thrust direction deviation on the aircraft structure are simulated, and the thrust disturbance is decomposed into two directions, including axial force and lateral force. Second, the thrust oscillation amplitude, frequency and phase angle are determined, and the thrust disturbance is simulated in the form of harmonics and introduced into the boundary conditions of the aircraft engine nacelle section. Finally, the aeroelastic response solution model under thrust disturbance is realized, and the aeroelastic response of rigid-elastic coupling under thrust disturbance of the aircraft is obtained.

[0017] Furthermore, in step one, the numerical solution of the structural module is based on the structural dynamics equations: , in, These are the structural mass matrix, damping matrix, and stiffness matrix, respectively. For the resultant external force of the structure, This represents its acceleration, velocity, and displacement; simultaneously, considering rigid-elastic coupling, the rigid body's degrees of freedom for motion and elastic vibration are taken into account in the structural governing equations: , , in, and These are the three translational velocities and three rotational angular velocities in the body coordinate system; and The aerodynamic forces and torques generated by rigid body motion; and The aerodynamic forces and torques generated by elastic vibration;m and These are the mass and inertia of the aircraft, respectively. It is the acceleration due to gravity; The transformation matrix from the ground coordinate system to the body coordinate system is given. The degrees of freedom of the rigid body motion are the three-axis translation and rotation around the axis of the aircraft's center of mass. The spatial solution format is constructed using finite element discretization technology and solved by combining it with the time integration algorithm to obtain the rigid-elastic coupled displacement response of the aircraft under aerodynamic loads.

[0018] The calculation of the flow field module is based on the numerical solution of the three-dimensional Reynolds-averaged three-dimensional Navier-Stokes equations: , in, For the solution vector, Let represent density, and u, v, w represent velocities in the three directions, respectively. For fluid domain The boundary, The outward normal vector of the boundary, vector flux From convection vector flux With viscous vector flux Composed of two parts ; In terms of numerical solution, the above fluid control equations are spatially discretized using the finite volume method and temporally discretized using the implicit second-order backward Euler method, employing the SST method. Turbulence models are used to obtain aerodynamic distribution.

[0019] The process of coupling modeling includes: (1) Solve the flow field module to determine the nodal pressure distribution and flow field characteristics at the next moment; (2) The obtained pressure distribution is transmitted to the structural module using an interpolation algorithm; (3) Under the newly established structural module boundary conditions, solve the structural dynamic equations to predict the structural deformation at the next moment; (4) Using dynamic mesh technology, the latest structural deformation information is fed back to the flow field module, thereby completing a complete coupled iteration step; The entire coupling process follows this data exchange process, iterating until the preset computation time is reached.

[0020] Furthermore, the heat conduction in step two is described by Fourier's law, and the three-dimensional transient heat conduction equation based on the isotropic assumption is as follows: , in, For heat flux density, The thermal conductivity of the material, This represents the temperature distribution.

[0021] Furthermore, the time-domain propagation of the thermal module in step two adopts a loosely coupled method, and its solution process mainly includes the following steps: (1) Calculation of heat conduction inside the structure; (2) Interpolate the calculated object surface temperature to the flow field grid; (3) Use a computational fluid dynamics solver to perform calculations for aerodynamic heating; (4) Interpolate the surface heat flux density to the heat conduction calculation grid. In the aerodynamic module, the thermal boundary conditions of the coupled boundary adopt the given temperature boundary conditions; while in the structural heat conduction calculation, the coupled boundary adopts the given heat flux boundary conditions; for the given temperature boundary conditions, the following conditions are satisfied on the boundary: For a given heat flux boundary condition, the following holds at the boundary: Where T is temperature, q is heat flux, w represents the wall surface, t represents time, and x, y, z represent position coordinates; In step two, when considering thermal effects, it is assumed that the material properties exhibit a linear trend within the temperature range. Therefore, the material properties C (elastic modulus and coefficient of linear expansion) at a specific temperature T can be expressed in the following form: , in, It is a constant characterizing the rate at which material properties change with temperature.

[0022] Furthermore, the thrust disturbance in step three is modeled using sinusoidal thrust oscillations and lateral disturbance forces to simulate high-frequency disturbance effects that may occur during engine operation, employing harmonic function forms: , in, For steady-state thrust, For the disturbance amplitude, For the perturbation frequency, The phase angle is one factor. On the other hand, a slight deviation in the direction of engine thrust may be caused by factors such as structural vibration, uneven flow in the pipeline, or nozzle sway, which in turn can lead to lateral loads on the aircraft. Lateral disturbance force can be expressed as: , in, The amplitude of the lateral disturbance force. Its dominant frequency, This is the phase angle. This term applies to the location of the aircraft's engine nacelle and affects the overall vibration response.

[0023] This invention also provides an aeroelastic modeling and analysis method for a boost-recharge glider. Specifically, it includes decoupling the rigid-elastic coupling signal from the aeroelastic response solution models under rigid-elastic coupling, the thermo-aeroelastic response solution models under rigid-elastic coupling, and the aeroelastic response solution models under thrust disturbance at monitoring points on the aircraft. The decoupling technique combines empirical mode decomposition (EMD) and filtering decoupling techniques. First, the signal is decomposed into different frequency components using EMD to obtain the cutoff frequency. Then, a high-pass filter is used to extract the high-frequency elastic vibration component using this cutoff frequency, while simultaneously separating the low-frequency rigid body motion component. Finally, effective decoupling of the rigid body motion and elastic vibration of the displacement signal is achieved.

[0024] Furthermore, the empirical mode decomposition mainly includes the following steps for signal decomposition: (1) Identification signal The upper envelope is obtained by fitting all local maxima and local minima using cubic spline interpolation. and lower envelope To cover the entire data range; (2) Calculate the local mean ; (3) Calculate the difference between the signal and the mean: ,like If an eigenmode function satisfies the conditions that the number of zero-crossings and the number of extreme points differ by no more than 1, and the local mean is close to 0, then it is considered a first-order eigenmode function, denoted as . Otherwise, for Continue iterating the same envelope fitting process until the intrinsic mode function condition is met; set the cutoff frequency of the high-pass filter by the frequency of the empirical mode decomposition signal. .

[0025] For a certain type of HTV-2 concept machine, such as Figure 2 As shown, numerical simulation analyses of aeroelasticity under rigid-elastic coupling of the entire aircraft, aeroelasticity under propulsion system disturbance, and thermo-aeroelasticity are conducted. The operating points are located within the altitude range [40, 50] km and the Mach number range [15, 20]. Based on the sensor layout strategy, approximately 75% of the positions in the front cabin of the aircraft are selected as characteristic monitoring points to record the dynamic response characteristics of the structure in the time domain. Elastic vibration frequencies were obtained through modal analysis and combined with high-pass filtering to separate motion components at different frequencies. The near-zero frequency signal primarily reflects the rigid body motion characteristics of the aircraft, while the high-frequency signal (tens of hertz) reveals the main features of elastic vibration. Since empirical modal decomposition requires upper and lower extrema, and some rigid body motions excessively cover elastic vibrations, a displacement signal in a direction with significant vibration was selected for trial calculation to obtain the cutoff frequency. To compare the impact of disturbances / extreme thermal environments on the structural dynamic response, a unified cutoff frequency was used in the example, with a normalized cutoff frequency of 0.64. Vibration displacement data were obtained after dimensionless conversion based on skin thickness, and time was dimensionless converted to the reciprocal of the first-order elastic modal frequency under no-load conditions.

[0026] To further simulate the impact of combustion instability on the aircraft structure, a sinusoidal excitation was applied to the engine nacelle section, with an average thrust set at 35 kN and an excitation frequency of 1000 Hz, to reproduce the dynamic characteristics of high-frequency fluctuations during combustion. Furthermore, to analyze the impact of directional disturbances on system stability, a lateral force with an amplitude of 10% of the main thrust was added to the simulation to simulate disturbances and asymmetric loads at the refueling moment.

[0027] Simultaneously, based on flight mission requirements, two refueling phases were planned, corresponding to three typical load states of the aircraft structure: fully loaded, half-loaded, and unloaded. Before the first refueling, the aircraft is in a fully loaded state, representing the condition of carrying all required energy at the beginning of the mission; before the second refueling, the state is adjusted to half-loaded, reflecting the operational characteristics after partial energy consumption in the middle of the mission; after refueling, it is in an unloaded state, simulating the situation where the aircraft enters the mission's endurance or return phase after refueling. These states are achieved by adjusting the mass distribution loading of the engine nacelle section and are used to evaluate the aeroelastic and thermo-aeroelastic coupling behavior under different load conditions.

[0028] In the aerodynamic-thermal-conduction coupling solution process, to more closely approximate the structural design in actual engineering, it is assumed that there is a 10mm thick insulation layer on the structural surface, and the heat transfer parameter is ρ = 73 kg / m². 3 c = 729 J / kg / K, k = 0.0258 W / m / K, treated as a wall boundary, and the elastic modulus is taken as e = -0.0005.

[0029] Figure 3 The results show the filtered response characteristics in the x, y, and z directions under full load, half load, and no load conditions. The results indicate that, under conditions of no thrust disturbance / no heating, the elastic vibrations of the aircraft in all three directions exhibit a clear convergence trend, demonstrating good stability in its aeroelastic response.

[0030] Figure 4 The dynamic responses of characteristic points of the structure under full load, half load, and no load conditions are shown. It can be seen that under no load conditions, the structure can maintain the stability of elastic vibration, and the vibration amplitude tends to converge, indicating good system stability. However, under full load and half load conditions, the dynamic response of the structure exhibits more complex coupling characteristics, especially since the rigid body displacement in the y-direction can no longer completely cover the elastic vibration. This indicates that under these load conditions, the influence of elastic vibration on the overall dynamic response becomes significant, and rigid body motion is no longer the dominant characteristic. The presence of thrust disturbance further exacerbates the aeroelastic response, reflecting the tendency of the aircraft vibration to become unstable under the combined action of load distribution and external excitation. This places higher demands on the aeroelastic servo suppression of the aircraft during mission execution.

[0031] Figure 5 This presents the thermo-aeroelastic response under a 200-second heating load, categorized as aerothermal-heat conduction. Although aerodynamic heating leads to a decrease in the natural frequencies of the structure's thermal modes compared to the unheated condition, the thermo-aeroelastic response exhibits non-negligible high-frequency vibrations. This is attributed to the influence of thermal bending moments and thermal stresses caused by the thermal load. Furthermore, rigid body displacements are more significant under the thermal load condition, exacerbating the coupling between rigid body motion and elastic vibrations. This indicates that under extreme thermal environments, the aeroelastic response of an aircraft must comprehensively consider high-frequency vibrations, coupling mechanisms, and nonlinear characteristics to ensure structural stability and reliable flight performance.

[0032] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. An aeroelastic modeling method for a boost-recharge glider, characterized in that, The aeroelastic modeling includes aeroelastic response modeling under rigid-elastic coupling, thermo-aeroelastic response modeling under rigid-elastic coupling, and aeroelastic response modeling under thrust disturbance, specifically including the following steps: Step 1: Solving and modeling the aeroelastic response under rigid-elastic coupling. A model for solving the aeroelastic response of an aircraft in a complex external flow field is established by using aeroelastic response modeling techniques that combine the finite element method with computational fluid dynamics. The structural module considers the coupling of rigid body motion and elastic vibration, allowing the aircraft to move freely in all six translational and rotational degrees of freedom, enabling the aircraft to translate and rotate freely under airflow excitation. The flow field module uses computational fluid dynamics to perform unsteady numerical simulations of the initial flow field mesh; The coupling interface between the flow field module and the structure module is modeled by dynamic meshing and data interpolation to obtain the transient displacement response of the aircraft monitoring point, i.e. the aeroelastic response under rigid-elastic coupling. Step 2: Solving and modeling the thermo-aeroelastic response under rigid-elastic coupling. Based on the aeroelastic response solution model established in step one, aeroelastic response solution modeling considering thermal effects is performed. Specifically, this includes: First, a thermal module solution model is established using a loosely coupled time-domain propagation method, which includes aerodynamic heating and heat conduction solutions. After obtaining the temperature field, the temperature change is regarded as an approximately steady-state process. Heat conduction and thermal stress effects are loaded into the aeroelastic response model. At the same time, the change of structural properties with temperature is modeled using a linear degradation model to achieve thermal aeroelastic response solution modeling under the influence of thermal effects, thus obtaining the thermal aeroelastic response of the aircraft under rigid-elastic coupling. Step 3: Modeling and solving the aeroelastic response of rigid-elastic coupling under thrust disturbance. Based on the aeroelastic response solution model established in step one, the aeroelastic response solution model considering thrust disturbance is carried out. First, the dynamic effects of combustion instability and thrust direction deviation on the aircraft structure are simulated, and the thrust disturbance is decomposed into two directions, including axial force and lateral force. Second, the thrust oscillation amplitude, frequency and phase angle are determined, and the thrust disturbance is simulated in the form of harmonics and introduced into the boundary conditions of the aircraft engine nacelle section. Finally, the aeroelastic response solution model under thrust disturbance is realized, and the aeroelastic response of rigid-elastic coupling under thrust disturbance of the aircraft is obtained.

2. The aeroelastic modeling method for a boost-recharge glider as described in claim 1, characterized in that, In step one, the numerical solution of the structural module is based on the structural dynamics equations: ， in, These are the structural mass matrix, damping matrix, and stiffness matrix, respectively. For the resultant external force of the structure, This represents its acceleration, velocity, and displacement; simultaneously, considering rigid-elastic coupling, the rigid body's degrees of freedom for motion and elastic vibration are taken into account in the structural governing equations: ， ， in, and These are the three translational velocities and three rotational angular velocities in the body coordinate system; and The aerodynamic forces and torques generated by rigid body motion; and The aerodynamic forces and torques generated by elastic vibration; m and These are the mass and inertia of the aircraft, respectively. It is the acceleration due to gravity; This is the transformation matrix from the ground coordinate system to the body coordinate system; The rigid body motion degrees of freedom mentioned above are the three-dimensional translation and rotation around the axis of the aircraft's center of mass. A spatial solution format is constructed using finite element discretization technology, and the solution is obtained by combining it with a time integration algorithm to obtain the rigid-elastic coupled displacement response of the aircraft under aerodynamic loads.

3. The aeroelastic modeling method for a boost-recharge glider as described in claim 1, characterized in that, The calculation of the flow field module in step one is based on the numerical solution of the three-dimensional Reynolds-averaged three-dimensional Navier-Stokes equations: ， in, For the solution vector, Let represent density, and u, v, w represent velocities in the three directions, respectively. For fluid domain The boundary, The outward normal vector of the boundary, the vector flux From convection vector flux With viscous vector flux Composed of two parts ; In terms of numerical solution, the above fluid control equations are spatially discretized using the finite volume method and temporally discretized using the implicit second-order backward Euler method, employing the SST method. Turbulence models are used to obtain aerodynamic distribution.

4. The aeroelastic modeling method for a boost-recharge glider as described in claim 1, characterized in that, The coupling modeling process in step one includes: (1) Solve the flow field module to determine the nodal pressure distribution and flow field characteristics at the next moment; (2) The obtained pressure distribution is transmitted to the structural module using an interpolation algorithm; (3) Under the newly established structural module boundary conditions, solve the structural dynamic equations to predict the structural deformation at the next moment; (4) Using dynamic mesh technology, the latest structural deformation information is fed back to the flow field module, thereby completing a complete coupled iteration step; The entire coupling process follows this data exchange process, iterating until the preset computation time is reached.

5. The aeroelastic modeling method for a boost-recharge glider as described in claim 1, characterized in that, The heat conduction in step two is described by Fourier's law, and the three-dimensional transient heat conduction equation based on the isotropic assumption is as follows: ， in, For heat flux density, The thermal conductivity of the material This represents the temperature distribution.

6. The aeroelastic modeling method for a boost-recharge glider as described in claim 1, characterized in that: The time-domain propagation of the thermal module in step two adopts a loosely coupled method, and its solution process mainly includes the following steps: (1) Calculation of heat conduction inside the structure; (2) Interpolate the calculated object surface temperature to the flow field grid; (3) Use a computational fluid dynamics solver to perform calculations for aerodynamic heating; (4) Interpolate the surface heat flux density to the heat conduction calculation grid. In the aerothermal module, the thermal boundary conditions of the coupled boundary are given temperature boundary conditions; while in the structural heat conduction calculation, the coupled boundary is given heat flux boundary conditions. For a given temperature boundary condition, the following conditions are satisfied on the boundary: ， For a given heat flux boundary condition, then at the boundary: Where T is temperature, q is heat flux, w represents the wall surface, t represents time, and x, y, z represent position coordinates.

7. The aeroelastic modeling method for a boost-recharge glider as described in claim 1, characterized in that, In step two, when considering thermal effects, it is assumed that the material properties exhibit a linear trend within the temperature range. Therefore, the material properties C (elastic modulus and coefficient of linear expansion) at a specific temperature T can be expressed in the following form: ， in, It is a constant characterizing the rate at which material properties change with temperature.

8. The aeroelastic modeling method for a boost-recharge glider as described in claim 1, characterized in that, The thrust disturbance in step three is modeled using sinusoidal thrust oscillations and lateral disturbance forces to simulate high-frequency disturbance effects that may occur during engine operation, employing harmonic function forms: , in, For steady-state thrust, For the disturbance amplitude, For the perturbation frequency, The phase angle is one factor. On the other hand, a slight deviation in the direction of engine thrust may be caused by factors such as structural vibration, uneven flow in the pipeline, or nozzle sway, which in turn can lead to lateral loads on the aircraft. Lateral disturbance force can be expressed as: ， in, The amplitude of the lateral disturbance force. Its dominant frequency, This is the phase angle; this term applies to the location of the aircraft's engine nacelle and affects the overall vibration response.

9. A method for aeroelastic modeling and analysis of a boost-recharge glider as described in any one of claims 1-8, characterized in that, This includes the rigid coupling signal in the aeroelastic response solution models under rigid-elastic coupling, the thermo-aeroelastic response solution models under rigid-elastic coupling, and the aeroelastic response solution models under thrust disturbance at the monitoring points on the aircraft. The decoupling technique for this signal combines empirical mode decomposition and filtering decoupling techniques, specifically: First, the signal is decomposed into different frequency components using the empirical mode decomposition method to obtain the cutoff frequency. Then, a high-pass filter is used to extract the high-frequency elastic vibration component using this cutoff frequency, while the low-frequency rigid body motion component is separated. Finally, the rigid body motion and elastic vibration of the displacement signal are effectively decoupled.

10. The aeroelastic modeling and analysis method for a boost-recharge glider as described in claim 9, characterized in that: The empirical mode decomposition mainly involves signal decomposition through the following steps: (1) Identification signal The upper envelope is obtained by fitting all local maxima and local minima using cubic spline interpolation. and lower envelope To cover the entire data range; (2) Calculate the local mean ; (3) Calculate the difference between the signal and the mean: ,like If the intrinsic mode function condition is met, i.e., the difference between the number of zero-crossing points and the number of extreme points does not exceed 1, and the local mean is close to 0, then it is considered a first-order intrinsic mode function, denoted as . Otherwise, for Continue iterating the same envelope fitting process until the intrinsic mode function condition is met; set the cutoff frequency of the high-pass filter by the frequency of the empirical mode decomposition signal. .

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