Calculation method for nonlinear flight mechanics fitting modal characteristic parameters of connected-wing aircraft

Abstract

This invention discloses a method for calculating the modal characteristic parameters of a connected-wing aircraft using nonlinear flight mechanics fitting, relating to the field of flight mechanics technology. The method includes the following steps: S1, preparation for full nonlinear modal characteristic analysis of the connected-wing aircraft, including initial flight state parameters, aerodynamic database, disturbance signals, time scale, and a full nonlinear flight dynamics model; S2, solving the nonlinear dynamics simulation of the connected-wing aircraft using an iterative simulation method; S3, fitting the modal characteristic parameters using the least squares method based on the simulation results to obtain the final processed result. This invention's method for calculating the modal characteristic parameters of a connected-wing aircraft using nonlinear flight mechanics fitting improves the accuracy of flight performance analysis, enhances the reliability of nonlinear dynamics simulation, and provides design basis and technical support for aircraft design and flight control.

Description

Technical Field

[0001] This invention relates to the field of flight mechanics technology, and in particular to a method for calculating the characteristic parameters of the nonlinear flight mechanics fitting mode of a connected-wing aircraft. Background Technology

[0002] Modal characteristic parameters are crucial features describing the dynamic stability of an aircraft system within its flight dynamics. These parameters depict the dynamic behavior patterns of an aircraft in response to external disturbances or control inputs. From a motion synthesis perspective, the motion of an aircraft after being disturbed can be composed of a series of simple motions superimposed; these simple motions are called typical modal motions. Generally, for conventionally configured aircraft, longitudinal motion mainly consists of two modes: long-period and short-period. Lateral motion mainly consists of three modes: roll, Dutch roll, and propulsion. Of course, different aircraft modes may be coupled. Typical modal motions correspond one-to-one with solutions to the characteristic equations in the system matrix. Therefore, modal characteristic parameters are derived from the linear dynamics system matrix, representing a way for linear dynamics systems to describe motion laws. Thus, modal characteristic parameters are defined in linear dynamics systems, and by solving the characteristic equations of the system matrix, the modal characteristic parameters and modal characteristic structure of the flight dynamics system can be obtained.

[0003] When studying flight stability and controllability issues, the full-scale nonlinear flight dynamics model is more adaptable than the linear dynamics model. In the derivation of the dynamics model, the full-scale nonlinear flight dynamics model of the aircraft can be obtained based on the momentum theorem and the angular momentum theorem. Under the quasi-steady assumption, when studying flight stability and controllability issues, the aircraft motion can be divided into reference motion and perturbation motion. Through the small perturbation linearization method, a small perturbation linear dynamics model for this quasi-steady motion can be obtained. Therefore, the linear dynamics model is a simplified form of the full-scale nonlinear dynamics model under certain basic assumptions.

[0004] Under typical flight conditions, for connected-wing aircraft, when using a nonlinear flight dynamics model to describe flight motion, it is impossible to directly calculate the modal characteristic parameters describing flight dynamics. Theoretically, for a given flight state, as long as the linearization condition is met, the modal characteristic parameters describing flight dynamics should be an inherent property of the aircraft, independent of whether the dynamics model providing the research basis is linear. Therefore, under identical simulation conditions, based on the results of linear and nonlinear dynamics simulations and according to the composition of the solution of the linear dynamics model, a method for calculating the nonlinear flight mechanics fitting modal characteristic parameters of connected-wing aircraft can be presented. Summary of the Invention

[0005] The purpose of this invention is to propose a method for calculating the modal characteristic parameters of a winged aircraft using nonlinear flight mechanics fitting. This method improves the accuracy of flight performance analysis, enhances the reliability of nonlinear dynamics simulation, and achieves accurate modal characteristic parameter fitting. It provides a basis for aircraft design and supports flight control, reduces the number of flight tests and costs, and promotes the development of winged aircraft technology and improves its overall performance.

[0006] To achieve the above objectives, this invention proposes a method for calculating the characteristic parameters of the nonlinear flight mechanics fitting mode of a connected-wing aircraft. The specific steps are as follows:

[0007] Step S1: Preparation for full nonlinear modal characteristic analysis of the connected-wing aircraft, including initial flight state parameters, aerodynamic database, disturbance signals, time scale, and full nonlinear flight dynamics model;

[0008] Step S2: Solve the nonlinear dynamics of the connected-wing aircraft through iterative simulation.

[0009] Step S3: Fit the modal characteristic parameters using the least squares method.

[0010] Preferably, the specific preparation steps in step S1 are as follows:

[0011] Step S11: Determine the initial flight state parameters of the connected-wing aircraft;

[0012] Step S12: Determine the aerodynamic database of all components' aerodynamic forces and aerodynamic moments under different flight conditions of the coupled-wing aircraft;

[0013] Step S13: Determine the disturbance signal of the connected-wing aircraft to be studied;

[0014] Step S14: Determine the time scale of the iterative simulation, i.e., determine the initial moment of the simulation. Length of time and the simulation end time ;

[0015] Step S15: Derive the full nonlinear flight dynamics model of the connected-wing aircraft.

[0016] Preferably, in step S11, the initial flight state parameters include flight vacuum speed. Angle of attack Flight sideslip angle Flight altitude It also includes the current aircraft weight. Ground velocity vector, attitude angle, and three-axis angular rate.

[0017] Preferably, in step S12, the aerodynamic database includes velocity. Angle of attack Flight sideslip angle Flight altitude Aerodynamic data information in four dimensions, as well as aerodynamic control data on changes in aerodynamic forces and torques caused by the deflection of all control surfaces of the aircraft, and dynamic aerodynamic data under dynamic changes of the aircraft.

[0018] Preferably, in step S13, the longitudinal dynamics process is studied using the elevator. Pulse signals and step signals; studying the lateral dynamics process using ailerons or rudder Pulse signals and step signals.

[0019] Preferably, in step S15, the dynamic model includes three force equations (drag, side force, and lift), three moment equations (roll, pitch, and yaw), three attitude equations, and two aerodynamic angle equations (angle of attack and sideslip angle). The equation set contains 11 variables and a total of 11 equations, and the equation set is closed.

[0020] Preferably, in step S2, the specific steps for solving the nonlinear dynamics simulation are as follows:

[0021] Step S21: Select a discrete iterative solution algorithm for the differential equation, including Runge-Kutta algorithm, Newton iteration method and bisection method;

[0022] Step S22: Select an appropriate time step;

[0023] Step S23: Substitute the initial flight state parameters and basic physical parameters into the solution algorithm, write the aerodynamic database, control aerodynamic data and dynamic aerodynamic data interpolation and retrieval program, and perform discrete solution through iterative simulation method;

[0024] Step S24: Adjust the time step, compare the simulation results from multiple simulations, conduct simulation time step independence analysis, and use error calculation methods to determine whether the simulation results are time step independent. If the simulation time step is not independent, adjust the time step and iterate again to solve the problem. If it is independent, proceed to the next step.

[0025] Preferably, in step S3, the steps for processing the fitted modal feature parameters are as follows:

[0026] Step S31: Based on the simulation results, separate the simulation data of the angle of attack representing longitudinal disturbance motion and the simulation data of the sideslip angle representing lateral disturbance motion;

[0027] Step S32: Derive the state equations of the linear dynamics system of the connected-wing aircraft; When the connected-wing aircraft is subjected to disturbances, its modal characteristic parameters represent the aircraft's motion state. As long as the small disturbance linearization condition is met, the linear simulation results and nonlinear simulation results should be quite close. Through derivation, its linear dynamics equations are as follows:

[0028] ;

[0029] in, For state variables, longitudinal motion , This is the change in airspeed. This represents the change in angle of attack. The change in pitch rate. The change in pitch angle; lateral motion. , This represents the change in sideslip angle. This is the change in roll rate. This represents the change in yaw rate. This represents the change in roll angle; A For the system matrix, B For the control matrix, C For the output matrix, D For direct matrix transmission; To control the quantity, longitudinal movement , This represents the change in elevator deflection angle. Change in throttle opening, lateral motion , This represents the change in aileron deflection angle. This represents the change in rudder deflection angle; State vector Time differential, This is the output vector;

[0030] Step S33: Determine the form of the time-domain solution of the linear dynamics system; for the linear dynamics equations, both the state vector and the matrix are defined in the real number space, under the initial constant-length level flight state. X =0, performing a Laplace transform on the above equation, we get:

[0031] ;

[0032] in, This is the input matrix in the state equation. The input vector for the state equation. The vector of state variables in the state equation. For Laplace variation notation, IGiven the identity matrix, the theoretical solution under the initial state conditions is:

[0033] ;

[0034] in, Let be the integral symbol; further expansion using Euler's formula yields the following solution:

[0035] ;

[0036] in, and These are the amplitudes of two modes, and These are the real parts of the eigenvalues ​​of the two modes, and These are the imaginary parts of the eigenvalues ​​of the two modes, and These are the initial phase angles of the two modes, It is a constant. The change in state parameters;

[0037] Step S34: The least squares method is used to numerically fit the solution of the full nonlinear flight dynamics equation of the underwing layout under disturbance. The fitting parameters are the longitudinal and lateral motion aerodynamic angles. During the fitting process, it is determined whether the fitting error is within the required range. If it is, proceed to the next step. If it is not, return to the fitting process iteration.

[0038] Step S35: Result processing, obtaining the dynamic change functions of the angle of attack and sideslip angle of the lower-wing configuration aircraft under the influence of pulse disturbance.

[0039] Therefore, this invention proposes a method for calculating the characteristic parameters of the nonlinear flight mechanics fitting mode of a connected-wing aircraft, the advantages of which are as follows:

[0040] (1) The method for calculating the nonlinear flight mechanics fitting modal characteristic parameters of the connected-wing aircraft proposed in this invention can accurately consider the aerodynamic forces and aerodynamic moments of the aircraft under different flight states by calculating the nonlinear flight mechanics fitting modal characteristic parameters of the connected-wing aircraft. It combines a multi-dimensional aerodynamic database including speed, flight angle of attack, flight sideslip angle, flight altitude, as well as dynamic aerodynamic data and control aerodynamic data, avoiding the flight performance analysis errors that may be caused by incomplete or inaccurate data in traditional methods, and making the performance prediction of the aircraft under various flight conditions more accurate.

[0041] (2) The method for calculating the nonlinear flight mechanics fitting modal characteristic parameters of the connected-wing aircraft proposed in this invention provides a scientific basis for the design of connected-wing aircraft through accurate calculation of nonlinear flight mechanics fitting modal characteristic parameters and reliable simulation results. It can help designers better understand the dynamic characteristics of the aircraft, optimize the layout, structure and performance of the aircraft, and improve the design quality and flight performance of the aircraft.

[0042] (3) The method for calculating the nonlinear flight dynamics fitting modal characteristic parameters of the connected-wing aircraft proposed in this invention considers the elevator pulse signal and step signal in the longitudinal dynamic process, as well as the aileron or rudder pulse signal or step signal in the lateral dynamic process as disturbance signals. This helps to more comprehensively evaluate the flight performance of the aircraft under different external disturbances, simulate various disturbances that may occur in actual flight, and provide more realistic data support for the stability and maneuverability analysis of the aircraft.

[0043] The technical solution of the present invention will be further described in detail below with reference to the accompanying drawings and embodiments. Attached Figure Description

[0044] Figure 1 This is a flowchart illustrating the overall analysis process of the method for calculating the nonlinear flight mechanics fitting modal characteristic parameters of the connected-wing aircraft according to the present invention. Detailed Implementation

[0045] To make the technical solutions, advantages, and objectives of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below. The described embodiments are only some, not all, of the embodiments of the present invention. All other embodiments obtained by those skilled in the art based on the described embodiments of the present invention without creative effort are within the protection scope of this application.

[0046] Unless otherwise defined, the technical or scientific terms used in this invention shall have the ordinary meaning as understood by one of ordinary skill in the art to which this invention pertains.

[0047] like Figure 1 As shown, this invention provides a method for calculating the characteristic parameters of the nonlinear flight mechanics fitting mode of a connected-wing aircraft. The specific steps are as follows:

[0048] Step S1: Preparation for full nonlinear modal characteristic analysis of the coupled-wing aircraft, specifically including:

[0049] Step S11: Determine the initial flight state parameters of the coupled-wing aircraft, including flight vacuum speed. Angle of attack Flight sideslip angle Flight altitude It also includes the current aircraft weight. Ground velocity vector, attitude angle, and three-axis angular rate. The values ​​of the initial parameters are shown in Table 1.

[0050] Table 1 Initial parameters for full-scale nonlinear simulation of a coupled-wing aircraft

[0051]

[0052] Step S12: Determine the aerodynamic database of all components' aerodynamic forces and aerodynamic moments under different flight conditions of the coupled-wing aircraft. The aerodynamic database includes velocity... Angle of attack Flight sideslip angle Flight altitude Aerodynamic data information in four dimensions, as well as aerodynamic control data on changes in aerodynamic forces and torques caused by the deflection of all control surfaces of the aircraft, and dynamic aerodynamic data under dynamic changes of the aircraft.

[0053] Step S13: Determine the disturbance signal of the connected-wing aircraft to be studied. To study the longitudinal dynamics, take the pulse signal and step signal of the elevator; to study the lateral dynamics, take the pulse signal and step signal of the aileron or rudder.

[0054] Step S14: Determine the time scale of the iterative simulation, i.e., determine the initial moment of the simulation. Length of time and the simulation end time ;

[0055] Step S15: Derive the full nonlinear flight dynamics model of the coupled-wing aircraft. The dynamics model includes three force equations (drag, side force, and lift), three moment equations (roll, pitch, and yaw), three attitude equations, and two aerodynamic angle equations (angle of attack and sideslip). The equation set contains 11 variables and a total of 11 equations. The equation set is closed, and the formulas for the 11 equations are as follows:

[0056] ;

[0057] ;

[0058] ;

[0059] ;

[0060] ;

[0061] in, For the angle of attack, Sideslip angle, These are the x, y, and z axis projections of the airspeed vector in the body axis system, respectively. These are the angular velocities about the body axis, respectively. These are the aircraft attitude angles, It is the acceleration due to gravity. , , These are drag, lateral force, and lift acting on the aircraft, respectively. These are the rolling moment, pitching moment, and yaw moment acting on the aircraft, respectively. For the weight of the aircraft, , , These are engine thrust, thrust angle, and thrust eccentricity, respectively. , , These are the aircraft's moments of inertia, Let be the inertial product of the aircraft about the xz plane.

[0062] Step S2: Solve the nonlinear dynamics of the connected-wing aircraft through iterative simulation. The specific steps are as follows:

[0063] Step S21: Select a discrete iterative solution algorithm for the differential equation, including Runge-Kutta algorithm, Newton iteration method and bisection method;

[0064] Step S22: Select an appropriate time step;

[0065] Step S23: Substitute the initial flight state parameters and basic physical parameters into the solution algorithm, write the aerodynamic database, control aerodynamic data and dynamic aerodynamic data interpolation and retrieval program, and perform discrete solution through iterative simulation method;

[0066] Step S24: Adjust the time step, compare the simulation results from multiple simulations, conduct simulation time step independence analysis, and use error calculation methods to determine whether the simulation results are time step independent. If the simulation time step is not independent, adjust the time step and iterate again to solve the problem. If it is independent, proceed to the next step.

[0067] Step S3: Processing the fitted modal feature parameters, the steps are as follows:

[0068] Step S31: Based on the simulation results, separate the simulation data of the angle of attack representing longitudinal disturbance motion and the simulation data of the sideslip angle representing lateral disturbance motion;

[0069] Step S32: Derive the state equations of the linear dynamics system of the connected-wing aircraft; When the connected-wing aircraft is subjected to disturbances, its modal characteristic parameters represent the aircraft's motion state. As long as the small disturbance linearization condition is met, the linear simulation results and nonlinear simulation results should be quite close. Through derivation, its linear dynamics equations are as follows:

[0070] ;

[0071] in, For state variables, longitudinal motion , This is the change in airspeed. This represents the change in angle of attack. The change in pitch rate. The change in pitch angle; lateral motion. , This represents the change in sideslip angle. This is the change in roll rate. This represents the change in yaw rate. This represents the change in roll angle; A For the system matrix, B For the control matrix, C For the output matrix, D For direct matrix transmission; To control the quantity, longitudinal movement , This represents the change in elevator deflection angle. Change in throttle opening, lateral motion , This represents the change in aileron deflection angle. This represents the change in rudder deflection angle; State vector Time differential, This is the output vector;

[0072] Step S33: Determine the form of the time-domain solution of the linear dynamics system; for the linear dynamics equations, both the state vector and the matrix are defined in the real number space, under the initial constant-length level flight state. X =0, performing a Laplace transform on the above equation, we get:

[0073] ;

[0074] in, This is the input matrix in the state equation. The input vector for the state equation. The vector of state variables in the state equation. For Laplace variation notation, I Given the identity matrix, the theoretical solution under the initial state conditions is:

[0075] ;

[0076] in, Let be the integral symbol; further expansion using Euler's formula yields the following solution:

[0077] ;

[0078] in, and These are the amplitudes of two modes, and These are the real parts of the eigenvalues ​​of the two modes, and These are the imaginary parts of the eigenvalues ​​of the two modes, and These are the initial phase angles of the two modes, It is a constant. The change in state parameters;

[0079] Step S34: The least squares method is used to numerically fit the solution of the full nonlinear flight dynamics equation of the underwing layout under disturbance. The fitting parameters are the longitudinal and lateral motion aerodynamic angles. During the fitting process, it is determined whether the fitting error is within the required range. If it is, proceed to the next step. If it is not, return to the fitting process iteration.

[0080] Step S35: Result processing. The dynamic change functions of the angle of attack and sideslip angle of the lower-wing configuration aircraft under the influence of the pulse disturbance are obtained, and the formulas are as follows:

[0081] ;

[0082] ;

[0083] The resulting fitted oscillation modes, fitted oscillation frequencies, and fitted eigenvalues ​​are shown in Tables 2 and 3, respectively. This significantly improves the accuracy of flight performance analysis, enhances the reliability of nonlinear dynamics simulation, and achieves accurate fitting of modal characteristic parameters.

[0084] Table 2 Oscillation parameters fitted to longitudinal motion

[0085]

[0086] Table 3 Oscillation parameters fitted by lateral motion

[0087]

[0088] Therefore, this invention provides a method for calculating the nonlinear flight mechanics fitting modal characteristic parameters of a connected-wing aircraft. Through accurate simulation and calculation, it can predict the performance and response of the aircraft under various flight conditions to a certain extent, providing a more economical and efficient technical means for the research and development and use of the aircraft. At the same time, since it takes into account complex nonlinear dynamics and various disturbances, it can better cope with the complex and ever-changing environment and various possible problems in actual flight, thereby improving the overall performance and competitiveness of the connected-wing aircraft.

[0089] 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 preferred embodiments, those skilled in the art should understand that modifications or equivalent substitutions can still be made to the technical solutions of the present invention, and these modifications or equivalent substitutions cannot cause the modified technical solutions to deviate from the spirit and scope of the technical solutions of the present invention.

Claims

1. A method for calculating the characteristic parameters of the nonlinear flight mechanics fitting mode of a connected-wing aircraft, characterized in that, The specific steps are as follows: Step S1: Preparation for full nonlinear modal characteristic analysis of the connected-wing aircraft, including initial flight state parameters, aerodynamic database, disturbance signals, time scale, and full nonlinear flight dynamics model; Step S2: Solve the nonlinear dynamics of the connected-wing aircraft through iterative simulation. Step S3: Fit the modal characteristic parameters using the least squares method; In step S3, the steps for processing the fitted modal feature parameters are as follows: Step S31: Based on the simulation results, separate the simulation data of the angle of attack representing longitudinal disturbance motion and the simulation data of the sideslip angle representing lateral disturbance motion; Step S32: Derive the state equations of the linear dynamics system of the coupled-wing aircraft; When the coupled-wing aircraft is subjected to disturbances, its modal characteristic parameters represent the aircraft's motion state. As long as the small disturbance linearization condition is satisfied, its linear dynamics equations are as follows: ； in, For state variables, longitudinal motion , This is the change in airspeed. This represents the change in angle of attack. The change in pitch rate. The change in pitch angle; lateral motion. , This represents the change in sideslip angle. This is the change in roll rate. This represents the change in yaw rate. This represents the change in roll angle; A For the system matrix, B For the control matrix, C For the output matrix, D For direct matrix transmission; To control the quantity, longitudinal movement , This represents the change in elevator deflection angle. Change in throttle opening, lateral motion , This represents the change in aileron deflection angle. This represents the change in rudder deflection angle; State vector Time differential, This is the output vector; Step S33: Determine the form of the time-domain solution of the linear dynamics system; for the linear dynamics equations, both the state vector and the matrix are defined in the real number space, under the initial constant-length level flight state. X =0, performing a Laplace transform on the above equation, we get: ； in, This is the input matrix in the state equation. The input vector for the state equation. The vector of state variables in the state equation. For Laplace variation notation, I Given the identity matrix, the theoretical solution under the initial state conditions is: ； in, Let be the integral symbol; further expansion using Euler's formula yields the following solution: ； in, and These are the amplitudes of two modes, and These are the real parts of the eigenvalues ​​of the two modes, and These are the imaginary parts of the eigenvalues ​​of the two modes, and These are the initial phase angles of the two modes, It is a constant. The change in state parameters; Step S34: The least squares method is used to numerically fit the solution of the full nonlinear flight dynamics equation of the underwing layout under disturbance. The fitting parameters are the longitudinal and lateral motion aerodynamic angles. During the fitting process, it is determined whether the fitting error is within the required range. If it is, proceed to the next step. If it is not, return to the fitting process iteration. Step S35: Result processing, obtaining the dynamic change functions of the angle of attack and sideslip angle of the lower-wing configuration aircraft under the influence of pulse disturbance.

2. The method for calculating the nonlinear flight mechanics fitting modal characteristic parameters of a connected-wing aircraft according to claim 1, characterized in that, In step S1, the specific preparation steps are as follows: Step S11: Determine the initial flight state parameters of the connected-wing aircraft; Step S12: Determine the aerodynamic database of all components' aerodynamic forces and aerodynamic moments under different flight conditions of the coupled-wing aircraft; Step S13: Determine the disturbance signal of the connected-wing aircraft to be studied; Step S14: Determine the time scale of the iterative simulation, i.e., determine the initial moment of the simulation. Length of time and the simulation end time ; Step S15: Derive the full nonlinear flight dynamics model of the connected-wing aircraft.

3. The method for calculating the nonlinear flight mechanics fitting modal characteristic parameters of a connected-wing aircraft according to claim 2, characterized in that, In step S11, the initial flight state parameters include flight vacuum speed. Angle of attack Flight sideslip angle Flight altitude It also includes the current aircraft weight. Ground velocity vector, attitude angle, and three-axis angular rate.

4. The method for calculating the nonlinear flight mechanics fitting modal characteristic parameters of a connected-wing aircraft according to claim 2, characterized in that, In step S12, the aerodynamic database includes velocity. Angle of attack Flight sideslip angle Flight altitude Aerodynamic data information in four dimensions, as well as aerodynamic control data on changes in aerodynamic forces and torques caused by the deflection of all control surfaces of the aircraft, and dynamic aerodynamic data under dynamic changes of the aircraft.

5. The method for calculating the nonlinear flight mechanics fitting modal characteristic parameters of a connected-wing aircraft according to claim 2, characterized in that, In step S13, the longitudinal dynamics process is studied using the elevator. Pulse signals and step signals; studying the lateral dynamics process using ailerons or rudder Pulse signals and step signals.

6. The method for calculating the nonlinear flight mechanics fitting modal characteristic parameters of a connected-wing aircraft according to claim 2, characterized in that, In step S15, the dynamic model includes three force equations (drag, side force, and lift), three moment equations (roll, pitch, and yaw), three attitude equations, and two aerodynamic angle equations (angle of attack and sideslip angle). The equation set contains 11 variables and a total of 11 equations, and the equation set is closed.

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