Method for calculating nonlinear flight mechanics fitting modal characteristic parameters of connected-wing aircraft

Abstract

The invention discloses a method for calculating nonlinear flight mechanics fitting modal characteristic parameters of a linked-wing aircraft, and relates to the technical field of flight mechanics. The method comprises the following steps: S1, carrying out analysis preparation on full-amount nonlinear modal characteristics of a connected-wing aircraft, including initial flight state parameters, a pneumatic database, a disturbance signal, a time scale and a full-amount nonlinear flight dynamics model; s2, performing nonlinear dynamic simulation solution of the connected-wing layout aircraft through an iterative simulation method; and S3, according to a simulation result, fitting modal characteristic parameters by adopting a least square method to obtain a final processing result. According to the method for calculating the nonlinear flight mechanical fitting modal characteristic parameters of the connected-wing aircraft, the accuracy of flight performance analysis is improved, the reliability of nonlinear dynamic simulation is enhanced, and a design basis and technical support are provided for aircraft design and flight control.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of flight dynamics, in particular to a fitting modal characteristic parameter calculation method for a tandem wing aircraft nonlinear flight dynamics. BACKGROUND

[0002] The modal characteristic parameter is an important characteristic parameter of the body dynamics of the aircraft flight dynamics, which describes the dynamic stability of the system. The parameter describes the dynamic behavior mode of the aircraft in response to external disturbance or control input. From the perspective of motion synthesis, the motion of the aircraft after disturbance can be superimposed by a series of simple motions, which are called typical modal motions. Generally, for a conventional layout aircraft, the longitudinal motion mainly has two modes of long period and short period, the lateral motion mainly has three modes of roll, Dutch roll and spiral, and of course the modes of different characteristics may be coupled. The typical modal motion is one-to-one corresponding to the solution of the characteristic equation in the system matrix, so the modal characteristic parameter is derived from the linear dynamics system matrix, which is a way to describe the motion law of the linear dynamics system. Therefore, the modal characteristic parameter is defined in the linear dynamics system, and the modal characteristic parameter and the modal characteristic structure of the flight dynamics system can be obtained by solving the system matrix characteristic equation.

[0003] When studying flight stability and maneuverability, the full nonlinear flight dynamics model has better adaptability than the linear dynamics model. In the process of deriving the dynamics model, the aircraft full nonlinear flight dynamics model can be obtained according to the momentum theorem and the moment of momentum theorem. Under the quasi-constant assumption, when studying flight stability and maneuverability, the aircraft motion can be divided into reference motion and disturbance motion, and through the small disturbance linearization method, a small disturbance linear dynamics model for the quasi-constant motion can be obtained. Therefore, the linear dynamics model is a simplified form of the full nonlinear dynamics model based on certain basic assumptions.

[0004] Under the typical flight state, for a tandem wing aircraft, the modal characteristic parameter describing the flight dynamics behavior cannot be directly calculated when the nonlinear flight dynamics model is used to describe the flight motion. In theory, for a given flight state, as long as the linearization condition is met, the modal characteristic parameter describing the flight dynamics behavior should be an inherent property of the aircraft, and it is irrelevant to whether the dynamics model providing the research basis is a linear dynamics model. Therefore, under the same simulation conditions, based on the linear dynamics simulation and nonlinear dynamics simulation results, a fitting modal characteristic parameter calculation method for a tandem wing aircraft nonlinear flight dynamics can be given according to the composition of the linear dynamics model solution. SUMMARY

[0005] The application aims to provide a fitting mode characteristic parameter calculation method for nonlinear flight dynamics of a wing-in-ground effect vehicle, improve the accuracy of flight performance analysis, enhance the reliability of nonlinear dynamics simulation, realize accurate mode characteristic parameter fitting, provide a basis for vehicle design and support for flight control, reduce flight test times and costs, and promote the development and improvement of the comprehensive performance of the wing-in-ground effect vehicle.

[0006] To achieve the above-mentioned purpose, the application provides a fitting mode characteristic parameter calculation method for nonlinear flight dynamics of a wing-in-ground effect vehicle, and the specific steps are as follows: Step S1, preparation of full-quantity nonlinear mode characteristic analysis of a wing-in-ground effect vehicle, including initial flight state parameters, an aerodynamic database, disturbance signals, a time scale, and a full-quantity nonlinear flight dynamics model; Step S2, nonlinear dynamics simulation solving of the wing-in-ground effect vehicle is performed through iterative simulation; Step S3, mode characteristic parameters are fitted by using the least square method.

[0007] Preferably, in step S1, the specific preparation steps are as follows: Step S11, initial flight state parameters of the wing-in-ground effect vehicle are determined; Step S12, an aerodynamic database of aerodynamic forces and aerodynamic moments of all components of the wing-in-ground effect vehicle under different flight states is determined; Step S13, disturbance signals of the wing-in-ground effect vehicle to be studied are determined; Step S14, the time scale of iterative simulation is determined, that is, the initial time , the time length , and the end time of simulation are determined; Step S15, a full-quantity nonlinear flight dynamics model of the wing-in-ground effect vehicle is derived.

[0008] Preferably, in step S11, the initial flight state parameters include flight true airspeed , flight angle of attack , flight sideslip angle , flight altitude , current aircraft weight , ground speed vector, attitude angle, and three-axis angular rate.

[0009] Preferably, in step S12, the aerodynamic database includes speed , flight angle of attack , flight sideslip angle , flight altitude Four-dimension aerodynamic data information, and aerodynamic control data of all control surfaces deflection of the aircraft caused aerodynamic force and aerodynamic moment changes and dynamic aerodynamic data under dynamic change conditions of the aircraft.

[0010] Preferably, in step S13, the longitudinal dynamic process is studied by taking the elevator pulse signal and step signal; the lateral dynamic process is studied by taking the aileron or rudder pulse signal and step signal.

[0011] Preferably, in step S15, the dynamic model includes three force equations of drag, lift and side force, three moment equations of roll, pitch and yaw, three attitude equations and two aerodynamic angle equations of angle of attack and sideslip angle, the equation set contains 11 variables, a total of 11 equations, and the equation set is closed.

[0012] Preferably, in step S2, the specific steps of the nonlinear dynamic simulation solution are as follows: Step S21, select a differential equation discrete iterative solution algorithm, including Runge-Kutta algorithm, Newton iteration method and bisection method; Step S22, select an appropriate time step; Step S23, substitute the initial flight state parameters and basic physical parameters into the solution algorithm, write an interpolation and call program for the aerodynamic database, control aerodynamic data and dynamic aerodynamic data, and perform discrete solution by iterative simulation method; Step S24, adjust the time step, compare the simulation results, carry out simulation time step independence analysis, and use error calculation method to judge whether the simulation result has time step independence, when the simulation time step does not have independence, adjust the time step and reiterate the solution, when it has independence, go to the next step.

[0013] Preferably, in step S3, the fitting modal characteristic parameter processing steps are as follows: Step S31, according to the simulation results, separate the angle of attack simulation data representing the longitudinal disturbance motion and the sideslip angle simulation data representing the lateral disturbance motion; Step S32, derive the state equation of the linear dynamic system of the tandem wing layout aircraft; when the tandem wing layout aircraft is subjected to disturbance, its modal characteristic parameters represent the motion state of the aircraft, as long as the small disturbance linearization condition is met, the linear simulation result and the nonlinear simulation result should be relatively close, and the linear dynamic equation is derived as follows: ; Where, is the state variable, the longitudinal motion , is the change of airspeed, for the angle of attack change, for the pitch rate change, for the pitch angle change; lateral side movement , for the sideslip angle change, for the roll rate change, for the yaw rate change, for the roll angle change; A for the system matrix, B for the control matrix, C for the output matrix, D for the direct transfer matrix; for the control quantity, longitudinal movement , for the elevator deflection angle change, for the throttle opening change, lateral side movement , for the aileron deflection angle change, for the rudder deflection angle change; for the state vector time differential, for the output vector; Step S33, determine the form of the linear dynamic system time domain solution; for the linear dynamic equation, the state vector and the matrix are defined in the real number space range, and in the initial straight flight state X = 0, Laplace transform is performed on the above formula to obtain: ; wherein, is the input matrix in the state equation, is the state equation input vector, is the state variable vector in the state equation, is the Laplace change symbol, I is the unit matrix, and the theoretical solution under the initial state condition is: ; wherein, is the integral symbol; further expanded by Euler formula, the solution formula is as follows: ; wherein, and are two modal amplitudes, and are two modal characteristic value real parts, and are two modal characteristic value imaginary parts, and respectively two modal initial phase angles, is a constant, is a state parameter change amount; Step S34, using the least square method to carry out numerical fitting on the solution of the full-quantity nonlinear flight dynamics equation of the joined wing layout under the disturbance action; the fitting parameters are the longitudinal and lateral motion aerodynamic angles, in the fitting process, it is judged whether the fitting error is within the required range, if it is satisfied, the next step is entered, if it is not satisfied, it is returned to the iteration of the fitting process; Step S35, result processing, the dynamic change function of the angle of attack and the sideslip angle of the joined wing layout aircraft under the pulse disturbance action is obtained.

[0014] Therefore, the present application proposes a joined wing aircraft nonlinear flight dynamics fitting modal characteristic parameter calculation method, which has the following beneficial effects: (1) The joined wing aircraft nonlinear flight dynamics fitting modal characteristic parameter calculation method proposed by the present application can accurately consider the aerodynamic force and aerodynamic moment of the aircraft under different flight states through the calculation of the nonlinear flight dynamics fitting modal characteristic parameters of the joined wing aircraft. Combined with the multi-dimensional aerodynamic database including speed, flight attack angle, flight sideslip angle, flight height, and dynamic aerodynamic data and control aerodynamic data, the flight performance analysis error caused by the incomplete or inaccurate data of the traditional method is avoided, so that the performance prediction of the aircraft under various flight conditions is more accurate.

[0015] (2) The joined wing aircraft nonlinear flight dynamics fitting modal characteristic parameter calculation method proposed by the present application provides a scientific basis for the design of the joined wing layout aircraft through accurate nonlinear flight dynamics fitting modal characteristic parameter calculation and reliable simulation results, which 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.

[0016] (3) The joined wing aircraft nonlinear flight dynamics fitting modal characteristic parameter calculation method proposed by the present application considers the elevator pulse signal and step signal in the longitudinal dynamics process, and the aileron or rudder pulse signal or step signal in the lateral dynamics process as the disturbance signal, which helps to more comprehensively evaluate the flight performance of the aircraft under different external disturbances, simulates various disturbances that may occur in actual flight, and provides more actual situation data support for the stability and maneuverability analysis of the aircraft.

[0017] The technical solutions of the present application will be further described in detail below through the drawings and examples. BRIEF DESCRIPTION OF DRAWINGS

[0018] Figure 1The whole analysis flow chart of the wing-in-ground effect vehicle nonlinear flight dynamics fitting modal characteristic parameter calculation method. DETAILED DESCRIPTION

[0019] In order to make the technical solutions, advantages and purposes of the present application clearer, the technical solutions of the embodiments of the present application will be clearly and completely described below. The described embodiments are part of the embodiments of the present application, rather than all the embodiments. Based on the described embodiments of the present application, all other embodiments obtained by those skilled in the art without creative labor are within the protection scope of the present application.

[0020] Unless otherwise defined, the technical terms or scientific terms used in the present application shall have the usual meanings understood by those skilled in the art to which the present application belongs.

[0021] As shown in Figure 1 , the present application provides a wing-in-ground effect vehicle nonlinear flight dynamics fitting modal characteristic parameter calculation method, and the specific steps are as follows: Step S1, wing layout aircraft full quantity nonlinear modal characteristic analysis preparation, specifically including: Step S11, determining the initial flight state parameters of the wing layout aircraft, including flight true airspeed , flight angle of attack , flight sideslip angle , flight altitude , current aircraft weight , ground speed vector, attitude angle and three-axis angular rate. The numerical values of the initial parameters are shown in Table 1.

[0022] Table 1 Wing layout aircraft full quantity nonlinear simulation initial parameters

[0023] Step S12, determining the aerodynamic database of all components of the wing layout aircraft under different flight states, the aerodynamic database including aerodynamic data information of four dimensions of speed , flight angle of attack , flight sideslip angle , flight altitude , and aerodynamic control data caused by deflection of all control surfaces of the aircraft and dynamic aerodynamic data under dynamic change conditions of the aircraft.

[0024] Step S13, determining the disturbance signals of the wing layout aircraft to be studied, taking pulse signals and step signals of elevators for studying longitudinal dynamic processes; taking pulse signals and step signals of ailerons or rudders for studying lateral dynamic processes.

[0025] Step S14, determining the time scale of the iterative simulation, i.e., determining the initial time of simulation , the length of time , and the end time of simulation ; Step S15, deriving the full-quantity nonlinear flight dynamics model of the flying-wing aircraft, the dynamics model including three force equations of drag, side force, and lift, three moment equations of roll, pitch, and yaw, three attitude equations, and two aerodynamic angle equations of angle of attack and sideslip angle, the equation set containing 11 variables, a total of 11 equations, the equation set being closed, and the 11 equations of the equation set being as follows: ; ; ; ; ; wherein, is the angle of attack, is the sideslip angle, are respectively the x, y, and z axis projections of the airspeed vector under the body axis system, are respectively the angular rates around the body axis system, are respectively the aircraft attitude angles, is the gravitational acceleration, , , are respectively the drag, side force, and lift received by the aircraft, are respectively the roll moment, pitch moment, and yaw moment received by the aircraft, is the weight of the aircraft, , , are respectively the engine thrust, thrust angle, and thrust eccentricity. , , are respectively the moments of inertia of the aircraft, is the product of inertia of the aircraft about the xz plane.

[0026] Step S2, performing nonlinear dynamics simulation and solving of the flying-wing aircraft through iterative simulation, and the specific steps are as follows: Step S21, selecting a differential equation discrete iterative solving algorithm, including the Runge-Kutta algorithm, the Newton iteration method, and the bisection method; Step S22, selecting a suitable time step; Step S23, substituting the initial flight state parameters and basic physical parameters into the solving algorithm, compiling an aerodynamic database, a control aerodynamic data, and a dynamic aerodynamic data interpolation and calling program, and performing discrete solving through an iterative simulation method; Step S24, adjusting time step, comparing multiple simulation results, carrying out simulation time step independence analysis, using error calculation method to judge whether the simulation result has time step independence, when the simulation time step does not have independence, adjusting the time step to re-iterate the solution, when it has independence, entering the next step.

[0027] Step S3, fitting modal characteristic parameter processing, the steps are as follows: Step S31, according to the simulation results, the angle of attack simulation data representing longitudinal disturbance motion and the sideslip angle simulation data representing lateral disturbance motion are separated; Step S32, deriving the linear dynamics system state equation of the blended wing layout aircraft; when the blended wing layout aircraft is subjected to disturbance, the modal characteristic parameter represents the motion state of the aircraft, as long as the small disturbance linearization condition is met, the linear simulation result and the nonlinear simulation result should be relatively close, through derivation, the linear dynamics equation is as follows: ; Wherein, is a state variable, longitudinal motion , is the change amount of airspeed, is the change amount of angle of attack, is the change amount of pitch angle rate, is the change amount of pitch angle; lateral motion , is the change amount of sideslip angle, is the change amount of roll angle rate, is the change amount of yaw angle rate, is the change amount of roll angle; A is a system matrix, B is a control matrix, C is an output matrix, D is a direct transfer matrix; is a control variable, longitudinal motion , is the change amount of elevator deflection angle, is the change amount of throttle opening, lateral motion , is the change amount of aileron deflection angle, is the change amount of rudder deflection angle; is a state vector time differential, is an output vector; Step S33, determining the form of time domain solution of linear dynamics system; for linear dynamics equation, the state vector and the matrix are defined in real number space range, in the initial straight and level flight state X= 0, Laplace transform is performed on the above equation, and the following equation is obtained: ; wherein, is the input matrix in the state equation, is the input vector in the state equation, is the state variable vector in the state equation, is the Laplace transform symbol, I is the unit matrix, and the theoretical solution under the initial state condition is as follows: ; wherein, is the integral symbol; further expanded by the Euler formula, the solution formula is as follows: ; wherein, and are two modal amplitudes, and are two modal eigenvalue real parts, and are two modal eigenvalue imaginary parts, and are two modal initial phase angles, is a constant, is a state parameter change amount; Step S34, the least square method is used to perform numerical fitting on the solution of the full-quantity nonlinear flight dynamics equation of the joined wing layout under the disturbance action; the fitting parameters are the longitudinal and lateral motion aerodynamic angles, in the fitting process, it is determined whether the fitting error is within the required range, if yes, the next step is entered, if not, the fitting process is returned to iteration; Step S35, result processing, the obtained dynamic change function of the angle of attack and the sideslip angle of the joined wing layout aircraft under the pulse disturbance action is as follows: ; ; The fitted oscillation mode, the fitted oscillation frequency and the fitted eigenvalue derived therefrom are shown in Table 2 and Table 3 respectively, which significantly improves the accuracy of the flight performance analysis, enhances the reliability of the nonlinear dynamics simulation, and realizes the precise modal eigenvalue fitting.

[0028] Table 2 Longitudinal motion fitted oscillation parameters

[0029] Table 3 Lateral motion fitted oscillation parameters

[0030] Therefore, the application provides a method for calculating fitting modal characteristic parameters of a tandem wing aircraft nonlinear flight dynamics, which can predict the performance and reaction of the aircraft under various flight conditions to a certain extent through accurate simulation and calculation, and provides a more economical and efficient technical means for the research and use of the aircraft; meanwhile, the complex nonlinear dynamics and various disturbance conditions are considered, so that the complex and changeable environment and various possible problems in actual flight can be better coped with, and the comprehensive performance and competitiveness of the tandem wing aircraft are improved.

[0031] Finally, it should be noted that the above examples are only used to illustrate the technical solutions of the present application, but not to limit them. Although the present application has been described in detail with reference to the preferred embodiments, those skilled in the art should understand that the technical solutions of the present application can still be modified or replaced by equivalents, and these modifications or replacements cannot make the modified technical solutions deviate from the spirit and scope of the technical solutions of the present application.

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.

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. Duration 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.

7. 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 S2, the specific steps for solving the nonlinear dynamics simulation are as follows: Step S21: Select a discrete iterative solution algorithm for the differential equation, including Runge-Kutta algorithm, Newton iteration method and bisection method; Step S22: Select an appropriate time step; 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; 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.

8. The method for calculating the characteristic parameters of the nonlinear flight mechanics fitting mode of a connected-wing aircraft according to claim 1, characterized in that, In step S3, the steps for processing the feature parameters of the fitted mode 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 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: ； 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 movement , 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.

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