A method and apparatus for full aircraft coupled dynamics modeling applicable to coaxial twin rotor helicopters

Abstract

The application provides a full-machine coupling dynamics modeling method and device suitable for a coaxial dual-rotor helicopter, belongs to the field of helicopter dynamics, and aims at the complex component interference problem of a high-speed compound helicopter, constructs an elastic machine body model by using a modal synthesis method, combines an aerodynamic model of a horizontal and vertical tail considering slipstream interference and a rotor inflow model, and establishes a full-machine unified coupling dynamics equation. The method can effectively quantify the influence of the slipstream of a thrust propeller on the efficiency of a horizontal and vertical tail and the feedback mechanism of the elastic deformation of a machine body on a rotor, solves the problem of insufficient precision of a traditional single model under complex working conditions, and provides a high-confidence analysis tool for the dynamics design of a coaxial dual-rotor helicopter.

Description

Technical Field

[0001] This invention belongs to the interdisciplinary field of helicopter flight dynamics, structural dynamics and aerodynamics, and specifically relates to a whole-aircraft coupled dynamics modeling method and device suitable for coaxial twin-rotor helicopters. Background Technology

[0002] The coaxial twin-rotor configuration is widely regarded as one of the key technological paths to break through the speed bottleneck of traditional helicopters. In this configuration, the coaxial twin rotors are mainly responsible for providing lift and hovering low-speed heading control, while the tail thrust rotor is responsible for generating the main forward thrust, and the tail aerodynamic surfaces (horizontal tail and vertical tail) are used for longitudinal trim and forward heading control, respectively.

[0003] However, this configuration also introduces highly complex dynamic coupling problems. First, the shock waves generated by the fuselage during high-speed forward flight significantly alter the inflow environment of the rotor and thruster; simultaneously, the rotor's downwash and the thruster's slipstream strongly impact the horizontal and vertical stabilizers, leading to nonlinear changes in the tail's aerodynamic efficiency. Second, aerodynamic loads easily excite elastic modes of the fuselage (such as bending and torsion), and the elastic deformation of the fuselage, in turn, changes the boundary conditions of the rotor installation, thus affecting the rotor's excitation response to the fuselage. Furthermore, the high power output of the thruster causes violent fluctuations in the fuselage pitch moment, and the thruster's own vibration is also transmitted to the main fuselage through the tail boom.

[0004] Most existing modeling techniques employ a "rigid body assumption" or "one-way coupling" strategy, either ignoring fuselage elastic deformation or simplifying aerodynamic disturbances to fixed factors. While these methods are acceptable at low speeds, they struggle to accurately predict fuselage vibration response and control system redundancy trim characteristics during high-speed maneuvering flight, easily leading to insufficient design margins or control divergence. Therefore, it is necessary to develop a coupled dynamics modeling technique that comprehensively considers fuselage structural elasticity, multi-component aerodynamic disturbances, and the effects of the propulsion system. Summary of the Invention

[0005] The purpose of this invention is to provide a whole-aircraft coupled dynamics modeling method and device suitable for coaxial twin-rotor helicopters. It uses modal synthesis to handle the elasticity of the airframe, employs a high-confidence aerodynamic model to describe the interference between components, and establishes a unified coupling equation, thereby achieving an accurate description of the dynamic characteristics of the whole aircraft.

[0006] The first aspect of this invention provides a whole-aircraft coupled dynamics modeling method suitable for coaxial twin-rotor helicopters, comprising: S1. A finite element model of the entire coaxial twin-rotor helicopter, including the main fuselage structure, tail boom, horizontal tail, and vertical tail, is established using the finite element method. Based on the finite element model, low-frequency elastic modes are extracted through modal analysis, and then the dynamic equations of the elastic body are constructed by reducing the physical degrees of freedom using the modal synthesis method. The boundary conditions of the finite element model include: applying 6 spring elements at the rotor hub and applying 3 spring elements at the thrust rotor support. S2. Construct a coaxial twin-rotor unsteady aerodynamic model, a thrust propeller aerodynamic model, a horizontal and vertical tail interference aerodynamic model, and a fuselage aerodynamic model. S3. Establish the entire machine coupled dynamics equation set; S4. Numerical solution of the whole machine coupled dynamic equations is carried out. The vibration response characteristics of the whole machine are extracted from the solution results, and the engineering calculation of the complete coupled dynamic model is completed.

[0007] Optionally, the dynamic equations of the elastic organism can be constructed, including: First, low-frequency elastic modes of the fuselage are extracted from the finite element model of the whole machine through modal analysis. Then, the extracted low-frequency elastic modes are used as input, and the degrees of freedom in the physical coordinate system are reduced to the superposition of rigid body modal coordinates and elastic modal coordinates by modal synthesis method, so as to construct the dynamic equation of the elastic body.

[0008] Optionally, the order of low-frequency elastic modes extracted is 10 to 20, including the first-order vertical / lateral bending of the fuselage, the first-order torsion, the first-order bending of the vertical tail, and the symmetric / antisymmetric modes of the horizontal tail.

[0009] Optionally, in the modal synthesis method, the interface retention points include the rotor hub center point and the thrust propeller support center point; the constraint mode matrix is ​​obtained through static analysis, and the truncated regular modes are retained to within 30Hz.

[0010] Optionally, the coaxial dual-rotor unsteady aerodynamic model is based on the finite state inflow theory and introduces the superposition method of pressure potential of the upper and lower rotors, taking into account the wake tilt and upper and lower rotor interference. The induced velocity field of the coaxial dual-rotor system is formed by the superposition of contributions from the upper and lower rotors; The rotor aerodynamic load of the coaxial dual rotor system is obtained by integrating the blade element theory along the radial and azimuth angles.

[0011] Optionally, the thrust and power of the thruster in the aerodynamic model can be obtained by integrating along the blade using blade element momentum theory.

[0012] Optionally, the horizontal and vertical tail aerodynamic model is used to solve for the real-time aerodynamic loads of the horizontal and vertical tail by substituting the modified local angle of attack of the horizontal and vertical tail into the aerodynamic coefficient database. Among them, the local angle of attack of the horizontal and vertical stabilizers is corrected based on the vertical induced velocity of the horizontal stabilizer and the lateral induced velocity of the vertical stabilizer.

[0013] Optionally, establish a set of coupled dynamic equations for the entire machine, including: Establish the interface force transmission matrix between the rotor and the elastic body, and between the thruster and the elastic body; The periodic aerodynamic loads of the rotor and thruster are mapped to the generalized excitation force of the airframe modal equations by the interface force transfer matrix. The generalized excitation force induces elastic deformation of the airframe. Based on the interface force transfer matrix, the elastic displacement and velocity of the coupled interface are restored from the modal coordinates. The elastic displacement and velocity are fed back to the rotor and thruster aerodynamic equations as time-varying motion boundary conditions to complete the two-way coupling iteration and obtain the iterated interface force transfer matrix. Based on the iterated interface force transmission matrix, assemble the whole machine coupled dynamic equations.

[0014] Optionally, assemble the entire machine's coupled dynamic equations, including: Based on the Lagrange equations, the kinetic energy, potential energy, rotor / thrust propeller kinetic energy, and aerodynamic virtual work of the elastic body are assembled in a unified modal space. Based on the iterated interface force transfer matrix, a set of nonlinear coupled differential equations including rigid body motion, elastic vibration of the body, rotor flapping / swaying, and thrust propeller rotation is derived.

[0015] Optionally, the main fuselage structure adopts shell units, the tail boom adopts beam units, and the horizontal and vertical tail adopts shell units.

[0016] A second aspect of the present invention provides a whole-aircraft coupled dynamics modeling device suitable for coaxial twin-rotor helicopters, comprising: The full-aircraft finite element model creation module is used to create a full-aircraft finite element model of a coaxial twin-rotor helicopter, including the main fuselage structure, tail boom, horizontal stabilizer, and vertical stabilizer, using the finite element method. Based on the full-aircraft finite element model, low-frequency elastic modes are extracted through modal analysis, and then the dynamic equations of the elastic body are constructed by reducing the physical degrees of freedom using the modal synthesis method. The boundary conditions of the full-aircraft finite element model include: applying 6 spring elements at the rotor hub and applying 3 spring elements at the thrust rotor support. The aerodynamic model building module is used to construct unsteady aerodynamic models of coaxial twin rotors, thrust propeller aerodynamic models, horizontal and vertical tail interference aerodynamic models, and fuselage aerodynamic models. The equation system establishment module is used to establish the entire machine coupled dynamic equation system; The solver module is used to numerically solve the entire coupled dynamic equations of the aircraft, extract the vibration response characteristics of the entire aircraft body from the solution results, and complete the engineering calculation of the complete coupled dynamic model.

[0017] Optionally, the whole-machine finite element model establishment module is specifically used to: firstly extract the low-frequency elastic modes of the fuselage from the whole-machine finite element model through modal analysis; then, using the extracted low-frequency elastic modes as input, and using the modal synthesis method, reduce the degrees of freedom in the physical coordinate system to the superposition of rigid body modal coordinates and elastic modal coordinates, thereby constructing the dynamic equations of the elastic body.

[0018] Optionally, the order of low-frequency elastic modes extracted is 10 to 20, including the first-order vertical / lateral bending of the fuselage, the first-order torsion, the first-order bending of the vertical tail, and the symmetric / antisymmetric modes of the horizontal tail.

[0019] Optionally, in the modal synthesis method, the interface retention points include the rotor hub center point and the thrust propeller support center point; the constraint mode matrix is ​​obtained through static analysis, and the truncated regular modes are retained to within 30Hz.

[0020] The purpose of this invention is to provide a method and apparatus for whole-aircraft coupled dynamics modeling suitable for coaxial twin-rotor helicopters. The beneficial effects of this invention are reflected in the following aspects: 1. Full element coupling: For the first time, the coaxial twin rotor, thrust propeller, fuselage structure and horizontal and vertical tail aerodynamics are fully coupled under the same model framework, which can accurately capture the complex interaction of "thrust-aerodynamics-structure" in high-speed flight.

[0021] 2. High computational efficiency: The modal synthesis method significantly reduces the degrees of freedom of the machine structure, keeping the scale of the whole machine coupling equations within an acceptable range for engineering, and is suitable for long-term transient response analysis.

[0022] 3. Strong engineering applicability: The model can accurately predict the impact of thrust propeller slipstream on horizontal stabilizer efficiency and the feedback of fuselage elastic deformation on rotor, providing a reliable basis for predicting the vibration response of high-speed helicopters. Attached Figure Description

[0023] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. The drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0024] Figure 1 Flowchart for modeling the coupled dynamics of coaxial dual-rotor helicopter fuselage, rotor, thrust, and propeller; Figure 2 Schematic diagram of trim for a coaxial twin-rotor helicopter; Figure 3 Typical vibration response measurement point layout for the fuselage. Detailed Implementation

[0025] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0026] The features and illustrative embodiments of various aspects of the present invention will now be described in detail. Numerous specific details are set forth in the following detailed description to provide a thorough understanding of the invention. However, it will be apparent to those skilled in the art that the invention may be practiced without requiring some of these specific details. The following description of embodiments is merely intended to provide a better understanding of the invention by illustrating examples of the invention. The invention is by no means limited to any specific setups and methods set forth below, but covers any improvements, substitutions, and modifications to structures, methods, and devices without departing from the spirit of the invention. Well-known structures and techniques are not shown in the drawings and the following description to avoid unnecessarily obscuring the invention.

[0027] In the description of this invention, it should be noted that the directions or positional relationships indicated by terms such as "center," "upper," "lower," "left," "right," "vertical," "horizontal," "inner," and "outer" are based on the directions or positional relationships shown in the accompanying drawings and are only for the convenience of describing and simplifying the invention, and should not be construed as limiting the invention. Furthermore, the use of ordinal numbers (e.g., "first and second," etc.) is for distinguishing objects and is not limited to this order, and should not be construed as indicating or implying relative importance.

[0028] In the description of this invention, it should be noted that, unless otherwise explicitly specified and limited, the terms "installation," "connection," and "linking" should be interpreted broadly, encompassing both direct connection and indirect connection via an intermediate medium. Those skilled in the art can understand the specific meaning of these terms in this invention based on the specific circumstances.

[0029] It should be noted that, unless otherwise specified, the embodiments of the present invention and the features thereof can be combined with each other, and the various embodiments can be referenced and cited in each other. The present invention will now be described in detail with reference to the accompanying drawings and embodiments.

[0030] The present invention will be further described in detail below with reference to the embodiments and accompanying drawings, but the embodiments of the present invention are not limited thereto.

[0031] like Figure 1-3As shown, this invention provides a method for full-aircraft coupled dynamics modeling suitable for coaxial twin-rotor helicopters, comprising the following steps: Step 1: Constructing a dynamic model of an elastic organism based on the modal synthesis method A full-scale aircraft structural model, including the main fuselage structure, tail boom, horizontal stabilizer, and vertical stabilizer, was established using the finite element method. Low-frequency elastic modes of the fuselage (including first / second-order bending, torsional, and tail vibration modes) were extracted through modal analysis. Using modal synthesis, the numerous degrees of freedom in the physical coordinate system were reduced to a superposition of rigid body modal coordinates and elastic modal coordinates, thereby constructing the dynamic equations of the elastic body.

[0032]

[0033] Where, q f M is a generalized displacement vector containing both rigid body motion coordinates and elastic modal coordinates. f C f K f These are the reduced mass matrix, damping matrix, and stiffness matrix, respectively, Q. f This is the generalized force vector acting on the organism; In this context, a dot at the top of a variable indicates the first derivative of the variable with respect to time, while two dots indicate the second derivative of the variable with respect to time.

[0034] Step 2: Construct a distributed aerodynamic model that considers component interference. Rotor aerodynamics: Based on the finite-state inflow theory and introducing the superposition method of pressure potentials from the upper and lower rotors, an unsteady aerodynamic model of a coaxial dual-rotor system considering wake tilt and interference from the upper and lower rotors is established. The induced velocity field of the coaxial dual-rotor system is assumed to be the superposition of contributions from the upper and lower rotors, and its spatial state can be expressed as:

[0035] Where λ is the generalized coordinate vector of the induced velocity at the upper and lower rotor disks. M coax To add a mass matrix, L coax Let τ be the influence coefficient matrix, and τ be the generalized load vector of the rotor. M coax and L coax Each matrix contains non-focus block terms, which describe the mutual interference effects of the upper rotor on the lower rotor and the lower rotor on the upper rotor, respectively.

[0036] Thrust propeller and horizontal / vertical tail aerodynamics: A high-precision blade element momentum model of the thrust propeller is established. Based on momentum theory, a contracting slipstream forms behind the thrust propeller during operation. The induced velocity in the slipstream core region decays slowly along the flow direction, resulting in a significant increase in downwash / sidewash induced velocity for the downstream horizontal and vertical tail. Let the induced velocity at the thrust propeller disk be... v i The induced velocity of the slipstream reaching the horizontal tail can be approximated as: v i · k s (r), where k s (r) represents the slipstream contraction factor and attenuation function, which depend on the axial distance and radial offset of the vertical tail relative to the thruster. The thruster's thrust and power are obtained by integrating along the blades using blade element momentum theory:

[0037]

[0038] In the formula, ρ air density, V The combined velocity of the blades, c Let C1 and C be the chord lengths. d The lift-drag coefficient of the horizontal vertical tail. denoted as the inflow angle, r is the radial coordinate of a blade element on a single thrust blade from the center of the hub; R is the thrust blade radius; r / R is the dimensionless radial position, used to discretize blade elements at different radial positions of the blade, distinguishing the sections at the root, mid-section, and tip, and is the standard dimensionless radial parameter for the theoretical integral calculation of blade element momentum.

[0039] The local angle of attack corrections for the horizontal and vertical stabilizers are as follows: , ; in, w prop,ht and v prop,vt These represent the vertical induced velocity generated by the propeller slipstream at the horizontal stabilizer and the lateral induced velocity generated at the vertical stabilizer, respectively. V ∞ This refers to the forward flight speed of the helicopter.

[0040] Airframe aerodynamics: Based on wind tunnel test data or engineering estimation methods, an aerodynamic database of the fuselage, horizontal stabilizer, and vertical stabilizer is established. This database is a function of the fuselage angle of attack, sideslip angle, and elastic deformation.

[0041] Vertical induced velocity and lateral induced velocity are additional flow field parameters generated by the thruster slipstream in the horizontal and vertical tail regions, and are the calculation inputs for local angle of attack correction. Together, they constitute the core correction logic of the horizontal and vertical tail disturbance aerodynamic model. The complete logic chain is as follows: thruster slipstream → calculate horizontal tail vertical induced velocity and vertical tail lateral induced velocity → correct the local angle of attack of the horizontal and vertical tail based on the induced velocity → substitute the corrected angle of attack into the aerodynamic coefficient database to solve for the real-time aerodynamic loads of the horizontal and vertical tail.

[0042] Step 3: Establish the force transmission mechanism at the rotor-thrust propeller-airframe coupling interface The rotor hub and thruster support are defined as the coupling interface. The aerodynamic forces (including periodic alternating loads) and torques generated by the rotor and thruster are mapped to the modal coordinate equations of the elastic body through the interface force transfer matrix, serving as external excitations. Simultaneously, the elastic displacement and velocity of the body at the coupling interface, obtained from the modal coordinate solution, are used as time-varying boundary conditions and fed back to the motion equations of the rotor and thruster to correct their inflow field and motion attitude.

[0043] Among them, "load mapping and feedback" are two types of mathematical operations of the matrix: the forward operation is to obtain the generalized modal force of the airframe by multiplying the aerodynamic force by the transfer matrix; the reverse operation is to restore the physical displacement of the coupling interface by multiplying the modal coordinates by the transfer matrix, and then feed back the boundary conditions to the rotor and thrust propeller. Both steps of the operation are implemented by the matrix.

[0044] The specific mathematical description of this interface force transmission mechanism is as follows: Let the physical displacement of the coupling interface node be... The displacement reconstructed from the modal coordinates is u b =Φ b q f ,in Φ b This represents the modal matrix of the interface nodes. It also represents the aerodynamic forces acting on the interface. F b The corresponding modal generalized force is:

[0045] in, F b This includes the forces and moments at the rotor hub, as well as the tension and moments at the thrust rotor support. Simultaneously, the elastic displacement at the interface... u b and speed This will be fed back as boundary conditions to the dynamic models of the rotor and thruster.

[0046] Elastic deformation generation mechanism: The periodically alternating aerodynamic loads of the rotor and thrust propeller are mapped into generalized modal forces of the airframe through the interface force transfer matrix. F genGeneralized force input to the dynamic equations of elastic bodies Solving this equation yields the modal coordinate displacements. q Then through the interface modality matrix Φ int To restore the physical displacement of the coupling interface, i.e., the elastic deformation of the body; Feedback transmission path: The physical displacement and velocity of the interface are used as time-varying boundary conditions and directly substituted into the unsteady inflow equations of the coaxial rotor and thrust propeller to correct the local inflow velocity of the propeller disk and the motion attitude of the propeller blade. The formula reflects feedback: both the rotor induced velocity state equation and the thrust blade element integral equation introduce interface displacement correction terms. During the iterative solution of the coupled equations, the boundary displacement is updated at each time step, realizing a deformation feedback closed loop.

[0047] Step 4: Derive and solve the coupled dynamic equations of the whole machine. Based on the Lagrange equations, the kinetic energy, potential energy, rotor / thrust propeller kinetic energy, and aerodynamic virtual work of the elastic body are assembled in a unified modal space, and a set of nonlinear coupled differential equations including rigid body motion, elastic vibration of the body, rotor flapping / swaying, and thrust propeller rotation is derived.

[0048] The equations of motion for the system are established using the Lagrange equations. Let the generalized coordinates of the system be... ,in q prop and q r These represent the generalized coordinates of the rotor and thruster (including blade flapping, flaring, and the rotational degrees of freedom of the thruster). The Lagrangian function of the system is: L=TU ,in, T The total kinetic energy of the system. U Let be the total potential energy of the system. Based on the Lagrange equations, the coupled dynamic equations of the entire machine are derived as follows:

[0049] In the formula, M , C , K These are the system's mass matrix, damping matrix, and stiffness matrix, respectively, all of which contain coupling terms between the rigid body and the elastic body; F aero It is a generalized force vector that includes the aerodynamic forces of the rotor, thruster, fuselage, and horizontal and vertical tail, and is related to the system's state variables (position, velocity) and control inputs (collective pitch, cyclic pitch, etc.). F control A generalized force is input to the control system. The coupling equations are solved in the time domain using a numerical integration method (Runge-Kutta method).

[0050] For example, Figure 2 This is a schematic diagram of the overall machine's load balancing. Figure 2 All forces F and torques M are components of the generalized force vector of the coupled equation system; 1. Top rotor load: F x,2 ,F y,2 ,F z,2 M x,2 M y,2 M z,2 ; 2. Lower rotor load: F x,1 ,F y,1 ,F z,1 M x,1 M y,1 M z,1 ; 3. Fuselage aerodynamic load: F x,F ,F y,F ,F z,F M x,F M y,F M z,F ; 4. Thrust propeller and horizontal / vertical tail loads: F x,3 ,F y,3 ,F z,3 M x,3 M y,3 .

[0051] In a specific embodiment, the present invention provides a method for modeling the coupled dynamics of a coaxial dual-rotor helicopter fuselage / rotor / thrust rotor, comprising the following steps: Step 1: Modeling the elastic body The low-frequency elastic modes output from modal analysis serve as the basic input for modal synthesis, reducing the overall physical degrees of freedom of the helicopter. A full-scale finite element model of the helicopter is then created using finite element software (such as Patran / Nastran / Abaqus). During mesh generation, stress concentration areas such as the connection between the tail boom and fuselage, and the roots of the horizontal and vertical stabilizers are locally refined. The modal extraction order can be determined according to the analysis requirements, typically retaining the first 10-30 modes, mainly including the first-order vertical / lateral bending and first-order torsion of the fuselage, the first-order bending of the vertical stabilizer, and the symmetric / anti-symmetric modes of the horizontal stabilizer. Through modal synthesis, the physical model containing tens of thousands of degrees of freedom is reduced to a reduced-order model containing only rigid body degrees of freedom and a small number of elastic degrees of freedom.

[0052] Step 2: Aerodynamic Model Construction Rotor: For coaxial rigid rotors, a finite-state inflow model that takes into account the lift bias effect is used to calculate the aerodynamic interference when the upper and lower rotors are flying forward at high speed.

[0053] Thrust propeller and vertical tail: When the thrust propeller is working, it generates a contracting slipflow. The slipflow generates vertical induced velocity in the horizontal tail region and lateral induced velocity in the vertical tail region. Based on the two types of induced velocities, the local airflow angle of the horizontal and vertical tail is locally corrected. The corrected local angle of attack is then substituted into the wind tunnel aerodynamic database to complete the load calculation of the horizontal and vertical tail interference aerodynamic model.

[0054] Fuselage: The fuselage is considered as a slender beam element, and its aerodynamic coefficient is obtained by interpolation through wind tunnel test database, with an additional angle of attack correction term caused by the elastic deformation of the fuselage superimposed.

[0055] Step 3: Assembly and Solving of Coupled Equations The forces and moments at the rotor hub, and the thrust and moment of the thrust rotor, are mapped to the modal equations of the airframe using shape functions. For example, the periodic flapping moment of the rotor directly induces the bending and torsional modes of the fuselage. The equations of motion for the entire airframe system are constructed as follows:

[0056] Where M, C, and K are the system's mass matrix, damping matrix, and stiffness matrix, respectively. q Generalized displacement components, including rigid body coordinates, elastic modal coordinates, and generalized rotor / thrust rotor coordinates; F aero It is a generalized force vector that includes the aerodynamic forces of the rotor, thruster, fuselage, and horizontal and vertical tail.

[0057] The modeling method of this invention will be verified below using a coaxial twin-rotor helicopter as an example. This model adopts a coaxial rigid twin-rotor layout with a thrust propeller, and the maximum cruise speed forward ratio is about 0.35.

[0058] Step 1: Establishment and Validation of the Reduced-Order Model of the Elastic Mechanism A finite element model of the entire prototype structure was built using Patran / Nastran. Shell elements were used for the main fuselage structure, beam elements for the tail boom, and shell elements for the horizontal and vertical stabilizers. Boundary conditions: six spring elements were applied at the rotor hub, and three spring elements were applied at the thrust rotor support.

[0059] The first 30 modes were extracted using the Nastran modal analysis module. Based on energy distribution, the top 12 modes that significantly contribute to the dynamic response were selected, including 6 rigid body modes and 6 elastic body modes. The physical coordinates were reduced to modal coordinates using modal synthesis, with 12 inherent modal degrees of freedom (6 rigid body + 6 elastic body). After superimposing the coupling interface coordinates, the total generalized degrees of freedom of the system is 18. Interface points retained include the rotor hub center point and the thrust rotor support center point. The constraint mode matrix was obtained through static analysis, and truncated canonical modes were retained up to 30 Hz.

[0060] Step 2: Distributed Aerodynamic Model 1. Unsteady aerodynamic model of a coaxial twin rotor A coaxial dual-rotor pressure potential superposition model based on an extension of the finite-state inflow theory is adopted. The inflow harmonic order is set to N=4, and the number of induced velocity state variables for each rotor is 11, for a total of 22 state variables for both upper and lower rotors. The state equations are as follows: ; The rotor aerodynamic loads are obtained by integrating the blade element theory along the radial direction (20 blade element units) and the azimuth angle (one sampling point every 10°). The airfoil aerodynamic data are obtained using the wind tunnel test data table for the prototype (angle of attack range -12°~20°, Mach number 0.2~0.65).

[0061] 2. Aerodynamic model of interference between thrust propeller and horizontal / vertical tail Thrust propeller blade element model: Divided radially into 15 blade element elements, each using typical airfoil aerodynamic data. The geometric position parameters of the horizontal stabilizer and vertical stabilizer relative to the thrust propeller are as follows: the horizontal stabilizer is located 0.8m in front of the thrust propeller, and the vertical stabilizer is located 1.2m behind the thrust propeller. The slipstream influence at the horizontal stabilizer is relatively small, mainly affected by the rotor downwash; the vertical stabilizer is entirely within the slipstream core region.

[0062] 3. Fuselage and Vertical Aerodynamic Database Express the aerodynamic coefficients of the fuselage, horizontal stabilizer, and vertical stabilizer as angles of attack. (-10°~20°, step size 2°), sideslip angle (-20°~20°, step size 2°) and elastic deformation The functions are: horizontal stabilizer twist angle -2° to 2°, vertical stabilizer bending displacement -0.05m to 0.05m. The database contains 2000 data points obtained from full-aircraft wind tunnel tests. For flight conditions not observed at the test points, linear interpolation was used to obtain the aerodynamic coefficients.

[0063] Step 3: Assembly and solution of the whole machine coupled equations The generalized coordinate vector of the system is defined as:

[0064] in, q r It includes flapping angle, oscillation angle and pitch angle of 4 blades on each of the upper and lower rotors, totaling 24 degrees of freedom; q prop It includes 7 degrees of freedom, including the thruster's rotational speed and blade pitch angle. The total number of degrees of freedom for the entire system is 49.

[0065] Using the Lagrange equations to establish the equations of motion, the final coupled equations are in the form of:

[0066] The mass matrix includes the airframe-rotor coupling term, the damping matrix includes structural damping and aerodynamic damping, and the stiffness matrix includes structural stiffness, rotor centrifugal stiffness, and aeroelastic stiffness.

[0067] The fourth-order Rung-Kutta method was used for the solution, with an integration time step of 1 / 1024 seconds. The initial conditions were a hovering trimmed state (collective pitch 8.5°, periodic pitch 0°). The total simulation time was 15 seconds. The simulated prototype started from hovering and uniformly increased the collective pitch and pushed the periodic pitch lever forward within 5-10 seconds, accelerating the forward speed from 0 to 70 m / s.

[0068] Step 4: Simulation Results and Comparative Analysis Simulations were run under the above parameter settings to extract key dynamic responses, which were then compared with flight test data. Simultaneously, a traditional rigid body model (ignoring airframe elasticity, retaining only 6 rigid body degrees of freedom, and outputting rigid body acceleration) was established as a reference for the dynamic responses.

[0069] 1. Vibration response of the machine body At 60 m / s², the peak vertical NΩ vibration response of the prototype at the cockpit floor was 0.23 (dimensionless), while the flight test value was 0.21 (dimensionless), with an error of 9.5%. For comparison, traditional rigid body models cannot calculate elastic vibration acceleration; they can only output the rigid body vertical motion acceleration directly caused by rotor thrust and fuselage aerodynamic changes. The results show that the response at the same location is only 0.09 (dimensionless), far smaller than the actual flight test value. In summary, this indicates that neglecting airframe elasticity and relying solely on rigid body motion acceleration will severely underestimate the actual vibration level. The layout of key vibration measurement points for the entire aircraft is shown below. Figure 3 The location on the cockpit floor is Figure 3 Points 5 (to the left along the fuselage heading) and 6 (to the left along the fuselage heading).

[0070] 2. Tail boom elastic torsion and directional stability Simulation results show that when the thrust propeller power exceeds 800 kW, the elastic torsion angle of the tail boom reaches 2.1°, leading to a decrease in the actual sideslip angle of the vertical tail. The directional static stability derivative predicted by this model decreases from 0.0012 (dimensionless) in the rigid body model to 0.001 (dimensionless), which is in good agreement with the test flight result of 0.00105 (dimensionless). This result indicates that neglecting the elasticity of the tail boom would overestimate directional stability by approximately 15% or more.

[0071] 3. Computational efficiency This model takes approximately 23 minutes to complete a 15-second time-domain simulation. If a fully finite element model (without order reduction) were used, it would likely require tens of hours. The reduction in degrees of freedom brought about by the modal synthesis method improves computational efficiency by two orders of magnitude, meeting the needs of engineering iterations.

[0072] The above detailed embodiments are a description of the present invention. It should not be considered that the specific embodiments of the present invention are limited to these descriptions. For those skilled in the art, several simple deductions and substitutions can be made without departing from the concept of the present invention, and all of these should be considered to fall within the protection scope of the present invention.

Claims

1. A method for full-aircraft coupled dynamics modeling applicable to coaxial twin-rotor helicopters, characterized in that, include: S1. A finite element model of the entire coaxial twin-rotor helicopter, including the main fuselage structure, tail boom, horizontal tail, and vertical tail, is established using the finite element method. Based on the finite element model, low-frequency elastic modes are extracted through modal analysis, and then the dynamic equations of the elastic body are constructed by reducing the physical degrees of freedom using the modal synthesis method. The boundary conditions for the full-scale finite element model include: applying 6 spring elements at the rotor hub and 3 spring elements at the thrust propeller support. S2. Construct a coaxial twin-rotor unsteady aerodynamic model, a thrust propeller aerodynamic model, a horizontal and vertical tail interference aerodynamic model, and a fuselage aerodynamic model. S3. Establish the entire machine coupled dynamics equation set; S4. Numerical solution of the whole machine coupled dynamic equations, extract the whole machine vibration response characteristics through the solution results, and complete the engineering calculation of the complete coupled dynamic model.

2. The whole-aircraft coupled dynamics modeling method for coaxial twin-rotor helicopters according to claim 1, characterized in that, The dynamic equations of an elastic organism are constructed, including: First, low-frequency elastic modes of the fuselage are extracted from the finite element model of the whole machine through modal analysis. Then, the extracted low-frequency elastic modes are used as input, and the degrees of freedom in the physical coordinate system are reduced to the superposition of rigid body modal coordinates and elastic modal coordinates by modal synthesis method, so as to construct the dynamic equation of the elastic body.

3. The whole-aircraft coupled dynamics modeling method for coaxial twin-rotor helicopters according to claim 2, characterized in that, The low-frequency elastic modes are extracted in order of 10 to 20, including the first-order vertical / lateral bending of the fuselage, the first-order torsion, the first-order bending of the vertical tail, and the symmetric / antisymmetric modes of the horizontal tail.

4. The whole-aircraft coupled dynamics modeling method for coaxial twin-rotor helicopters according to claim 2, characterized in that, In the modal synthesis method, the interface retention points include the rotor hub center point and the thrust propeller support center point; the constraint mode matrix is ​​obtained through static analysis, and the truncated canonical modes are retained to within 30Hz.

5. The whole-aircraft coupled dynamics modeling method for coaxial twin-rotor helicopters according to claim 1, characterized in that, The unsteady aerodynamic model of the coaxial dual rotor is based on the finite state inflow theory and introduces the superposition method of pressure potential of the upper and lower rotors, taking into account the wake tilt and the disturbance of the upper and lower rotors. The induced velocity field of the coaxial dual-rotor system is formed by the superposition of contributions from the upper and lower rotors; The rotor aerodynamic load of the coaxial dual rotor system is obtained by integrating the blade element theory along the radial and azimuth angles.

6. The whole-aircraft coupled dynamics modeling method for coaxial twin-rotor helicopters according to claim 1, characterized in that, In the aerodynamic model of the thruster, the thrust and power of the thruster are obtained by integrating along the blade using the blade element momentum theory.

7. The whole-aircraft coupled dynamics modeling method for coaxial twin-rotor helicopters according to claim 1, characterized in that, The horizontal and vertical tail aerodynamic model is used to solve for the real-time aerodynamic loads of the horizontal and vertical tail by substituting the modified local angle of attack of the horizontal and vertical tail into the aerodynamic coefficient database. Among them, the local angle of attack of the horizontal and vertical stabilizers is corrected based on the vertical induced velocity of the horizontal stabilizer and the lateral induced velocity of the vertical stabilizer.

8. The whole-aircraft coupled dynamics modeling method for coaxial twin-rotor helicopters according to claim 1, characterized in that, Establish the entire coupled dynamic equation set, including: Establish the interface force transmission matrix between the rotor and the elastic body, and between the thruster and the elastic body; The periodic aerodynamic loads of the rotor and thruster are mapped to the generalized excitation force of the airframe modal equations by the interface force transfer matrix. The generalized excitation force induces elastic deformation of the airframe. Based on the interface force transfer matrix, the elastic displacement and velocity of the coupled interface are restored from the modal coordinates. The elastic displacement and velocity are fed back to the rotor and thruster aerodynamic equations as time-varying motion boundary conditions to complete the two-way coupling iteration and obtain the iterated interface force transfer matrix. Based on the iterated interface force transmission matrix, assemble the whole machine coupled dynamic equations.

9. The whole-aircraft coupled dynamics modeling method for coaxial twin-rotor helicopters according to claim 8, characterized in that, Assemble the coupled dynamic equations of the entire machine, including: Based on the Lagrange equations, the kinetic energy, potential energy, rotor / thrust propeller kinetic energy, and aerodynamic virtual work of the elastic body are assembled in a unified modal space. Based on the iterated interface force transfer matrix, a set of nonlinear coupled differential equations including rigid body motion, elastic vibration of the body, rotor flapping / swaying, and thrust propeller rotation is derived.

10. The whole-aircraft coupled dynamics modeling method for coaxial twin-rotor helicopters according to claim 1, characterized in that, The main fuselage structure uses shell units, the tail boom uses beam units, and the horizontal and vertical tails use shell units.

11. A whole-aircraft coupled dynamics modeling device suitable for coaxial twin-rotor helicopters, characterized in that, include: The whole-aircraft finite element model building module is used to build a whole-aircraft finite element model of a coaxial twin-rotor helicopter, including the main fuselage structure, tail boom, horizontal tail and vertical tail, using the finite element method; based on the whole-aircraft finite element model, low-frequency elastic modes are extracted through modal analysis, and then the dynamic equations of the elastic body are constructed by reducing the physical degrees of freedom using the modal synthesis method. The boundary conditions for the full-scale finite element model include: applying 6 spring elements at the rotor hub and 3 spring elements at the thrust propeller support. The aerodynamic model building module is used to construct unsteady aerodynamic models of coaxial twin rotors, thrust propeller aerodynamic models, horizontal and vertical tail interference aerodynamic models, and fuselage aerodynamic models. The equation system establishment module is used to establish the entire machine coupled dynamic equation system; The solver module is used to numerically solve the entire coupled dynamic equations of the aircraft, extract the vibration response characteristics of the entire aircraft body from the solution results, and complete the engineering calculation of the complete coupled dynamic model.

12. The whole-aircraft coupled dynamics modeling device for coaxial twin-rotor helicopters according to claim 11, characterized in that, The module for establishing the full-machine finite element model is specifically used to: firstly, extract the low-frequency elastic modes of the fuselage from the full-machine finite element model through modal analysis; then, using the extracted low-frequency elastic modes as input, reduce the degrees of freedom in the physical coordinate system to the superposition of rigid body modal coordinates and elastic modal coordinates using the modal synthesis method, thereby constructing the dynamic equations of the elastic body.

13. The whole-aircraft coupled dynamics modeling device for coaxial twin-rotor helicopters according to claim 12, characterized in that, The low-frequency elastic modes are extracted in order of 10 to 20, including the first-order vertical / lateral bending of the fuselage, the first-order torsion, the first-order bending of the vertical tail, and the symmetric / antisymmetric modes of the horizontal tail.

14. The whole-aircraft coupled dynamics modeling device for coaxial twin-rotor helicopters according to claim 12, characterized in that, In the modal synthesis method, the interface retention points include the rotor hub center point and the thrust propeller support center point; the constraint mode matrix is ​​obtained through static analysis, and the truncated canonical modes are retained to within 30Hz.

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