Dynamic modeling method for aircraft containing multi-degree-of-freedom executing mechanism

Abstract

The invention provides an aircraft dynamics modeling method comprising a multi-degree-of-freedom execution mechanism, the modeling method regresses advanced theories and ideas of general mechanics and mechanics basic disciplines, an aircraft is regarded as a multi-rigid-body system, and a recursive modeling method is based on a virtual power principle and a multi-body system topological configuration. A deformable multi-body system dynamics model suitable for an aircraft containing multiple execution mechanisms is established, and multi-degree-of-freedom and strong-coupling motion characteristics between external aircraft body attitude motion and internal component relative motion of the aircraft are accurately described.

Description

Technical Field

[0001] The invention belongs to the technical field of high-speed flight, and in particular relates to a method for dynamic modeling of an aircraft containing a multi-degree-of-freedom actuator. Background Art

[0002] Aircraft with multi-degree-of-freedom actuators are a type of aircraft whose internal components have a large range of relative motion. This type of aircraft can achieve continuous deformation in space and time according to pre-designed programs based on different flight environments, obtain optimal aerodynamic performance, and complete various types of flight missions. It has better maneuverability and flexibility and is currently a hot topic in the field of high-speed flight research.

[0003] Compared with traditional fixed-shape aircraft, aircraft with multi-degree-of-freedom actuators face more complex dynamic problems. On the one hand, the aircraft's aerodynamic characteristic parameters, center of mass and moment of inertia characteristic parameters all show significant time-varying characteristics during deformation, and traditional assumptions such as the "coefficient freezing method" are no longer applicable; on the other hand, there are complex interaction forces and torques between the moving parts inside the aircraft, which brings multi-degree-of-freedom and strong nonlinear characteristics to the dynamic model, which has a great impact on the stability and maneuverability of the aircraft, and even causes instability.

[0004] The dynamic modeling method under the traditional flight mechanics description framework regards the aircraft as a controlled rigid body and establishes a 6-DOF dynamic equation based on the Newton-Euler method of classical theoretical mechanics. This method can no longer meet the urgent requirements for the detailed description of the relative motion of aircraft components with multi-DOF actuators. Therefore, whether from the perspective of accurate description of dynamic behavior or from the need for precise design of high-performance control algorithms, it is urgent to study the dynamic modeling method of aircraft with multi-DOF actuators. Summary of the invention

[0005] The present invention aims to solve at least one of the technical problems existing in the prior art or related art.

[0006] To this end, the present invention provides a method for modeling the dynamics of an aircraft with multi-degree-of-freedom actuators. This modeling method returns to the advanced theories and ideas of general mechanics and basic mechanics, regards the aircraft as a multi-rigid body system, and establishes a deformable multi-body system dynamics model suitable for aircraft with multi-actuators based on the principle of virtual power and the recursive modeling method of multi-body system topology configuration, accurately describing the multi-degree-of-freedom and strong coupling motion characteristics between the external body attitude motion of the aircraft and the relative motion of internal components.

[0007] The technical solution of the present invention is as follows:

[0008] According to one aspect, a method for modeling aircraft dynamics including a multi-degree-of-freedom actuator is provided, the modeling method comprising:

[0009] Step 1, obtaining the relative motion form between the rigid bodies of the aircraft containing multi-degree-of-freedom actuators, and kinematically describing the rigid bodies of the aircraft and the position of the pressure center according to the vector transfer relationship and the principle of angular velocity superposition, wherein the aircraft includes two types of actuators: one is a sliding actuator, which controls component A to move axially with a displacement s; the other is a rotating actuator, which controls component B1 and component B2 to rotate along the fuselage around the rudder axis to change the aerodynamic layout, and the rigid bodies of the aircraft include the fuselage, component B1, component B2 and component A;

[0010] Step 2: Establish a conversion relationship matrix between kinematic parameters involved in the virtual power principle and system state variables based on the kinematic description obtained in step 1;

[0011] Step 3: Based on the virtual power principle and the conversion relationship matrix obtained in step 2, the virtual power equations of each rigid body with the virtual velocity of the system state variable as the virtual variation are obtained, and the virtual power equations of each rigid body are superimposed to obtain the virtual power equation of the system;

[0012] Step 4: Obtain the free flight simulation dynamics equation and the virtual flight test dynamics equation of the aircraft multi-body system according to the system virtual power equation obtained in step 3 to construct a multi-body dynamics model

[0013] Step 5: Incorporate the multi-body dynamics model obtained in step 4 into the original flight mechanics description framework to complete the aircraft dynamics modeling.

[0014] Furthermore, in step 1, the fuselage dynamics is described as follows:

[0015] The conversion relationship between the fuselage coordinate system and the ground coordinate system is expressed as:

[0016]

[0017] Among them, x1, y1, z1 are the three basic vectors of the fuselage coordinate system respectively; x, y, z are the three basic vectors of the ground coordinate system respectively; the rotation of the fuselage coordinate system relative to the ground coordinate system is decomposed into three fixed-axis rotations, ① the first rotation is a rotation of the yaw angle ψ around the y axis of the ground coordinate system; ② the second rotation is a rotation of the pitch angle θ around the z′ axis corresponding to the original z axis after the first rotation; ③ the third rotation is a rotation of the roll angle γ around the x″ axis corresponding to the x axis after the second rotation; the x″ axis is the x1 axis of the fuselage coordinate system; R1 is the rotation matrix;

[0018]

[0019] According to the angular velocity superposition principle, the angular velocity ω1 is the sum of the product of the rate of change of the rotation angle and the axis of rotation, as shown in the following formula:

[0020]

[0021] Furthermore, in step 1, components B1 and B2 are dynamically described in the following manner:

[0022] 1) The conversion relationship between the component B1 / B2 coordinate system and the ground coordinate system is expressed as

[0023]

[0024] Among them, x2, y2, z2 are the three basic vectors of the component B1 / B2 coordinate system respectively; the variable sweep motion of the aircraft is realized by the rotation motion of the component B1 / B2 around the rotation axis, and the component B1 / B2 coordinate system is regarded as the body coordinate system rotated around the y axis by an angle of x, wherein the component B1 rotation angle x is positive, the component B2 rotation angle x is negative, and the two are opposite numbers; R2 is the rotation matrix;

[0025]

[0026] 2) According to the principle of angular velocity superposition, the angular velocity ω2 of component B1 / B2 is:

[0027]

[0028] 3) The components of the centroid of component B1 / B2 in the ground coordinate system are expressed as

[0029]

[0030] Its time derivative is expressed as:

[0031]

[0032] Among them, the hinge point of the rotation connection between component B1 / B2 and the fuselage is A, and the body hinge vector component fixed in the fuselage coordinate system is The hinge vector components of the body fixed in the B1 / B2 coordinate system are: The component of the center of mass of the fuselage in the ground coordinate system is r1.

[0033] Furthermore, in step 1, component A is dynamically described in the following manner:

[0034] The real-time position of the centroid of component A in the ground coordinate system is expressed as:

[0035]

[0036] Its time derivative can be:

[0037]

[0038] Among them, in the initial state, the component of the center of mass of component A in the fuselage coordinate system is Afterwards, it moves in a straight line along the axial direction at a speed v and a straight line displacement s. During the movement, the coordinate system of component A is always consistent with the coordinate system of the fuselage, and its angular velocity and angular acceleration are the same as those of the fuselage.

[0039] Furthermore, in step 1, the compression center is dynamically described by the following method:

[0040] Assume that the coordinates of the pressure center relative to the fuselage coordinate system are Its components in the ground coordinate system are

[0041]

[0042] Its time derivative is expressed as

[0043]

[0044] Furthermore, in step 2, the conversion relationship between each rigid body kinematic parameter and the system description variable is as follows:

[0045] 1) Relationship between the virtual angular velocity of the fuselage and the virtual angular velocity of the attitude

[0046]

[0047] 2) Relationship between body angular acceleration and attitude angular acceleration

[0048]

[0049] 3) Relationship between virtual angular velocity of component B1 and virtual angular velocity of attitude and virtual rotation angular velocity

[0050]

[0051] 4) Relationship between the virtual center of mass velocity of component B1 and the virtual center of mass velocity of the fuselage, attitude virtual angular velocity, and virtual rotation angular velocity

[0053]

[0054] 5) Relationship between component B1 angular acceleration and attitude angular acceleration and angular acceleration

[0055]

[0056] 6) Relationship between the mass center acceleration of component B1 and the mass center acceleration of the fuselage, attitude angular acceleration, and angular acceleration

[0057]

[0058] 7) Relationship between component B2 virtual angular velocity and attitude virtual angular velocity and virtual rotation angular velocity

[0059] 8) Relationship between the virtual center of mass velocity of component B2 and the virtual center of mass velocity of the fuselage, attitude virtual angular velocity, and virtual rotation angular velocity

[0060]

[0061] 9) Relationship between component B2 angular acceleration and attitude angular acceleration and rotational angular acceleration

[0062] 10) Relationship between the mass center acceleration of component B2 and the mass center acceleration of the fuselage, attitude angular acceleration, and angular acceleration

[0063]

[0064] 11) Relationship between virtual angular velocity of component A and virtual angular velocity of attitude

[0065]

[0066] 12) Relationship between the virtual center of mass velocity of component A and the virtual center of mass velocity of the fuselage, attitude virtual angular velocity, and virtual slip velocity

[0067]

[0068] 13) Relationship between component A angular acceleration and attitude angular acceleration

[0069]

[0070] 14) Relationship between the center-of-mass acceleration of component A and the center-of-mass acceleration of the fuselage, attitude angular acceleration, and slip acceleration

[0071]

[0072] The conversion matrix and acceleration margin matrix involved are defined as follows

[0073] Table 1 Definition of transformation matrix between rigid body kinematic parameters and system description variables

[0074] Conversion matrix between body virtual angular velocity and attitude virtual angular velocity <![CDATA[T 1ω ]]> Component B1 virtual angular velocity and virtual rotation angular velocity conversion matrix <![CDATA[T χ,R ]]> Component B2 virtual angular velocity and virtual rotation angular velocity conversion matrix <![CDATA[T χ,L ]]> Conversion matrix between virtual center of mass velocity of component B1 and virtual angular velocity and virtual rotation angular velocity <![CDATA[T 2r,R ,T 2χ,R ]]> Conversion matrix between virtual center of mass velocity of component B2 and virtual angular velocity and virtual rotation angular velocity <![CDATA[T 2r,L ,T 2χ,L ]]> Conversion matrix between virtual center of mass velocity of component A and virtual angular velocity and virtual slip velocity of attitude <![CDATA[T 3r ,T 3s ]]> The margin matrix between the body angular acceleration and the attitude angular acceleration <![CDATA[α 1ω ]]> Residual matrix for angular acceleration of component B1 <![CDATA[α 2ω,R ]]> Residual matrix for angular acceleration of component B2 <![CDATA[α 2ω,L ]]> Residual matrix for the center of mass acceleration of component B1 <![CDATA[α 2r,R ]]> Residual matrix for the center of mass acceleration of component B2 <![CDATA[α 2r,L ]]> The residual matrix of the center of mass acceleration of component A <![CDATA[α 3r ]]>

[0075] Furthermore, in step 3, the system virtual power equation is obtained by the following formula:

[0076]

[0077] Where M1, F1 are the generalized mass matrix and force matrix of the fuselage, M 2,R ,F 2,R is the generalized mass matrix and force matrix of component B1, M 2,L ,F 2,Lis the generalized mass matrix and force matrix of component B2, M3, F3 are the generalized mass matrix and force matrix of component A, and the specific representation is as follows:

[0078]

[0079]

[0080] Where m1 is the mass of the fuselage, m2 is the mass of components B1 and B2, J1 is the moment of inertia of the fuselage, and J 2,R is the moment of inertia of component B1, J 2,L is the moment of inertia of component B2, J3 is the moment of inertia of component A, F air is the aerodynamic force vector borne by the fuselage, M χ,R is the control torque of component B1, M χ,L is the control torque of component B2, F s is the control thrust of component A, g is the acceleration due to gravity, T sr is the conversion matrix between the virtual velocity of the center of pressure and the virtual angular velocity of the attitude, E 3×3 is a 3×3 identity matrix.

[0081] Furthermore, the step 4 specifically includes:

[0082] 4.1 Determine whether the aircraft is constrained. If the aircraft is not constrained, go to step 4.2; if the aircraft is constrained, go to step 4.3;

[0083] 4.2 According to the system virtual power equation, remove the virtual variation to obtain the system free flight simulation dynamic equation;

[0084] 4.3 According to the specific constraint form, constraint processing is performed, in which the state variables are divided into independent terms and dependent terms, and the conversion relationship between independent terms and dependent terms is established. The independent terms are substituted into the system virtual power equation to obtain the system virtual power equation corresponding to the independent virtual variation. The independent virtual variation is removed to obtain the system virtual flight test dynamics equation.

[0085] According to another aspect, a computer device is provided, comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor implements the above-mentioned prediction method when executing the computer program.

[0086] Compared with the prior art, the present invention has the following beneficial effects:

[0087] 1) The present invention gives full play to the advantages of combining the basic principles of dynamic modeling with the dynamic modeling method of multi-body system. On the one hand, it utilizes the universality of the virtual power principle, and on the other hand, it introduces the mathematical description of the topological configuration of the multi-body system and utilizes the vector transfer relationship between components to make the modeling more systematic.

[0088] 2) The present invention not only considers the external variable configuration of the conventional variable sweep (i.e., component B1\B2), but also considers the internal variable configuration of the internal object (i.e., component A), as well as the dynamic influence of the mutual coupling between them. The variants are more complex and the factors considered are more comprehensive.

[0089] 3) From the perspective of overall and sub-professional coordination and engineering application, the present invention reincorporates the constructed multi-body dynamics model into the traditional flight mechanics description framework, forming a deformable aircraft multi-body system dynamics model under the traditional flight mechanics description framework, which is more applicable.

[0090] 4) The present invention can take into account both the refined characterization of numerical simulation and the concise description of engineering applications, avoiding the errors caused by the decoupling of the translational and rotational coupling terms of the mass matrix of the rigid body dynamics model under the traditional flight mechanics description framework. BRIEF DESCRIPTION OF THE DRAWINGS

[0091] The included drawings are used to provide a further understanding of the embodiments of the present invention, which constitute a part of the specification, are used to illustrate the embodiments of the present invention, and together with the text description, explain the principles of the present invention. Obviously, the drawings in the following description are only some embodiments of the present invention, and for ordinary technicians in this field, other drawings can be obtained based on these drawings without creative work.

[0092] Figure 1 It is a schematic diagram of a multi-degree-of-freedom actuator aircraft.

[0093] Figure 2 It is a topological configuration diagram of a multi-body system of an aircraft with multiple degrees of freedom actuators. DETAILED DESCRIPTION

[0094] It should be noted that, in the absence of conflict, the embodiments in this application and the features in the embodiments can be combined with each other. The technical solutions in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. The following description of at least one exemplary embodiment is actually only illustrative and is by no means intended to limit the present invention and its application or use. Based on the embodiments in the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0095] It should be noted that the terms used herein are only for describing specific embodiments and are not intended to limit the exemplary embodiments according to the present application. As used herein, unless the context clearly indicates otherwise, the singular form is also intended to include the plural form. In addition, it should be understood that when the terms "comprise" and / or "include" are used in this specification, it indicates the presence of features, steps, operations, devices, components and / or combinations thereof.

[0096] Unless otherwise specifically stated, the relative arrangement of the parts and steps described in these embodiments, numerical expressions and numerical values ​​do not limit the scope of the present invention. At the same time, it should be understood that, for ease of description, the sizes of the various parts shown in the accompanying drawings are not drawn according to the actual proportional relationship. The technology, method and equipment known to ordinary technicians in the relevant field may not be discussed in detail, but in appropriate cases, the technology, method and equipment should be regarded as a part of the authorization specification. In all examples shown and discussed here, any specific value should be interpreted as being merely exemplary, rather than as a limitation. Therefore, other examples of exemplary embodiments may have different values. It should be noted that similar reference numerals and letters represent similar items in the following drawings, so once a certain item is defined in an accompanying drawing, it does not need to be further discussed in subsequent drawings.

[0097] like Figure 1 As shown, in one embodiment of the present invention, a method for modeling aircraft dynamics including a multi-degree-of-freedom actuator is provided, the modeling method comprising:

[0098] Step 1: Obtain the relative motion form between the rigid bodies of the aircraft containing multi-degree-of-freedom actuators, and perform kinematic description of the rigid bodies of the aircraft and the position of the pressure center according to the vector transfer relationship and the principle of angular velocity superposition. The aircraft includes two types of actuators: one is a sliding actuator, which controls component A to move axially with a displacement s; the other is a rotating actuator, which controls component B1 and component B2 to rotate around the rudder axis along the fuselage to change the aerodynamic layout. The rigid bodies of the aircraft include the fuselage, component B1, component B2 and component A. Their topological structures are shown in the figure below. Figure 2 As shown;

[0099] Step 2: Establish a conversion relationship matrix between kinematic parameters involved in the virtual power principle and system state variables based on the kinematic description obtained in step 1;

[0100] Step 3: Based on the virtual power principle and the conversion relationship matrix obtained in step 2, the virtual power equations of each rigid body with the virtual velocity of the system state variable as the virtual variation are obtained, and the virtual power equations of each rigid body are superimposed to obtain the virtual power equation of the system;

[0101] Step 4: According to the system virtual power equation obtained in step 3, obtain the free flight simulation dynamics equation and the virtual flight test dynamics equation of the aircraft multi-body system to construct a multi-body dynamics model.

[0102] Step 5: Incorporate the multi-body dynamics model obtained in step 4 into the original flight mechanics description framework to complete the aircraft dynamics modeling.

[0103] The embodiment of the present invention is modeled based on the virtual power principle and the multi-body system topological configuration graph theory. For a multi-rigid body system, the virtual power principle can be expressed as follows: the sum of the virtual power of the translational inertia force and the virtual power of the rotational inertia force is equal to the virtual power of the external force. The specific expression is as follows:

[0104]

[0105] In the formula, r c,i ,ω i F is the radius vector and angular velocity vector of the rigid body’s center of mass; a,i ,M a,i is the external force and torque on the rigid body; m i is the rigid body mass; J i is the moment of inertia; g is the gravitational acceleration vector; n is the number of rigid bodies.

[0106] The following describes in detail each step of the embodiment of the present invention:

[0107] The kinematic description in step 1 is as follows:

[0108] The aircraft can be regarded as a four-rigid body system consisting of the fuselage, component B1, component B2, and component A, and the kinematic description of the rigid body and the position of the center of pressure is given separately.

[0109] 1.1 Body

[0110] According to the traditional flight mechanics description framework, the rotation of the fuselage coordinate system relative to the ground coordinate system can be decomposed into three fixed-axis rotations, namely: ① The first rotation is a yaw angle ψ around the y-axis of the ground coordinate system; ② The second rotation is a pitch angle θ around the z′ axis corresponding to the original z axis after the first rotation; ③ The third rotation is a roll angle γ around the x″ axis corresponding to the x axis after the second rotation (i.e., the x1 axis of the fuselage coordinate system); The conversion relationship between the fuselage coordinate system and the ground coordinate system can be expressed as

[0111]

[0112] Among them, x1, y1, z1 are the three basis vectors of the fuselage coordinate system; the rotation matrix is

[0113]

[0114] According to the principle of angular velocity superposition, the angular velocity is the sum of the product of the rate of change of the angle of rotation and the axis of rotation, that is,

[0115]

[0116] 1.2 Part B1 or B2

[0117] The variable sweep motion of the aircraft is achieved by the rotation of component B1 / B2 around the rotation axis. The component B1 / B2 coordinate system can be regarded as the body coordinate system rotated around the y axis by an angle x. Therefore, the conversion relationship between the component B1 / B2 coordinate system and the ground coordinate system can be expressed as

[0118]

[0119] Among them, x2, y2, z2 are the three basic vectors of the component B1 / B2 coordinate system respectively; the rotation matrix is

[0120]

[0121] Similarly, according to the principle of angular velocity superposition, the angular velocity of rotation is

[0122]

[0123] Among them, the rotation angle χ of component B1 is positive, and the rotation angle χ of component B2 is negative, and the two are opposite numbers.

[0124] The hinge point of the rotation connection between component B1 / B2 and the fuselage is A, and the body hinge vector component fixed in the fuselage coordinate system is The hinge vector components fixed in the coordinate system of component B are: According to the vector transfer relationship, the component of the centroid of component B in the ground coordinate system can be expressed as

[0125]

[0126] Its time derivative can be expressed as

[0127]

[0128] 1.3 Part A

[0129] In the initial state, the component of the center of mass of component A in the fuselage coordinate system is Then it moves linearly along the axial direction at a speed v (linear displacement s); then the real-time position of the center of mass of component A in the ground coordinate system can be expressed as

[0130]

[0131] Its time derivative can be expressed as

[0132]

[0133] Among them, the component of the center of mass of the fuselage in the ground coordinate system is r1; during the movement of component A, the coordinate system is always consistent with the fuselage coordinate system, and its rotational angular velocity and angular acceleration are the same as those of the fuselage.

[0134] 1.4 Heart Pressure

[0135] During the virtual flight test, the aircraft is mainly affected by gravity and aerodynamic forces. Gravity acts directly on the center of mass, vertically downward; aerodynamic forces act on the center of pressure. The coordinates of the center of pressure relative to the fuselage coordinate system are set as Its components in the ground coordinate system are

[0136]

[0137] Its time derivative can be expressed as

[0138]

[0139] This completes the kinematic description of each rigid body.

[0140] In the embodiment of the present invention, in step 2:

[0141] According to the motion characteristics of the deformable aircraft, the body center of mass vector r1, pitch angle θ, yaw angle ψ, roll angle γ, translation displacement s of component A, and rotation angle χ of component B are selected as system description variables. For the convenience of description, the vector is defined

[0142]

[0143] From the above kinematic description, the conversion relationship between kinematic parameters and description variables can be obtained, which is summarized in the following table (the definition of the conversion matrix is ​​shown in Table 1).

[0144] Table 2 Relationship matrix between kinematic parameters and description variables of deformable aircraft

[0145]

[0146] In the embodiment of the present invention, in step three:

[0147] Substituting the kinematic relationship in the above table into the virtual power principle equation, combined with the force characteristics of the aircraft, we can obtain the virtual power equations of the fuselage, component B1, component B2, and component A respectively. By superimposing them, we can obtain the system virtual power equation, which is as follows

[0148]

[0149] Where M1, F1 are the generalized mass matrix and force matrix of the fuselage, M 2,R ,F 2,R is the generalized mass matrix and force matrix of component B1, M2,L ,F 2,L is the generalized mass matrix and force matrix of component B2, M3, F3 are the generalized mass matrix and force matrix of component A, and the specific representation is as follows:

[0150]

[0151]

[0152] Where m1 is the mass of the fuselage, m2 is the mass of components B1 and B2, J1 is the moment of inertia of the fuselage, and J 2,R is the moment of inertia of component B1, J 2,L is the moment of inertia of component B2, J3 is the moment of inertia of component A, F air is the aerodynamic force vector borne by the fuselage, M χ,R is the control torque of component B1, M χ,L is the control torque of component B2, F s is the control thrust of component A, g is the acceleration due to gravity, T sr is the conversion matrix between the virtual velocity of the center of pressure and the virtual angular velocity of the attitude, E 3×3 is a 3×3 identity matrix.

[0153] At this point, the virtual power equations of each rigid body are superimposed to obtain the system virtual power equation.

[0154] In the embodiment of the present invention, in step 4, the free flight simulation dynamics equations and the virtual flight test dynamics equations of the deformable aircraft multi-body system are respectively given.

[0155] 4.1 Free flight dynamics equations

[0156] In actual free flight, the aircraft is not subject to any constraints and has 6+2 independent degrees of freedom. The center of mass vector r1, pitch angle θ, yaw angle ψ, roll angle γ, translation displacement s of component A, and rotation angle χ of component B1 / B2 are independent of each other. The dynamic equation can be directly written as

[0157]

[0158] Where M1, F1 are the generalized mass matrix and force matrix of the fuselage, M 2,R ,F 2,R is the generalized mass matrix and force matrix of component B1, M 2,L ,F 2,L are the generalized mass matrix and force matrix of component B2, M3, F3 are the generalized mass matrix and force matrix of component A.

[0159] 4.2 Virtual flight dynamics equations

[0160] During the virtual flight test, the aircraft is fixed by the support system and there is a clamping point. The three translational degrees of freedom constraints at this point are fixed, and the system's independent degrees of freedom become 3+2. The fuselage center of mass vector r1, pitch angle θ, yaw angle ψ, roll angle γ, component A translational displacement s, and component B1 / B2 rotation angle χ are not independent, and the descriptive variables need to be tested for independence.

[0161] 1) Constraint processing

[0162] Assume that the component of the clamping point relative to the fuselage coordinate system is The constraint equations are expressed as

[0163] ① Displacement constraint equation

[0164]

[0165] ②Speed ​​constraint equation

[0166]

[0167] ③Acceleration constraint equation

[0168]

[0169] Among them, r jx is the position coordinate of the clamping point.

[0170] 2) Kinetic equation

[0171] According to the above constraint equation, we can get the relationship matrix between the variables virtual velocity and acceleration, which can be expressed as follows:

[0172]

[0173] Among them, T jr is the relationship matrix between the velocity of the center of mass of the fuselage and the angular velocity of the attitude, α jr is the residual matrix between the body mass center acceleration and the attitude angular acceleration, T c is the conversion matrix between the first-order time derivative of the system description variable and the first-order time derivative of the independent description variable, α c is the residual matrix between the second-order time derivatives of the system description variables and the second-order time derivatives of the independent description variables, E 3×3 is a 3×3 identity matrix.

[0174] Combining the system virtual power equation and the constraint relationship matrix, the independent description variable can be obtained The corresponding virtual power equation, and then the system dynamics equation, is as follows:

[0175]

[0176] Where M1, F1 are the generalized mass matrix and force matrix of the fuselage, M 2,R ,F 2,R is the generalized mass matrix and force matrix of component B1, M 2,L ,F 2,L are the generalized mass matrix and force matrix of component B2, M3, F3 are the generalized mass matrix and force matrix of component A.

[0177] According to the formula - we can find the term describing the variable acceleration A closed second-order dynamic differential equation is formed, which meets the completeness of dynamic solution.

[0178] In the embodiment of the present invention, for step five:

[0179] The constructed multi-body dynamics model is incorporated into the traditional flight mechanics description framework. Therefore, the system description variables are reselected as

[0180] ① The radius of the center of mass of the fuselage r1;

[0181] ② Speed ​​V, angle of attack α, sideslip angle β;

[0182] ③ Pitch angle θ, yaw angle ψ, roll angle γ;

[0183] ④ Time derivative of pitch angle Time derivative of yaw angle Roll angle time derivative

[0184] ⑤ Translation s, rotation angle χ;

[0185] ⑥Translation speed Corner speed

[0186] According to the traditional flight mechanics description framework, the velocity coordinate system can be regarded as the fuselage coordinate system first rotating around the y1 axis by a sideslip angle β, and then rotating around the z1′ axis corresponding to the original z1 axis after the first rotation by an attack angle α. The relationship between the two can be described as

[0187]

[0188] in, are the three basis vectors of the velocity coordinate system; x1, y1, z1 are the three basis vectors of the fuselage coordinate system. Then the first-order time derivative of the center of mass of the fuselage can be expressed as

[0189]

[0190] The second-order time derivative can be expressed as

[0191]

[0192] Then we can get

[0193]

[0194] In the formula

[0195] b1=-V(sinαcosβx1+cosαcosβy1)

[0196] b2=V(sinαsinβy1-cosαsinβx1+cosβz1)

[0197] It can be seen that the above content gives the following important relationship: given the velocity V, angle of attack α, and sideslip angle β, find the first-order time derivative of the center of mass vector of the fuselage The second-order time derivative of the center of mass radius of the fuselage is known. Find the magnitude of acceleration Time derivative of angle of attack Sideslip angle time derivative Combining the dynamic equations and the above relationship, the first-order time derivative describing the variable can be obtained. Thus, the first-order solution model of the dynamic equations of free flight and virtual flight tests under the framework of traditional flight mechanics description is completed.

[0198] The above technical scheme is different from the traditional fixed shape aircraft dynamics modeling method and the general variant aircraft dynamics modeling method, mainly reflected in: 1) The present invention gives full play to the advantages of the combination of the basic principle of dynamics modeling + multi-body system dynamics modeling method, on the one hand, it utilizes the universality of the virtual power principle, on the other hand, it introduces the mathematical description of the topological configuration of the multi-body system, and uses the vector transfer relationship between components to make the modeling more systematic. 2) The present invention not only considers the external variable configuration of the conventional variable sweep (i.e., component B1\B2), but also considers the internal variable configuration of the internal object (i.e., component A), as well as the dynamic influence of the mutual coupling between them. The variant is more complex and the factors considered are more comprehensive. 3) From the perspective of overall and sub-professional coordination and engineering application, the present invention re-incorporates the constructed multi-body dynamics model into the traditional flight mechanics description framework, forming a dynamics model of the multi-body system of the deformable aircraft under the traditional flight mechanics description framework, which is more applicable. 4) The present invention can take into account the refined characterization of numerical simulation and the concise description of engineering applications, avoiding the errors caused by the decoupling of the translational and rotational coupling terms of the mass matrix of the rigid body dynamics model under the traditional flight mechanics description framework.

[0199] Features described and / or illustrated above for one embodiment may be used in the same or similar manner in one or more other embodiments, and / or combined with or used in place of features in other embodiments.

[0200] It should be emphasized that the term "include / comprises" when used herein refers to the presence of features, integers, steps or components, but does not exclude the presence or addition of one or more other features, integers, steps, components or combinations thereof.

[0201] The above method of the present invention can be implemented by hardware, or by hardware combined with software. The present invention relates to such a computer-readable program, which, when executed by a logic component, enables the logic component to implement the above-mentioned device or component, or enables the logic component to implement the above-mentioned various methods or steps. The present invention also relates to a storage medium for storing the above program, such as a hard disk, a magnetic disk, an optical disk, a DVD, a flash memory, etc.

[0202] The many features and advantages of these embodiments are apparent from this detailed description, and thus the appended claims are intended to cover all such features and advantages of these embodiments that fall within their true spirit and scope. Furthermore, since numerous modifications and changes will readily occur to those skilled in the art, it is not intended that the embodiments of the invention be limited to the exact construction and operation illustrated and described, but rather all suitable modifications and equivalents falling within the scope thereof are intended to be covered.

[0203] Parts of the present invention that are not described in detail are well known to those skilled in the art.

Claims

1. A method for dynamic modeling of an aircraft with a multi-degree-of-freedom actuator, characterized in that: The modeling method comprises: Step 1, obtaining the relative motion form between the rigid bodies of the aircraft containing multi-degree-of-freedom actuators, and kinematically describing the rigid bodies of the aircraft and the position of the pressure center according to the vector transfer relationship and the principle of angular velocity superposition, wherein the aircraft includes two types of actuators: one is a sliding actuator, which controls component A to move axially with a displacement s; the other is a rotating actuator, which controls component B1 and component B2 to rotate along the fuselage around the rudder axis to change the aerodynamic layout, and the rigid bodies of the aircraft include the fuselage, component B1, component B2 and component A; Step 2: Establish a conversion relationship matrix between kinematic parameters involved in the virtual power principle and system state variables based on the kinematic description obtained in step 1; Step 3: Based on the virtual power principle and the conversion relationship matrix obtained in step 2, the virtual power equations of each rigid body with the virtual velocity of the system state variable as the virtual variation are obtained, and the virtual power equations of each rigid body are superimposed to obtain the virtual power equation of the system; Step 4: Obtain the free flight simulation dynamics equation and the virtual flight test dynamics equation of the aircraft multi-body system according to the system virtual power equation obtained in step 3 to construct a multi-body dynamics model Step 5: Incorporate the multi-body dynamics model obtained in step 4 into the original flight mechanics description framework to complete the aircraft dynamics modeling.

2. The method for dynamic modeling of an aircraft with a multi-degree-of-freedom actuator according to claim 1, characterized in that: In step 1, the fuselage dynamics are described as follows: The conversion relationship between the fuselage coordinate system and the ground coordinate system is expressed as: Among them, x1, y1, z1 are the three basic vectors of the fuselage coordinate system; x, y, z are the three basic vectors of the ground coordinate system; the rotation of the fuselage coordinate system relative to the ground coordinate system is decomposed into three fixed-axis rotations, ① the first rotation is around the y axis of the ground coordinate system to rotate the yaw angle ψ; ② the second rotation is around the z′ axis corresponding to the original z axis after the first rotation to rotate the pitch angle ③ The third rotation is a roll angle γ around the x″ axis corresponding to the x axis after the second rotation; the x″ axis is the x1 axis of the fuselage coordinate system; R1 is the rotation matrix; According to the angular velocity superposition principle, the angular velocity ω1 is the sum of the product of the rate of change of the rotation angle and the axis of rotation, as shown in the following formula: 。 3. The method for dynamic modeling of an aircraft with a multi-degree-of-freedom actuator according to claim 2, characterized in that: In step 1, components B1 and B2 are dynamically described in the following way: 1) The conversion relationship between the component B1 / B2 coordinate system and the ground coordinate system is expressed as Among them, x2, y2, z2 are the three basic vectors of the component B1 / B2 coordinate system respectively; the variable sweep motion of the aircraft is realized by the rotation motion of the component B1 / B2 around the rotation axis, and the component B1 / B2 coordinate system is regarded as the body coordinate system rotated around the y axis by an angle of x, wherein the component B1 rotation angle x is positive, the component B2 rotation angle x is negative, and the two are opposite numbers; R2 is the rotation matrix; 2) According to the principle of angular velocity superposition, the angular velocity ω2 of component B1 / B2 is: 3) The components of the centroid of component B1 / B2 in the ground coordinate system are expressed as Its time derivative is expressed as: Among them, the hinge point of the rotation connection between component B1 / B2 and the fuselage is A, and the body hinge vector component fixed in the fuselage coordinate system is The hinge vector components of the body fixed in the B1 / B2 coordinate system are: The component of the center of mass of the fuselage in the ground coordinate system is r1.

4. The method for dynamic modeling of an aircraft with a multi-degree-of-freedom actuator according to claim 3, characterized in that: In step 1, component A is dynamically described in the following way: The real-time position of the centroid of component A in the ground coordinate system is expressed as: Its time derivative can be: Among them, in the initial state, the component of the center of mass of component A in the fuselage coordinate system is Afterwards, it moves in a straight line along the axial direction at a speed v and a straight line displacement s. During the movement, the coordinate system of component A is always consistent with the coordinate system of the fuselage, and its angular velocity and angular acceleration are the same as those of the fuselage.

5. The method for dynamic modeling of an aircraft with a multi-degree-of-freedom actuator according to claim 4, characterized in that: In step 1, the compression center is dynamically described in the following way: Assume that the coordinates of the pressure center relative to the fuselage coordinate system are Its components in the ground coordinate system are Its time derivative is expressed as 。 6. The method for dynamic modeling of an aircraft with a multi-degree-of-freedom actuator according to claim 5, characterized in that: In step 2, the conversion relationship between each rigid body kinematic parameter and the system description variable is as follows: 1) Relationship between the virtual angular velocity of the fuselage and the virtual angular velocity of the attitude 2) Relationship between body angular acceleration and attitude angular acceleration 3) Relationship between virtual angular velocity of component B1 and virtual angular velocity of attitude and virtual rotation angular velocity 4) Relationship between the virtual center of mass velocity of component B1 and the virtual center of mass velocity of the fuselage, attitude virtual angular velocity, and virtual rotation angular velocity 5) Relationship between component B1 angular acceleration and attitude angular acceleration and angular acceleration 6) Relationship between the mass center acceleration of component B1 and the mass center acceleration of the fuselage, attitude angular acceleration, and angular acceleration 7) Relationship between component B2 virtual angular velocity and attitude virtual angular velocity and virtual rotation angular velocity 8) Relationship between the virtual center of mass velocity of component B2 and the virtual center of mass velocity of the fuselage, attitude virtual angular velocity, and virtual rotation angular velocity 9) Relationship between component B2 angular acceleration and attitude angular acceleration and rotational angular acceleration 10) Relationship between the mass center acceleration of component B2 and the mass center acceleration of the fuselage, attitude angular acceleration, and angular acceleration 11) Relationship between virtual angular velocity of component A and virtual angular velocity of attitude 12) Relationship between the virtual center of mass velocity of component A and the virtual center of mass velocity of the fuselage, attitude virtual angular velocity, and virtual slip velocity 13) Relationship between component A angular acceleration and attitude angular acceleration 14) Relationship between the center-of-mass acceleration of component A and the center-of-mass acceleration of the fuselage, attitude angular acceleration, and slip acceleration The conversion matrix and acceleration margin matrix involved are defined as follows Table 1 Definition of transformation matrix between rigid body kinematic parameters and system description variables 7. A method for dynamic modeling of an aircraft with a multi-degree-of-freedom actuator according to any one of claims 1 to 6, characterized in that: In step 3, the system virtual power equation is obtained by the following formula: Where M1, F1 are the generalized mass matrix and force matrix of the fuselage, M 2,R ,F 2,R is the generalized mass matrix and force matrix of component B1, M 2,L ,F 2,L is the generalized mass matrix and force matrix of component B2, M3, F3 are the generalized mass matrix and force matrix of component A, and the specific representation is as follows: Where m1 is the mass of the fuselage, m2 is the mass of components B1 and B2, J1 is the moment of inertia of the fuselage, and J 2,R is the moment of inertia of component B1, J 2,L is the moment of inertia of component B2, J3 is the moment of inertia of component A, F air is the aerodynamic force vector borne by the fuselage, M χ,R is the control torque of component B1, M χ,L is the control torque of component B2, F s is the control thrust of component A, g is the acceleration due to gravity, T sr is the conversion matrix between the virtual velocity of the center of pressure and the virtual angular velocity of the attitude, E 3×3 is a 3×3 identity matrix.

8. A method for dynamic modeling of an aircraft with a multi-degree-of-freedom actuator according to any one of claims 1 to 7, characterized in that: The step 4 specifically includes: 4.1 Determine whether the aircraft is constrained. If the aircraft is not constrained, go to step 4.2; if the aircraft is constrained, go to step 4.3; 4.2 According to the system virtual power equation, remove the virtual variation to obtain the system free flight simulation dynamic equation; 4.3 According to the specific constraint form, constraint processing is performed, in which the state variables are divided into independent terms and dependent terms, and the conversion relationship between independent terms and dependent terms is established. The independent terms are substituted into the system virtual power equation to obtain the system virtual power equation corresponding to the independent virtual variation. The independent virtual variation is removed to obtain the system virtual flight test dynamics equation.

9. A computer device comprising a memory, a processor and a computer program stored in the memory and executable on the processor, characterized in that: When the processor executes the computer program, the prediction method according to any one of claims 1 to 8 is implemented.

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