Control method and system for a variable attitude aircraft

Abstract

The application discloses a control method and system of a variable posture aircraft, and comprises the following steps: step one, defining a commonly used coordinate system of the aircraft, relative distances of key points and commonly used motion variables; step two, establishing a multi-rigid-body six-degree-of-freedom dynamics model of the variable posture aircraft based on the Kane method; step three, performing Jacobian linearization processing on the dynamics model of the aircraft; step four, based on the linearized model, obtaining system state feedback gain by using a tensor product modeling method combined with a PDC-LMI thought, so as to design a flight control law. The method provided by the application can simplify the control law design steps of a nonlinear complex variable posture aircraft, and ensures the attitude consistency stability of the aircraft in different change processes of the down angle under the premise of initial disturbance, and has good engineering value.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of flight control, in particular to a control method and system of a variable attitude aircraft. BACKGROUND

[0002] With the continuous development of modern science and technology, aircraft has made great progress in maneuverability and reliability, among which the variable geometry aircraft can adapt to different flight conditions and task requirements by actively changing the body geometry, has high development potential, and is the focus of future aircraft development.

[0003] In the field of variable geometry aircraft control, there are many time-varying parameters and the parameters are closely related to each other, which poses a formidable challenge to the attitude control technology of the aircraft. The LPV strategy has a wide range of applications in time-varying systems, by obtaining a linearized model under the scheduling of each time-varying parameter and using linear control to obtain control gain, but this method has the problem of unreasonable selection of working points when facing densely sampled working points, and the subsequent controller calculation efficiency problem cannot be ignored.

[0004] Therefore, how to develop a simple and reliable polytope LPV control design method for variable attitude aircraft has become a technical problem that needs to be solved by those skilled in the art. SUMMARY

[0005] The present application provides a variable attitude aircraft control method based on polytope LPV, which solves the problem of control difficulty of the aircraft under the premise of initial disturbance.

[0006] The technical problem solved by the present application adopts the following technical scheme

[0007] A control method of a variable attitude aircraft, characterized in that the method adopts a tensor product modeling method based on a linearized model combined with a PDC-LMI idea to obtain system state feedback gain and design a flight control law.

[0008] Further, a control method of a variable attitude aircraft, comprising the following steps:

[0009] Step (1), defining the commonly used coordinate system of the aircraft and the relative distance of the key points and the commonly used motion variables;

[0010] Step (2), establishing a multi-rigid-body six-degree-of-freedom dynamics model of the variable down angle aircraft based on the Kane method;

[0011] Step (3), performing Jacobian linearization processing on the aircraft dynamics model;

[0012] Step (4), based on the linearized model, the system state feedback gain is obtained by using the tensor product modeling method combined with the PDC-LMI idea, so as to design the flight control law.

[0013] Further, the method for defining the commonly used coordinate system of the aircraft in step (1) is as follows: a ground coordinate system (x g ,y g ,z g ) is defined, the coordinate origin O is at the launch point of the aircraft, the Ox g axis can point to any direction in the horizontal plane, the Oy g axis is plumb upward, the Oz g axis is perpendicular to the Ox g axis and the Oy g axis to form a right-handed coordinate system; a body coordinate system (x b ,y b ,z b ) is defined, the coordinate origin O b is at the center of mass of the aircraft, the Ox b axis coincides with the longitudinal axis of the aircraft and points to the head as positive, the Oy b axis is in the longitudinal symmetry plane of the aircraft and is perpendicular to the Ox b axis and points upward as positive, the Oz b axis is perpendicular to the Ox b y b plane and its direction is determined according to the right-handed rectangular coordinate system; a trajectory coordinate system (x k ,y k ,z k ) is defined, the coordinate origin O k is at the instantaneous center of mass of the aircraft, the Ox k axis coincides with the velocity vector V of the aircraft, the Oy k axis is in the plumb plane containing the velocity vector V and is perpendicular to the Ox k axis and points upward as positive, the Oz k axis is perpendicular to the Ox k axis and the Oy k axis to form a right-handed coordinate system; a velocity coordinate system (x a ,y a ,z a ) is defined, the coordinate origin O a is at the center of mass of the aircraft, the Ox a axis coincides with the velocity vector V of the center of mass of the aircraft, the Oy a axis is in the longitudinal symmetry plane of the aircraft and is perpendicular to the Ox a axis and points upward as positive, the Oz a axis is perpendicular to the Ox a y a plane and its direction is determined according to the right-handed rectangular coordinate system.

[0014] Further, the step (1) defining the key point relative distance of the aircraft includes:

[0015] Ox b Axial includes: the distance between the body mass center and the midpoint of the winglet connecting line;

[0016] Oz b Axial includes: the distance between the midpoint of the winglet connecting line and the winglet connecting point, and the distance between the winglet connecting point and the winglet mass center.

[0017] Further, the step (1) defining the commonly used motion variables of the aircraft includes:

[0018] Attitude angle includes: pitch angle, yaw angle and roll angle;

[0019] Trajectory angle includes: trajectory inclination angle, trajectory deflection angle and trajectory tilt angle;

[0020] Deformation amount includes: wing tip winglet down angle.

[0021] Further, the step (2) of establishing a six-degree-of-freedom dynamics model of the variable down angle aircraft is based on the Kane method for multi-rigid-body modeling. The Kane method is based on the principle of D'Alembert, and the generalized velocity is selected to describe the motion of the object.

[0022] Further, the step (3) of Jacobian linearization processing of the aircraft dynamics model is obtained by using the Taylor series expansion method near the equilibrium point to obtain the linearization model of the original system expanded at different working points.

[0023] Further, the step (4) is based on the linearization model, and the tensor product modeling method is combined with the PDC-LMI idea to obtain the state feedback gain of the vertex control system by solving a series of matrix inequalities, and the gain of the vertex outer system is obtained by means of interpolation and weighted combination. Finally, the control law of the whole flight system is obtained.

[0024] A control system of a variable attitude aircraft, the system includes the following modules:

[0025] Space modeling module: for defining the commonly used coordinate system, key point relative distance and commonly used motion variables of the aircraft;

[0026] Mechanical modeling module: for establishing a multi-rigid-body six-degree-of-freedom dynamics model of the variable down angle aircraft;

[0027] Mechanical model processing module: for Jacobian linearization processing of the aircraft dynamics model;

[0028] Data processing module: for obtaining system state feedback gain based on linearization model, using tensor product modeling method combined with PDC-LMI idea, so as to design flight control law.

[0029] A storage device stores a polyhedral LPV-based variable posture aircraft control method.

[0030] The beneficial effects of the present application are:

[0031] 1、The present application selects the Kane method for multi-rigid-body modeling, compared with the traditional Newton-Euler method, the Kane method ignores the constraint force and interaction force between each rigid body, and the modeling is more efficient; compared with the Lagrange method, the Kane method does not use the energy function, and does not need to calculate the differential of the scalar energy function (including kinetic energy and potential energy), so it is more suitable for computer programming implementation and engineering application;

[0032] 2、The present application is based on a polyhedral LPV system, combines the PDC framework and the LMI method, obtains the linearization model of the complex nonlinear system through Jacobian linearization, obtains eight polyhedral vertex systems that can approximate the original system by means of high-order singular value decomposition method, realizes the design of the final control law by solving the gain of the vertex system, and greatly improves the calculation efficiency of the model and the robust performance of the control system. BRIEF DESCRIPTION OF DRAWINGS

[0033] Figure 1 A flowchart of the polyhedral LPV-based variable posture aircraft control method of the present application;

[0034] Figure 2 A simplified model and key point coordinate diagram of the variable posture aircraft of the present application;

[0035] Figure 3 An initial singular vector diagram obtained through high-order singular value decomposition of the present application;

[0036] Figure 4 A singular vector diagram meeting the SN and NN requirements of the present application;

[0037] Figure 5 A singular vector diagram meeting all requirements of the present application. DETAILED DESCRIPTION

[0038] In order to make the purpose, technical scheme and advantages of the embodiments of the present application clearer, the technical scheme in the embodiments of the present application will be described clearly and completely below with reference to the drawings in the embodiments of the present application. Obviously, the described embodiments are part of the embodiments of the present application, rather than all the embodiments. Based on the embodiments in the present application, all other embodiments obtained by those skilled in the art without creative labor fall within the scope of protection of the present application.

[0039] Embodiment 1: Reference Figure 1 The application discloses a variable posture aircraft control method based on a polyhedral LPV, comprising the following steps:

[0040] Step (1), defining a commonly used coordinate system of the aircraft and key point relative distance and commonly used motion variables;

[0041] Step (2), establishing a multi-rigid-body six-degree-of-freedom dynamics model of the variable posture aircraft based on the Kane method; the Kane method is selected for multi-rigid-body modeling, in the process of establishing the system dynamics equation, the interaction force and the constraint force between various systems do not need to be considered, and the derivative of the energy function of the system with respect to time does not need to be calculated, so that the modeling efficiency and the engineering implementation capability are improved;

[0042] Step (3), performing Jacobian linearization processing on the aircraft dynamics model, so as to provide a linearization model basis for subsequent controller design;

[0043] Step (4), based on the linearization model, adopting a tensor product modeling method combined with a PDC-LMI idea to obtain system state feedback gain, so as to design a flight control law.

[0044] The method for defining the commonly used coordinate system of the aircraft in the step (1) is as follows: a ground coordinate system (x g ,y g ,z g ) is defined, a coordinate origin O is on a launch point of the aircraft, an Ox g axis can point to an arbitrary direction in a horizontal plane, an Oy g axis is plumb upward, and an Oz g axis is perpendicular to the Ox g axis and the Oy g axis to form a right-hand coordinate system; a body coordinate system (x b ,y b ,z b ) is defined, a coordinate origin O b is on a mass center of the aircraft, an Ox b axis is coincided with a longitudinal axis of the aircraft, points to a head portion and is positive, an Oy b axis is located in a longitudinal symmetry plane of the aircraft and is perpendicular to the Ox b axis, points upward and is positive, and an Oz b axis is perpendicular to an Ox b y b plane, and its direction is determined according to a right-hand rectangular coordinate system; a trajectory coordinate system (x k ,y k ,z k ) is defined, a coordinate origin O k is taken on an instantaneous mass center of the aircraft, an Ox k axis is coincided with a velocity vector V of the aircraft, and an Oyk The axis lies in a plumb plane containing the velocity vector V and is perpendicular to Ox. k The axis, pointing upwards, is positive (Oz). k Axis and Ox k Axis and Oy k The axes perpendicularly form a right-handed coordinate system; define a velocity coordinate system (x... a ,y a ,z a O, the origin of the coordinate system a Taken at the center of mass of the spacecraft, Ox a The velocity vector V of the axis coincides with that of the spacecraft's center of mass, Oy a The axis is located within the longitudinal symmetry plane of the aircraft and intersects with Ox. a The axis is vertical, with upward pointing being positive, Oz a The axis is perpendicular to Ox a y a The plane, whose orientation is determined according to the right-hand rectangular coordinate system.

[0045] Specifically, commonly used coordinate systems include:

[0046] 1) Ground coordinate system S g

[0047] Ground coordinate system S g It is a coordinate system fixed to the Earth's surface, with the origin O usually chosen at the launch point of the spacecraft. g The axis can point in any direction within the horizontal plane, Oy g The shaft is plumb upward, Oz g Axis and Ox g Axis and Oy g The axes perpendicularly form a right-handed coordinate system.

[0048] 2) Body coordinate system S b

[0049] Body coordinate system S b The origin O is taken at the center of mass of the spacecraft, Ox b The axis coincides with the longitudinal axis of the aircraft, with the direction pointing towards the head being positive. b The axis is located within the longitudinal symmetry plane of the aircraft and intersects with Ox. b The axis is vertical, with upward pointing being positive, Oz b The axis is perpendicular to Ox b y b The plane, whose orientation is determined according to the right-hand rectangular coordinate system.

[0050] 3) Ballistic coordinate system S k

[0051] Ballistic coordinate system S k The origin O is taken at the instantaneous center of mass of the spacecraft, Ox kThe axis coincides with the velocity vector V of the aircraft, Oy k The axis lies in a plumb plane containing the velocity vector V and is perpendicular to Ox. k The axis, pointing upwards, is positive (Oz). k Axis and Ox k Axis and Oy k The axes perpendicularly form a right-handed coordinate system.

[0052] 4) Velocity coordinate system (airflow coordinate system) S a

[0053] Velocity coordinate system S a The origin O is taken at the center of mass of the spacecraft, Ox a The velocity vector V of the axis coincides with that of the spacecraft's center of mass, Oy a The axis is located within the longitudinal symmetry plane of the aircraft and intersects with Ox. a The axis is vertical, with upward pointing being positive, Oz a The axis is perpendicular to Ox a y a The plane, whose orientation is determined according to the right-hand rectangular coordinate system.

[0054] The commonly used motion variables of the aircraft defined in step (1) include:

[0055] 1) Attitude angle

[0056] Attitude angles represent the body coordinate system S b With ground coordinate system S g The conversion relationships between them include:

[0057] A) Pitch angle The longitudinal axis of the aircraft (Ox b (axis) and horizontal plane (Ox) g z g The angle between the two planes is positive when the longitudinal axis of the aircraft points upwards onto the horizontal plane.

[0058] B) Yaw angle ψ: Aircraft longitudinal axis (Ox b (axis) in the horizontal plane (Ox) g z g Projection in the plane and ground coordinate system Ox g The angle between the axes, facing Oy g Viewed from the top of the axis, if Ox g If the axis rotates counterclockwise to the projection, then the angle ψ is positive.

[0059] C) Roll angle γ: Body coordinate system S b Oy b The angle between the longitudinal axis and the plumb plane containing the longitudinal axis of the aircraft, viewed forward along the longitudinal axis from the tail of the aircraft, if Oy b If the axis is located on the right side of the vertical plane, then the angle γ is positive.

[0060] 2) Ballistic angle

[0061] Ballistic angle represents the transformation relationship between ballistic coordinate system S k and ground coordinate system S g and velocity coordinate system S a , including:

[0062] A) Ballistic angle of inclination θ: the angle between the velocity vector V(Ox k axis) of the aircraft and the horizontal plane (Ox g z g plane), the velocity vector points upwards, and the θ angle is positive;

[0063] B) Ballistic angle of declination ψ V : the angle between the projection of the velocity vector V(Ox k axis) of the aircraft in the horizontal plane (Ox g z g plane) and the Ox g axis of the ground coordinate system S g , looking down the Oy g axis, if the projection is counterclockwise rotation from the Ox g axis, then the ψ V angle is positive;

[0064] C) Ballistic angle of inclination μ: the angle between the Ox a axis in the longitudinal symmetry plane of the aircraft and the vertical plane Ox k y k containing the velocity vector V, looking forward from the tail of the aircraft, if the longitudinal symmetry plane is inclined to the right, then the μ angle is positive.

[0065] 3) Airflow angle

[0066] Airflow angle represents the transformation relationship between velocity coordinate system S a and body coordinate system S b , including:

[0067] A) Angle of attack (angle of incidence) α: the angle between the projection of the velocity vector V(Ox a axis) of the aircraft's center of mass on the longitudinal symmetry plane Ox g y g of the body and the Ox g axis, if the Ox g axis is above the projection line of V (i.e. generates lift), then the α angle is positive;

[0068] B) Angle of sideslip β: the angle between the velocity vector V(Ox a axis) and the longitudinal symmetry plane, looking in the direction of flight, if the incoming flow flows from the right side of the body (i.e. generates negative lateral force), then the β angle is positive.

[0069] 4) Downward negative angle Γ: the angle between the body coordinate system S b Oz b axis and the winglet longitudinal axis, the winglet is downwardly biased Γ angle is positive.

[0070] The step (2) is based on the Kane method to establish a variable down angle aircraft multi-rigid body six degree of freedom dynamics model;

[0071] Specifically, the Kane method is used for multi-rigid body modeling, which is also known as "Lagrange form of D'Alembert's principle". Based on D'Alembert's principle, the generalized velocity is selected to describe the motion of the object, and the interaction force and constraint force between the systems are not considered in the process of establishing the system dynamics equation, and the derivative of the energy function of the system with respect to time is not calculated.

[0072] The structure of the variable down angle aircraft and the coordinate diagram are shown in Figure 2 The variable down angle aircraft is divided into three rigid bodies: the main body, the left winglet and the right winglet, with masses m b , m, m, and the total mass of the aircraft is m T . Assuming that the center of mass of the fuselage is O b , the origin of the body coordinate system, O b x b , O b y b , O b z b are the three axes of the body coordinate system. The winglet is regarded as a uniform rectangular thin surface, P and Q are the connection points of the fuselage and the left and right wings, M and N are the connection points of the left and right wings and the winglet, and the winglet can rotate in the O b y b z b plane, and it is assumed that the downward negative angles of the left and right winglets are the same, both being Γ. The longitudinal distance from the center of mass of the fuselage O b to MN is h, and the lateral distance from the center of mass of the fuselage O b to M and N is i, and the winglet has a spanwise length of l. The fuselage and wings always move in the same plane.

[0073] The six degree of freedom mathematical model of the aircraft is as follows:

[0074]

[0075] Where V x , V y , V z are the velocity components of the body center of mass velocity in the body axes; ω x , ω y, ω z are the angular velocities of the body coordinate system x, y, z axes respectively; V is the vehicle mass center speed; α, β, μ are the angle of attack, sideslip angle and ballistic tilt angle of the variable attitude vehicle respectively; θ is the ballistic tilt angle; ψ v is the ballistic deflection angle; M x , M y , M z are the roll moment, yaw moment and pitch moment of the vehicle respectively; L, D, C are the lift, drag and lateral force acting on the variable attitude vehicle respectively; m T is the total mass of the vehicle.

[0076] c1~c 10 The expressions are as follows.

[0077]

[0078] Wherein, h is the distance from the body mass center to MN in the O b x b direction; i is the distance from the body mass center to M, N in the O b z b direction; m is the small wing mass; l is the small wing spanwise length; Γ and are the downwash angle and the downwash angle change speed (clockwise rotation is positive along the negative x b axis) respectively. J bx , J by , J bz are the moments of inertia in the body coordinate system x, y, z axes respectively.

[0079] The step (3) is to perform Jacobian linearization processing on the vehicle dynamics model. For a nonlinear system, the classical linearization processing method is to perform Taylor expansion at the equilibrium point of the system to obtain a linear approximation, so as to obtain the corresponding Jacobian matrix. Similarly, feedback linearization, including coordinate transformation and state feedback or output feedback method, can also realize the establishment of LPV model. The present application selects the classical Jacobian linearization method, and the subsequent controller design is mainly for the Jacobian matrix.

[0080] Specifically, in order to ensure the accurate tracking of the command by the airflow angle α, β and the ballistic tilt angle μ and the stabilization of the angular velocity itself, the state quantity of the controlled system is selected as x=[α, β, μ, ω x , ω y , ω z ] T , and the control quantity is selected as u=[δ e , δ a ] T, respectively, are generated by the FLAP rudders at the tail of the aircraft, and the error amount is defined as Δx = x - x s , Δu = u - u s The longitudinal fixed trim angle of attack is selected as 20°, and the lateral balance point is zero. The scheduling variables are selected as the down angle velocity and the down angle, and the linearized state space model of the aircraft is established as follows:

[0081]

[0082] The Jacobian matrices A and B are function matrices of the scheduling variables, and the specific expressions are as follows:

[0083]

[0084] The specific elements are represented as follows:

[0085]

[0086]

[0087] The step (4) is based on the linearized model, and the tensor product modeling method is combined with the PDC-LMI idea to obtain the system state feedback gain, so as to design the flight control law.

[0088] Specifically, the scheduling variables The state space equation of the variable attitude aircraft after the substitution is as follows:

[0089]

[0090] The system matrix is defined as:

[0091] S = [A[p(t)], B[p(t)]]

[0092] The system matrix is discretized, the scheduling variable space p(t) ∈ Ω = [-10, 10] × [0, 90] is defined, and the grid point density is selected as 41 × 91, that is, the down angle velocity range is -10 10° / s, and each 0.5° / s is divided into an area, and the down angle amplitude range is 0 90°, and each 1° is divided into an area. The system matrix is sampled on the given discrete grid points and stored in the high-order tensor S D

[0093] Wherein, O and I are the sizes of the system matrix itself, M1 and M2 represent the number of down angle velocity and down angle grid points, respectively.

[0094] Considering the change of the shape parameter on the system matrix, only S D ​The last two dimensions (M1, M2) are subjected to high-order singular value decomposition (HOSVD) to obtain an LTI vertex system, and singular values and corresponding singular vectors with small numerical values or equal to 0 are discarded, and finally non-zero singular values are obtained, as shown in Table 1.

[0095] Table 1 Non-zero singular values

[0096]

[0097] Non-zero singular values of two speed dimensions and three angle dimensions are reserved respectively, and a core tensor of the system is obtained The original system is approximated as:

[0098]

[0099] U n The singular vector corresponding to the n-th dimension of the high-order tensor is represented, and the approximation error e is determined by the following formula:

[0100]

[0101] The singular vector obtained through HOSVD is shown in Table 2. Figure 3

[0102] To ensure that the vertex system composed of the selected working points can meet the design requirements of the subsequent controller, the singular vector needs to be transformed into SN (Sum Normalized), NN (Non-Negative) and NO (Normal) types. To meet the SN requirement, SN conversion matrices T_SN_3 and T_SN_4 are designed for U3 and U4 respectively as follows:

[0103] T_SN_3 = I 2×2 + [sum((U3 (r) ) T )-1 2×1 ,0 2×1 ]

[0104] T_SN_4 = pinv(U4 (r) )U4'

[0105] Wherein:

[0106]

[0107] Φ a = diag[sum((U4 (r) ) T ]

[0108] To meet the NN requirement, NN conversion matrices are designed for U3 and U4 respectively as follows:

[0109] ​

[0110] After satisfying the above two requirements, the singular vector is shown as Figure 4 To meet the orthogonal requirement, the singular vector needs to be reduced dimensionally from the original n col dimensional space to n col -1 dimensional space, which can be achieved by multiplying the initial vector by the conversion matrix DR, and an additional vertex needs to be added in the angle dimension to achieve a strict NO conversion. The singular vector after completing the entire conversion is shown as Figure 5

[0111]

[0112] Since the sampling points are discrete, and the actual running state values are continuously changing, a weight function that changes continuously with the scheduling variable is needed to ensure that the polytope system can exhibit the original system characteristics. The singular vector obtained by the above transformation has met the weight function standard, so only the core tensor needs to be converted. The specific steps are as follows:

[0113]

[0114] where: the new core tensor is Finally, an eight-vertex system is obtained. Combined with the PDC-LMI idea, the dynamic characteristics of the original nonlinear system can be represented by the vertex system contained in the core tensor, that is:

[0115]

[0116] where represents the weight function of the corresponding dimension. Only the state feedback gain needs to be designed for each vertex system, and the control function of the entire control system can be determined by the following formula:

[0117]

[0118] Under the PDC framework, the dynamic characteristics of the system are converted to the following formula:

[0119]

[0120] The control law is also expressed as:

[0121]

[0122] Substituting the above formula into the control law gives:

[0123]

[0124] Further arrangement can be obtained:

[0125] ​

[0126] where: G r,s = A r -B r K s Based on Lyapunov stability theory, if there exists a positive definite matrix P such that the following inequality holds, the system can be made to achieve asymptotic stability:

[0127] G r,r T P+PG r,r <0, r = 1, 2, …, R

[0128]

[0129] Define a new positive definite matrix X = P -1 and a new controller solution matrix M r = K r X, then the controller design criterion based on the convex polyhedral tensor product model is: find a positive definite matrix X and a controller solution matrix M r such that they satisfy the following matrix inequalities:

[0130] -XA r T -A r X+M r T B r T +B r M r >0, r = 1, 2, …, R

[0131] -XA r T -A r X-XA s T -A s X

[0132] +M s T B r T +B r M s +M r T B s T +B s M r ≥0, r < s ≤ R

[0133] Given the initial disturbance α0=5°, μ0=5°, under five different dynamic changes of the lower angle, the airflow angle is uniformly tracked and the angular velocity is uniformly stabilized.

[0134] The present application selects the Kane method for multi-rigid-body modeling, compared with the traditional Newton-Euler method, the Kane method ignores the constraint force and interaction force between each rigid body, and the modeling is more efficient; compared with the Lagrange method, the Kane method does not use the energy function, and does not need to calculate the differential of the scalar energy function (including kinetic energy and potential energy), so it is more suitable for computer programming implementation and engineering application.

[0135] Based on the polytopic LPV system, the PDC framework and the LMI method are combined, the linearization model of the complex nonlinear system is obtained through Jacobian linearization, eight polytopic vertex systems which can approximate the original system are obtained by means of high-order singular value decomposition method, the gain of the vertex system is solved to realize the design of the final control law, the solving step of the control law of the traditional LPV method is greatly simplified, and the robustness of the control system is improved. It should be pointed out that the method proposed in the present application is not only suitable for the variable downward angle aircraft, but also suitable for other flight systems such as different variable configuration aircraft or rotary wing aircraft, and the corresponding scheduling variables are changed for the control problems faced.

[0136] Finally, it should be pointed out that: the above embodiments are only used to illustrate the technical solutions of the present application, and not to limit them; although the present application has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that: it can still modify the technical solutions recorded in the foregoing embodiments, or make equivalent replacement for part of the technical features; and these modifications or replacements do not make the essence of the corresponding technical solutions deviate from the spirit and scope of the technical solutions of the embodiments of the present application.

Claims

1. A control method of a variable configuration aircraft, characterized in that, the method comprises the following steps: Step 1, defining the commonly used coordinate system of the aircraft and the relative distance of the key points and the commonly used motion variables; Step 2, establishing a multi-rigid-body six-degree-of-freedom dynamic model of the variable configuration aircraft; Step 3, performing Jacobian linearization processing on the dynamic model of the aircraft; Step 4, based on the linearized model, using the tensor product modeling method combined with the PDC-LMI idea to obtain the state feedback gain of the system, thereby designing the flight control law; the six-degree-of-freedom dynamic model of the variable configuration aircraft is established based on the Kane method for multi-rigid-body modeling, and the Kane method selects the generalized velocity to describe the motion of the object on the basis of the d'Alembert principle; in Step 3, the Jacobian linearization processing is performed on the dynamic model of the aircraft, and the linearized model of the original system is obtained by using the Taylor series expansion method near the equilibrium point; in Step 4, based on the linearized model, the tensor product modeling method combined with the PDC-LMI idea is used to obtain the state feedback gain of the vertex control system by solving a series of matrix inequalities, and the gain of the vertex outer system is obtained by means of interpolation and weighted combination, and finally the control law of the entire flight system is obtained; the control law is expressed as: 2.The control method of a variable configuration aircraft according to claim 1, characterized in that, the coordinate system in Step 1 includes: ground coordinate system, body coordinate system, trajectory coordinate system, and velocity coordinate system. 3.The control method of a variable configuration aircraft according to claim 1, characterized in that, the relative distance of the key points in Step 1 includes: the distance of the body center of mass to the midpoint of the small wing connecting line, the distance of the midpoint of the small wing connecting line to the small wing connecting point, and the distance of the small wing connecting point to the small wing center of mass. 4.The control method of a variable configuration aircraft according to claim 1, characterized in that, the commonly used motion variables in Step 1 include: attitude angle includes: pitch angle, yaw angle and roll angle; trajectory angle includes: trajectory inclination angle, trajectory deflection angle and trajectory tilt angle; airflow angle includes: angle of attack and side slip angle; deformation amount includes: wing tip small wing down angle.

5. A control system for a variable configuration aircraft, characterized in that, a system for implementing the control method of a variable configuration aircraft according to claim 1; the system comprises the following modules: a space modeling module for defining the commonly used coordinate system of the aircraft and the relative distance of the key points and the commonly used motion variables; a mechanics modeling module for establishing a multi-rigid-body six-degree-of-freedom dynamic model of the variable down angle aircraft; a mechanics model processing module for performing Jacobian linearization processing on the dynamic model of the aircraft; a data processing module for obtaining the state feedback gain of the system based on the linearized model, using the tensor product modeling method combined with the PDC-LMI idea, thereby designing the flight control law. 6.A storage device, characterized in that, the storage device stores the control method of a variable configuration aircraft according to any one of claims 1-4.

Images