CN113885358B - Hybrid configuration fixed wing unmanned aerial vehicle maneuver simulation control law design method - Google Patents
Hybrid configuration fixed wing unmanned aerial vehicle maneuver simulation control law design method Download PDFInfo
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Abstract
The invention discloses a method for designing a hybrid configuration fixed wing unmanned aerial vehicle maneuvering simulation control law, which is used for establishing a six-degree-of-freedom nonlinear mathematical model of the fixed wing unmanned aerial vehicle; constructing a flight control law core based on a nonlinear dynamic inverse method; defining a special maneuver needing to be simulated, and analyzing a maneuver section needing to be simulated; selecting a proper control law mixing configuration according to the maneuvering section; and (5) inputting a control instruction to develop simulation according to the control law configuration design, and finally obtaining a maneuvering simulation numerical result. The method simulates actual movement to the maximum extent, and the design method has universal applicability. The method fully utilizes the flight dynamics knowledge, designs the control law feedback configuration based on the nonlinear dynamic inverse control core according to the maneuvering characteristics of the required simulation, thereby obtaining the corresponding numerical simulation result, and the design method is easy to encode by using programming languages such as C++, and has universal applicability.
Description
Technical Field
The invention relates to the technical field of unmanned aerial vehicle control, in particular to a method for designing a fixed-wing unmanned aerial vehicle trick maneuvering flight control law.
Background
The unmanned aerial vehicle is widely applied to activities such as aerial photography, geographical mapping, image acquisition, aerial inspection, agriculture and forestry spraying, communication relay and the like. However, the existing unmanned aerial vehicle has a lot of flying activities in a stable flying state, and only requires the unmanned aerial vehicle to have low pitching and rolling capabilities. The special flight of the unmanned aerial vehicle is used as the flight activity of performance property, the unmanned aerial vehicle is required to carry out great maneuver including aircraft nose pointing change, heading change and the like, the unmanned aerial vehicle is required to have higher maneuverability and stability, and new requirements are provided for the control law design of the unmanned aerial vehicle.
In order to solve the problem, a special effect flight control system and method of an unmanned aerial vehicle based on a PID control method are disclosed in Chinese patent application No. CN111580537A, an attitude control angle is calculated through inertia moment, an outer loop determines an acceleration instruction by using nonlinear tracking guidance control, an inner loop tracks the acceleration instruction through PI control rate, and the unmanned aerial vehicle with fixed wings can realize high-difficulty flight actions such as agile flight, special effect flight and the like.
Disclosure of Invention
In order to solve the requirement of complex maneuver simulation of the fixed-wing unmanned aerial vehicle, the six-degree-of-freedom hybrid configuration control law is designed in the full-flight envelope according to the motion characteristics of the fixed-wing unmanned aerial vehicle trick flight maneuver, and specific input can be designed according to the motion characteristics of a given maneuver, so that the maneuver simulation problem in the full-envelope is solved, and large-scale parameter adjustment work is avoided.
The invention is realized by the following technical scheme:
a fixed-wing unmanned aerial vehicle trick maneuver simulation control law design method based on nonlinear control comprises the following steps:
step 1, according to characteristic parameters of a fixed-wing unmanned aerial vehicle, including mass, moment of inertia, product of inertia, chord length, span length, wing area, aerodynamic static derivative, dynamic derivative and the like, six-degree-of-freedom modeling is carried out on the fixed-wing unmanned aerial vehicle according to aircraft flight dynamics, and the modeling process comprises a common ground coordinate system, a machine body coordinate system, an airflow coordinate system, a track coordinate system and a conversion matrix among the coordinate systems;
the six-degree-of-freedom nonlinear mathematical model of the fixed-wing unmanned aerial vehicle described in the step 1 has the following basic equation:
wherein m represents the mass of the unmanned aerial vehicle, g represents the local gravity acceleration, alpha and beta represent the attack angle and sideslip angle of the unmanned aerial vehicle respectively, u, v and w represent the triaxial speed components of the unmanned aerial vehicle under the body coordinate system respectively, p, q and r represent the pitching, rolling and yawing angle rates of the unmanned aerial vehicle under the body coordinate system respectively, phi, theta and psi represent the rolling angle, the pitching angle and the yawing angle of the unmanned aerial vehicle respectively, and x represents the three-axis speed components of the unmanned aerial vehicle under the body coordinate system respectively E 、y E 、z E Respectively representing projection positions of the unmanned aerial vehicle in three directions of a ground coordinate system; q 1 To q 4 Representing a quaternion for computing pose, c 1 To c 9 Representing a parameter related to the moment of inertia of the body,and->Respectively represents aerodynamic force and aerodynamic moment of the unmanned aerial vehicle in three directions under the machine body coordinate system, and for X th 、Y th 、Z th And M T 、N T 、L T Respectively representing the engine thrust and the engine thrust moment of the unmanned aerial vehicle in three directions under the machine body coordinate system.
Step 2, a control law core module based on a nonlinear dynamic inverse method is established, a control matrix in any state is obtained through solving by a numerical method, namely, a control matrix F is solved by substituting an input value 0 method, a control matrix G is solved by a numerical differential method, the control law core comprises an inner ring and an outer ring, the outer ring obtains a triaxial angular velocity instruction according to attack angle, sideslip angle and velocity rolling angle instruction solution, and the inner ring obtains a control surface instruction according to the angular velocity instruction solution;
step 3, analyzing a flight profile of the required simulated maneuver, dividing maneuver phases according to kinematic features, extracting key parameters of each phase, determining a flight instruction required to be controlled, and making a maneuver control flow chart;
step 4, determining a required control law configuration according to the control instruction type and the maneuvering control flow chart obtained in the step 3, and then determining a specific numerical value of a flight instruction according to the flight characteristics of the required simulated maneuvering;
and 5, integrating the six-degree-of-freedom body motion model of the fixed-wing unmanned aerial vehicle established in the step 1, the unmanned aerial vehicle control law core established in the step 2 and the control law specific configuration established in the step 4, and completing simulation of the required unmanned aerial vehicle maneuver by combining a flight instruction.
In summary, the core idea of the present invention is to select the most appropriate external control law configuration based on the target selection route, i.e. according to the maneuvering characteristics required to complete the simulation, on the premise of keeping the core technology unchanged, so as to complete the whole closed-loop simulation.
Compared with the prior art, the invention has the following beneficial technical effects:
aiming at the problem of design of a special maneuver flight control law of a mixed configuration of a fixed-wing unmanned aerial vehicle, a nonlinear dynamic inverse control method is adopted, three control law configurations of an attack angle-a roll angle speed-an accelerator, a pitch angle speed-a roll angle speed-an accelerator and a height-a track deflection angle-speed are provided based on three control parameters of the unmanned aerial vehicle, three fixed-wing unmanned aerial vehicle special maneuvers of S turning, S breaking maneuver and barrel rolling maneuver in a horizontal plane are selected, the control quantity required by each section maneuver is determined according to kinematic characteristics by adopting a segmentation analysis method, corresponding control law configurations are selected, the design control parameters are designed, and the accurate modeling and simulation of the fixed-wing unmanned aerial vehicle on the special maneuver in the whole envelope are realized.
Drawings
FIG. 1 is a diagram of a nonlinear dynamic inverse control law core module;
FIG. 2 is a schematic diagram of a control law configuration of a height-roll angle-speed configuration;
FIG. 3 is a schematic diagram of pitch rate-roll angular rate-throttle control law configuration;
FIG. 4 is a schematic view of the angle of attack-roll angle speed-throttle control law configuration;
FIG. 5 is a schematic illustration of a drum maneuver;
FIG. 6 is a flow chart of a drum roll motor control law design;
FIG. 7 is a simulated time domain response of a simulated drum maneuver;
fig. 8 is a three-dimensional simulated track of a barrel roll maneuver.
Detailed Description
The technical contents of the present invention will be described in detail with reference to the accompanying drawings and specific examples.
Aiming at the requirements of unmanned aerial vehicle trick flight control, the invention adopts a Nonlinear Dynamic Inversion (NDI) method to design unmanned aerial vehicle flight control law, and the basic structure is shown in figure 1. Because the unmanned plane has stronger nonlinear characteristic, if the control law is designed by adopting the traditional PID gain method, the design period is long, and the stability of the flight control system is difficult to ensure when the aircraft makes high-frequency intense maneuver. Among the various parameter variables of the unmanned aerial vehicle, the attitude variable belongs to the variable with slower change, including the attack angle, the roll angle, the accelerator and the like, and the angular velocity variable belongs to the variable with faster change, including the pitching, the rolling and the yaw angular velocity. In the nonlinear dynamic inverse control structure, a three-axis angular velocity instruction is obtained by resolving a gesture instruction through a control law slow variable loop, a control surface instruction is obtained by resolving an input fast variable loop, and the control surface instruction is obtained by resolving a control surface actuator link and is input into a six-degree-of-freedom nonlinear model of the unmanned aerial vehicle, so that a flight state at the next moment is obtained.
Numerous publications currently exist that use theoretical derivation methods to solve and controlThe matrix F and the matrix G are manufactured, the method is accurate but inconvenient to encode, the method is highly dependent on an aircraft mathematical model, the method does not have a general purpose type, and the numerical method adopts a difference method and a substitution solving method and can be suitable for various aircraft models. The present invention is directed to a nonlinear systemThe method for solving F and G is as follows:
(1) Let u=0, substituting into the system to obtain
(2) Let u=u respectively 0 +du and u=u 0 Du, du is small, available
Unmanned plane has various special effects, different flying maneuvers have different control requirements on gestures, tracks and movement angular speeds, and in order to complete the control effect of maneuver, the requirements on control laws at different stages of the same maneuver are also different
FIG. 2 is a schematic diagram of a control law configuration of a height-roll angle-speed configuration, with an input command being a target height variation ΔH c Target track deflection delta χ c Target speed variation Δv c 。
The climbing angle instruction can be obtained according to the altitude instruction:
γ c ＝K H ΔH c
wherein K is H Representing the gain factor between the climb angle and the height difference. The response characteristic of the track instruction is designed as a first-order dynamic characteristic, and is obtained by arrangement:
wherein K is V 、K γ And K χ The bandwidth of the three-axis track command loop is V and χ respectively represent the current speed and track deflection angle of the unmanned aerial vehicle. Thus obtaining a desired rate of change of the velocity vectorI.e.
According to the unmanned plane barycenter dynamics equation set, the thrust command T can be solved c Angle of attack command alpha c And a speed roll angle command mu c Wherein mu c Can be expressed as:
thrust command T c And angle of attack instruction alpha c The method can be obtained by iterative calculation of an unmanned aerial vehicle dynamic equation:
wherein M represents the unmanned aerial vehicle mass, L (alpha c ) Representing the angle of attack alpha c Unmanned aerial vehicle lift when.
Angle of attack instruction alpha obtained through flight path instruction calculation c And a speed roll angle command mu c Entering into nonlinear dynamic inverse control law to make sideslip angle command beta c =0, resulting in angle of attack, speed roll angle, and side slip angle rate of change:
k, K and K represent the system bandwidths of the response variables, respectively. And after the attitude angle change rate is obtained, a corresponding control surface instruction can be obtained through a nonlinear dynamic inverse slow variable loop and a fast variable loop.
This configuration is suitable for accurately controlling the maneuver of the attitude angle of the unmanned aerial vehicle.
FIG. 3 is a schematic diagram of a pitch rate-roll angle rate-throttle control law configuration with an input command of pitch rate command q c Roll angle speed command p c Throttle command Thr c ，Thr c The value range of (2) is [0,1 ]]。
Because the pitch angle speed, the roll angle speed and the throttle control law are directly input with the angular speed command, the external command directly controls the angular speed and the engine power of the unmanned aerial vehicle, and compared with the altitude-track deflection angular speed control law, the configuration structure is simpler, and a track resolving and NDI slow variable resolving module is not needed.
This configuration is suitable for directly controlling the unmanned aerial vehicle attitude and changing the maneuver of speed.
Fig. 4 is a schematic diagram of an angle of attack-roll angle speed-throttle control law configuration based on a pitch angle speed-roll angle speed-throttle control law configuration, changing the longitudinal input from a pitch angle speed command to an angle of attack command.
The input instruction of the attack angle, the rolling angle speed and the throttle control law configuration is an attack angle instruction alpha similar to the pitch angle speed, the rolling angle speed and the throttle configuration c Roll angle speed command p c Throttle command Thr c ，Thr c The value range of (2) is [0,1 ]]. The angle of attack instruction is solved through the NDI slow variable loop to obtain a corresponding pitch angle speed instruction input fast variable loop, and the roll angle speed instruction and the throttle instruction are directly input into the fast variable loop, so that a corresponding control surface instruction is obtained. The configuration is suitable for controlling the mechanical movement of the longitudinal attitude angle and the transverse attitude change speed of the unmanned plane at the same time.
The method for analyzing the special maneuver comprises the steps of segmenting a flight process according to the flight characteristics of the special maneuver of the unmanned aerial vehicle, selecting a proper control law configuration according to the control requirement of each stage, designing corresponding control instructions, and finally splicing the instructions of each segment to obtain the completed special maneuver process.
Fig. 5 is a schematic illustration of a barrel roll maneuver, in which the aircraft enters in a horizontal manner, a barrel roll maneuver is performed about a speed axis, i.e., one cycle movement is completed in both the longitudinal and transverse directions, and the speed roll angle is changed by 360 °.
The maneuver can be divided into three phases:
(1) Stage 1: the unmanned aerial vehicle enters maneuver from a fixed straight and flat flying state, and establishes an attack angle;
(2) Stage 2: the unmanned aerial vehicle rolls 360 degrees around the speed shaft while maintaining the attack angle;
(3) Stage 3: and after the speed roll angle of the unmanned aerial vehicle is changed by 360 degrees, unloading and recovering the initial flat flight state.
Fig. 6 is a flow chart of a barrel roll maneuver control design, for which the desired control law configurations are the angle of attack-roll angle speed-throttle configuration and altitude-track yaw-speed configuration. Fig. 7 is a simulated time domain response of a barrel roll maneuver obtained by simulation according to the method disclosed by the invention, and fig. 8 is a three-dimensional simulated track of the barrel roll maneuver.
Claims (2)
1. The design method of the hybrid configuration fixed wing unmanned aerial vehicle maneuvering simulation control law is characterized by comprising the following steps of:
step 1, according to characteristic parameters of a fixed-wing unmanned aerial vehicle, including mass, moment of inertia, product of inertia, chord length, span length, wing area, aerodynamic static derivative and dynamic derivative, carrying out six-degree-of-freedom modeling on the fixed-wing unmanned aerial vehicle according to aircraft flight dynamics, wherein the modeling process comprises a ground coordinate system, a machine body coordinate system, an airflow coordinate system, a track coordinate system and a conversion matrix among the coordinate systems;
step 2, a control law core module based on a nonlinear dynamic inverse method is established, a control matrix in any state is obtained through solving by a numerical method, namely, a control matrix F is solved by substituting an input value 0 method, a control matrix G is solved by a numerical differential method, the control law core comprises an inner ring and an outer ring, the outer ring obtains a triaxial angular velocity instruction according to attack angle, sideslip angle and velocity rolling angle instruction solution, and the inner ring obtains a control surface instruction according to the angular velocity instruction solution;
step 3, analyzing a flight profile of the required simulated maneuver, dividing maneuver phases according to kinematic features, extracting key parameters of each phase, determining a flight instruction required to be controlled, and making a maneuver control flow chart;
step 4, determining a required control law configuration according to the control instruction type and the maneuvering control flow chart obtained in the step 3, and then determining a specific numerical value of a flight instruction according to the flight characteristics of the required simulated maneuvering;
step 5, integrating the six-degree-of-freedom body motion model of the fixed-wing unmanned aerial vehicle established in the step 1, the unmanned aerial vehicle control law core established in the step 2 and the control law specific configuration established in the step 4, and completing simulation of the required unmanned aerial vehicle maneuver by combining a flight instruction;
the six-degree-of-freedom nonlinear mathematical model of the fixed-wing unmanned aerial vehicle described in the step 1 has the following basic equation:
wherein m represents the mass of the unmanned aerial vehicle, g represents the local gravity acceleration, alpha and beta respectively represent the attack angle and sideslip angle of the unmanned aerial vehicle, u, v and w respectively represent the triaxial speed components of the unmanned aerial vehicle under the body coordinate system, p, q and r respectively represent the pitching, rolling and yaw angle rates of the unmanned aerial vehicle under the body coordinate system, and phi, theta and psi respectively representIllustrating the roll angle, pitch angle and yaw angle, x of the unmanned aerial vehicle E 、y E 、z E Respectively representing projection positions of the unmanned aerial vehicle in three directions of a ground coordinate system; q 1 To q 4 Representing a quaternion for computing pose, c 1 To c 9 Representing a parameter related to the moment of inertia of the body,and->Respectively represents aerodynamic force and aerodynamic moment of the unmanned aerial vehicle in three directions under the machine body coordinate system, and for X th 、Y th 、Z th And M T 、N T 、L T Respectively representing engine thrust and engine thrust moment of the unmanned aerial vehicle in three directions under a machine body coordinate system;
the control law core in the step 2 has the following basic ideas:
for multiple-input multiple-output nonlinear systems
Order theThen
u＝G -1 (x)(-F(x)+K c (x c -x))
Wherein u represents a control matrix, x c Representing a target control quantity matrix, K c Representing a gain coefficient matrix;
the target control amount matrix used includes:
(1) Inner ring variable: pitch angle speed, roll angle speed, yaw angle speed;
(2) Outer loop variable: angle of attack, sideslip angle, speed roll angle;
the maneuvering phase described in step 3 includes:
(1) S turns continuously in the horizontal plane;
(2) S breaking the power;
(3) The barrel roller is powered;
the control law configuration described in step 4 includes:
(1) Altitude-track yaw-speed configuration;
(2) Pitch angle speed-roll angle speed-throttle configuration;
(3) Angle of attack-roll angle speed-throttle configuration;
the flight instruction of step 5 includes:
(1) A height instruction;
(2) Track deflection angle instructions;
(3) A speed command;
(4) A pitch angle rate command;
(5) A roll angle speed command;
(6) A throttle command;
(7) Angle of attack instruction.
2. The method for designing the hybrid configuration fixed wing unmanned aerial vehicle maneuver simulation control law according to claim 1, wherein the matrices F (x) and G (x) used are solved by numerical methods, i.e. for any x, let u=0, thenAnd G (x) is obtained by means of numerical differentiation.
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