CN105334735A - Flying wing layout unmanned aerial vehicle control law based on angular rate - Google Patents

Abstract

The invention discloses a flying wing layout unmanned aerial vehicle control law based on an angular rate. By adopting a robust servo linear quadratic regulator (being called as LQR for short) and root locus method, the control laws of longitudinal and transverse courses are designed respectively. Compared with a conventional control law, the longitudinal control law taking a pitching angular rate and a height change rate as control targets is added in a pitching angle and height control; and system damping is enhanced and the longitudinal stability is improved. Meanwhile, the transverse control law taking a rolling angular rate as a control target is added in rolling angular control, and the control problems that the rolling manipulation efficiency of an unmanned aerial vehicle is too high and the low stability is caused by small rotary inertia can be effectively overcome. According to the method provided by the invention, the accurate tracking on appointed height and track can be realized, and the control accuracy, robustness and reliability can meet the design requirements on gesture and track control by the flying wing layout unmanned aerial vehicle; and the method is applied and verified in the multi-manipulation-face flying wing layout unmanned aerial vehicle.

Description

Flying wing layout unmanned aerial vehicle control law based on angular rate

The technical field is as follows:

the invention relates to a flying wing layout unmanned aerial vehicle control law based on angular rate, in particular to an angular rate control law and a route tracking guidance law.

Background art:

the unmanned aerial vehicle with the flying wing layout has important application value in aspects of missile interception, penetration attack and the like, the small-aspect-ratio and large-sweepback pneumatic layout improves the flight speed and response capability, meanwhile, the variation range of the flight speed and the height is expanded, and higher requirements are provided for a flight control law. As the flying height increases, the air becomes thinner and thinner, resulting in a smaller and smaller damping moment of the aircraft itself. Therefore, the damping of the self angular motion of the airplane is reduced, and the nose swings, so that the airplane is difficult to finish the tasks of aiming, shooting and the like. Therefore, a flight control system must be added to meet the performance of large envelope and high maneuverability. In the past, the control law with attitude angle proportional-integral control as an inner loop is difficult to overcome the control problems caused by high roll control efficiency and poor stability of the flying wing layout unmanned aerial vehicle.

The invention content is as follows:

in order to realize the stable control of the flying wing layout unmanned aerial vehicle, the invention provides a longitudinal and transverse control law and a guidance law based on angular rate control, and an attitude angle control instruction and a task route can be smoothly, stably and quickly tracked.

The technical scheme of the invention is as follows: a flying wing layout unmanned aerial vehicle control law based on angular rate is characterized in that longitudinal and transverse course control laws are respectively designed, and a longitudinal control law with the pitch angle rate and the altitude change rate as control targets is added in pitch angle and altitude control; the roll angle control is added with a course control law taking the roll angle rate as a control target. The method comprises the steps of designing a pitch angle rate control law and a roll angle rate control law by using a robust servo LQR theory, introducing an angular rate deviation amount into a system, and regulating the deviation amount to be zero to enable a system state variable to track a control instruction; designing an attitude angle control law by combining a root track design method according to the bandwidth matching relationship between the inner loop and the outer loop of the loop, taking an attitude angle deviation signal as an input signal of an angular rate control loop, giving an attitude angle control instruction by an outer loop guidance law, and realizing stable tracking of a task route; in the guidance loop, a high-degree change rate control loop is added to increase the long-period damping of the unmanned aerial vehicle, and a track control loop is added to enhance the lateral-course damping.

The invention takes the angular rate as a control target, combines the robust servo LQR control theory, can enhance the damping of an attitude control loop, relieves the impulse effect of control surface output on a system, reduces the overshoot of the system, and achieves the aim of stably controlling the flying wing layout unmanned aerial vehicle to fly.

The beneficial effects obtained by the invention are as follows: the overshoot of the system response is reduced, the unmanned aerial vehicle can quickly track and reserve the task route, and the robustness of the system is enhanced.

Description of the drawings:

FIG. 1 is a flying wing layout drone control law structure diagram;

FIG. 2 is a diagram of a pitch rate control law structure;

FIG. 3 is a view of a pitch control law structure;

FIG. 4 is a diagram of a height control law structure;

FIG. 5 is a diagram of a roll angle control law structure;

FIG. 6 is a side-offset control law structure diagram;

FIG. 7 is a diagram of a track angular rate control law.

The specific implementation method comprises the following steps:

the technical scheme of the invention is explained in detail in the following with the accompanying drawings.

The parameters herein are defined as follows:

fig. 1 is a diagram of a longitudinal control law structure of the unmanned aerial vehicle of the present invention, including a pitch angle rate control law loop, a pitch angle control loop, and an altitude control loop, as shown in fig. 2, 3, and 4.

The angular rate control law is designed by adopting a robust servo LQR control theory, the angular rate deviation is introduced into the system, the deviation is adjusted to be zero, the system state variable tracks a control instruction, and rapidity and robustness are ensured; the influence of external interference on the system can be better inhibited by putting the integral link into the angular rate control loop.

The pitch angle rate control law is as shown in formula 1, a feedforward control link is added on the basis of a proportional-integral control structure, the response speed of the system can be accelerated, and after the peak time, the response attenuation is small and the change is smooth, so that the smoothness of the control response of the subsequent pitch angle rate control law is ensured.

${\mathrm{\δ}}_e=K_e^{IQ}\∫(Q-Q_g)dt+K_e^{Q}Q$ (1) Formula (II)

The control structure is shown in figure 2, for the pitch angle rate deviation signal (Q-Q)_g) Proportional-integral and amplitude limiting processing, proportional processing to the elevation rate signal (Q), adding the two processing results, and settling in real time to obtain the elevator control signal ((_e)。

Selecting a state variable x₁Q, output variable y_eAdditionally adding a state variable x₂＝Q-Q_gThereby obtaining a state coefficient matrix A and a control coefficient matrix B of the new system. Resolving the control gain K by the Riccati equation_C＝[K_IK_x]First, control weights are definedMatrix arrayPerformance weighting matrixThen selecting suitable one through iteration of system loopAfter the formation, finally solving the Riccati equation to obtain the control gain. Wherein,the larger the value the faster the system response speed, where a₁The system steady-state error is related, the larger the value of the error is, the smaller the steady-state error is, but the larger the value of the error is, the system can oscillate; a is₂Has the function of increasing the damping of the system.

The pitch angle control law is as shown in formula 2, on the basis of the pitch angle rate control circuit, the angular rate control law is used as an inner circuit, the attitude angle deviation signal is used as an input signal of the angular rate control circuit, an attitude angle control instruction is given by an outer circuit guidance law, and the control surface output of the attitude angle control law is smoother while the attitude angle control law quickly tracks a given value.

$Q_g=K_e^{\mathrm{\Θ}}({\mathrm{\Θ}}_g-\mathrm{\Θ})$ (2) Formula (II)

The control structure is shown in figure 3, and the pitch angle deviation signal (theta) is obtained_g-theta) is subjected to scaling and amplitude limiting processing, and a pitch angle rate given signal (Q) is obtained through real-time settlement_g)。

The altitude control law is as shown in formula 3, the pitch angle control circuit is used as an inner circuit of the altitude control circuit, and the pitch attitude is changed through the altitude deviation signal, so that the track inclination angle is changed to realize closed-loop stabilization and control of the flying altitude. Meanwhile, in order to further enhance the damping characteristic of the system, a height change rate control loop is added, and the effects of smooth connection of height control and attitude control are achieved, as shown in formula 3.

${\mathrm{\Θ}}_g=K_e^{\stackrel{\·}{H}}({\stackrel{\·}{H}}_g-\stackrel{\·}{H})+K_e^{I\stackrel{\·}{H}}\∫({\stackrel{\·}{H}}_g-\stackrel{\·}{H})dt$ (3) Formula (II)

${\stackrel{\·}{H}}_g=K_e^H(H_g-H)$

The structure is shown in FIG. 4 for the height deviation signal (H)_g-H) scaling and amplitude limiting to obtain high degree of changeGiven a signal, and then for a height rate of change deviation signalProportional and integral amplitude limiting processing is carried out, and a pitch angle rate given signal (theta) is obtained through real-time settlement_g) As an input signal to the pitch angle control loop.

The unmanned aerial vehicle lateral course control law comprises a roll angle control loop and a lateral offset control loop, as shown in fig. 5 and 6.

The roll angle control law is shown as formula 4, and similar to the longitudinal control strategy, the roll angle control law takes a roll angle rate control loop as an inner loop to enhance the smoothness and damping characteristics of the control effect and reduce the possibility of oscillation of the roll angle response, as shown in formula 4.

${\mathrm{\δ}}_a=K_a^{IP}\∫(P-P_g)dt+K_a^{P}P$ (4) Formula (II)

$P_g=K_a^{\mathrm{\Φ}}({\mathrm{\Φ}}_g-\mathrm{\Φ})$

The control structure is shown in FIG. 5 for the roll angle deviation signal (phi)_gProportional amplitude limiting processing is carried out on phi), and a roll angular rate given signal (P) is obtained through real-time settlement_g) Then to the roll rate deviation signal (P)_g-P) is subjected to proportional-integral and amplitude limiting treatment, the roll angle rate signal (P) is subjected to proportional treatment, the two treatment results are added, and real-time settlement is carried out to obtain the control signal (phi) of the aileron rudder_g-Φ)。

Assuming that the unmanned plane tracks the course flight to perform uniform circular motion, the following can be obtained approximately:

$V\stackrel{\·}{\mathrm{\Ψ}}=\stackrel{\·\·}{Y}\⇒V\mathrm{\Ψ}=\stackrel{\·}{Y}\≈K(Y_g-Y)\⇒{\mathrm{\Ψ}}_g=K_a^Y(Y_g-Y)$ (5) formula (II)

On the basis of a roll angle control loop, a track angle proportional-integral control law is designed by using a method of regulating parameters by a root track, the track angle deviation signal is limited in the control law resolving process, so that the roll angle instruction can be prevented from being changed violently, the control authority of the integral signal can be reduced by limiting the amplitude of the integral signal, and the system is prevented from generating large overshoot.

Yaw control law is as shown in equation 6, and yaw angle control command is obtained from yaw deviation signal, and appropriate control parameter is selected

Number ofAnd the lateral offset deviation signal is limited, so that the lateral offset angle control instruction can be ensured to be within the working range of the sensor.

${\mathrm{\Φ}}_g=K_a^{\mathrm{\Ψ}}({\mathrm{\Ψ}}_g-\mathrm{\Ψ})+K_a^{I\mathrm{\Ψ}}\∫({\mathrm{\Ψ}}_g-\mathrm{\Ψ})dt$ (6) Formula (II)

${\mathrm{\Ψ}}_g=K_a^Y(Y_g-Y)$

The control structure is shown in fig. 7, the side offset deviation signal is subjected to amplitude limiting and proportional processing, real-time settlement is carried out to obtain a side offset angle given signal, then proportional-integral and amplitude limiting processing is carried out on the side offset angle deviation signal, real-time settlement is carried out to obtain a roll angle given signal, and the roll angle given signal is used as an input signal of a roll angle controller.

As shown in equation 7, the rudder control signal is obtained by filtering, scaling and limiting the yaw rate.

${\mathrm{\Ψ}}_g=K_a^Y(Y_g-Y)$ (7) Formula (II) is shown.

Claims (4)

1. The utility model provides a flying wing overall arrangement unmanned aerial vehicle control law based on angular rate which characterized in that: the control law comprises a longitudinal control law and a transverse course control law; in a longitudinal control law, the pitch angle rate and the altitude change rate are increased in pitch angle and altitude control as control targets; in the course control law, the roll angle rate is increased as a control target in roll angle control.

2. The flying wing layout drone law according to claim 1, wherein: the longitudinal control law and the transverse course control law take angular rate control as a main control loop;

the longitudinal control law, that is, the longitudinal pitch angle rate (Q) of the unmanned aerial vehicle with the flying wing layout is taken as a longitudinal main control loop, and the pitch angle rate (Q) is given_g) By longitudinal control law

${\mathrm{\δ}}_e=K_e^{\mathrm{IQ}}\∫(Q-Q_g)\mathrm{dt}+K_e^{Q}Q---\left(1\right)$ Formula (II)

Real-time calculation of elevator control signals (_e) To the elevator;

the lateral course control law is that the flying wing layout unmanned aerial vehicle takes the roll angle rate (P) as a main control loop in the lateral direction and gives the roll angle rate (P)_g) By roll angle control law

${\mathrm{\δ}}_a=K_a^{\mathrm{IP}}\∫(P-P_g)\mathrm{dt}+K_a^{P}P---\left(2\right)$ Formula (II)

Real-time solution of aileron control signals (_a) To aileron executive machineAnd (5) forming.

3. The flying wing layout drone law according to claim 1, wherein: designing an attitude control loop by taking angular rate control as an inner loop;

longitudinal passing pitch angle control law of flying wing layout unmanned aerial vehicle

$Q_g=K_e^{\mathrm{\Θ}}({\mathrm{\Θ}}_g-\mathrm{\Θ})---\left(3\right)$ Formula (II)

Solving a given value of the pitch angle rate in real time;

law of transverse course of flying wing layout unmanned aerial vehicle through roll angle control law

$P_g=K_a^{\mathrm{\Φ}}({\mathrm{\Φ}}_g-\mathrm{\Φ})---\left(4\right)$ Formula (II)

And calculating the given value of the roll angular rate in real time.

4. The flying wing layout drone law according to claim 1, wherein: designing an airway guidance loop by taking the attitude control loop as an inner loop;

flying wing layout unmanned aerial vehicle longitudinal proportional-integral control with high change rate as outer loop

${\mathrm{\Θ}}_g=K_e^{\stackrel{\·}{H}}({\stackrel{\·}{H}}_g-\stackrel{\·}{H})+K_e^{I\stackrel{\·}{H}}$

${\stackrel{\·}{H}}_g=K_e^H(H_g-H)---\left(5\right)$ Formula (II)

Settling in real time according to the formula (5) by the height deviation signal to obtain an input given signal of the pitch angle control loop;

transverse course of flying wing layout unmanned aerial vehicle takes track proportional-integral control as outer loop

${\mathrm{\Φ}}_g=K_a^{\mathrm{\Ψ}}({\mathrm{\Ψ}}_g-\mathrm{\Ψ})+K_a^{I\mathrm{\Ψ}}\∫({\mathrm{\Ψ}}_g-\mathrm{\Ψ})dt$

${\mathrm{\Ψ}}_g=K_a^Y(Y_g-Y)---\left(6\right)$ Formula (II)

And (4) calculating in real time from the lateral offset deviation signal according to the formula (6) to obtain an input given signal of the roll angle control loop.