CN112506046A - Stability augmentation control method for wingtip hinged combined type flight platform - Google Patents

Stability augmentation control method for wingtip hinged combined type flight platform Download PDF

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CN112506046A
CN112506046A CN202011075019.4A CN202011075019A CN112506046A CN 112506046 A CN112506046 A CN 112506046A CN 202011075019 A CN202011075019 A CN 202011075019A CN 112506046 A CN112506046 A CN 112506046A
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谢长川
安朝
孟杨
刘晨宇
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Beihang University
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    • G05B13/00—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion
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Abstract

The invention provides a stability augmentation control method for a wingtip hinged combined type flying platform, which cannot fly stably under an uncontrolled condition. After the flight mechanics model is established, an independent PID control loop is established for each sub aircraft, the control parameters are adjusted to realize the stable control of the whole flight platform, and the control loops among the sub aircraft are not coupled. Simulation analysis shows that the control strategy of the invention is practical and effective.

Description

Stability augmentation control method for wingtip hinged combined type flight platform
Technical Field
The invention provides a stability augmentation control method for a wingtip hinged combined type flight platform, and belongs to the field of flight mechanics simulation and control strategy design.
Background
Endurance is the core component of flight performance, and directly determines factors such as the activity range, the lasting flight capability and the economy of the airplane. The improvement of voyage and endurance is always the core and the aim of the research work of the endurance of the airplane. The combined flying platform formed by hinging and compositely flying a plurality of airplane wingtips is an effective way for improving the endurance performance of the airplane. Such as a straight layout conjoined aircraft in CN102658866A, a wingtip butt-joint parallel flying wing unmanned aerial vehicle system in CN103963972A, and the like. The uniqueness of the aircraft concept lies in that other auxiliary models and supporting facilities are not needed, the oil consumption during cruising can be effectively reduced, the cruising range and the cruising duration are improved, the structural weight of the aircraft is reduced, and meanwhile, the requirements on maneuverability and flexibility during task execution can be met.
The prior art shows that the wingtip hinged combined flying platform has different flying characteristics from the traditional aircraft, has unstable flying mode when being fixed and flat, and cannot keep stable. This feature affects the practical application of the relevant type of aircraft, and there is currently no control method for trim-state stable flight of wingtip articulated combined flying platforms.
Disclosure of Invention
The invention provides a control method for a wingtip hinged combined type flight platform, which realizes stable control of the wingtip hinged combined type flight platform.
According to one aspect of the invention, a wing tip hinged combined type flight platform balancing solution is given according to a flight dynamics model, and a system state space equation is obtained after linear processing is carried out on a flight mechanics equation near the balancing solution; and establishing an independent PID (proportion-integral-derivative) control loop for each sub aircraft in the combined type flight platform, and setting PID control parameters to realize stable flight of the whole flight platform.
The control method for the wingtip articulated combined flying platform according to one embodiment of the invention comprises the following steps:
step 1: and (4) computing initialization, including information such as design parameters, an analysis coordinate system and aerodynamic force derivatives of each aircraft in a given computational analysis model.
Step 2: establishing a model of flight mechanics comprising
Assuming each sub aircraft as a rigid body, establishing a force and moment balance equation, giving a hinge constraint to obtain a dynamic model, giving aerodynamic expression by using a proper aerodynamic model, and combining the dynamic model to obtain a flight mechanics model;
Figure BDA0002716395160000021
wherein q is a parameter of the motion state of the airplane, lambda is a constraint force variable, and F is a dynamic equilibrium equation established according to the flight mechanics model.
And step 3: calculating a balancing state, performing linearization treatment, including giving a flight working condition, solving the flight mechanical equation in the step 2, calculating the balancing state, performing linearization treatment on the flight mechanical equation under balancing, reasonably and effectively solving the stability problem under the system linearization state, and representing the state space equation of the flight system as follows:
Figure BDA0002716395160000022
wherein x is a state vector, u is an input vector, A is a system matrix, and B is an input matrix;
and 4, step 4: giving a PID control loop
And (3) establishing a PID control loop for each flight platform in the wingtip hinged combined type flight platform, wherein the PID loops are not coupled. The feedback quantity is set as the degree of freedom consistent with the degree of freedom of the hinge joint, and when the hinge joint releases the relative rolling degree of freedom of the aircraft, the feedback quantity is the rolling angle of each sub-aircraft; when the articulated connection releases the aircraft to attack the angle degree of freedom, the feedback quantity is the pitch angle of each sub aircraft.
The feedback loop forms a control signal number through the PID controller to operate the control surface to realize the control and the operation of the aircraft. The corresponding relationship between the feedback signal and the control surface is roll angle feedback-aileron control surface deflection; pitch angle feedback-elevator control surface deflection; yaw angle feedback-rudder control surface deflection.
And 5: setting PID parameter to realize stable control
In a PID stability augmentation loop in each sub aircraft, PID control parameters are respectively and independently given, pitch angle, yaw angle and roll angle instructions of each sub aircraft are given according to a trim state, dynamic response of the aircraft is obtained through simulation, and the control parameters are adjusted until the whole flying platform is stable, so that stable control is realized. And the control systems of the sub aircrafts are not coupled.
The invention provides a stable control method for a wingtip hinged combined type flight platform, which defines the control idea of independent loops of each sub aircraft; the beneficial effects of the invention include: (1) the PID control law is simple and easy to implement, and the engineering applicability is very strong; (2) each sub aircraft control loop is independent, control parameters are not coupled, and debugging and engineering realization are facilitated; (3) the method is convenient for realizing time domain simulation, can adopt a ready-made Matlab.
According to one aspect of the present invention, there is provided a method of controlling a wingtip-articulated combined flying platform comprising a plurality of sub-aircraft, characterized by comprising:
A) an initialization step, which comprises the steps of setting the design parameters of each sub-aircraft, analyzing a coordinate system and an aerodynamic derivative,
B) establishing a flight mechanics model, comprising:
each sub aircraft is regarded as a rigid body, the flight mechanics equation of the flight platform is determined,
C) calculating a trim state and performing linearization processing, wherein the method comprises the following steps:
under a given flight condition, solving the flight mechanics equation, and calculating a balancing result of each balancing degree of freedom, wherein the balancing degree of freedom comprises an attack angle of each of the plurality of sub aircrafts, an elevator deflection angle of each of the plurality of sub aircrafts, a roll angle of each of the plurality of sub aircrafts, and an aileron deflection angle of each of the plurality of sub aircrafts,
the dynamic linear processing is carried out under the balancing state,
D) establishing a PID control loop, comprising:
establishing respective PID control loops for the plurality of sub-aircraft, wherein:
there is no coupling between the PID control loops,
the plurality of sub aircrafts respectively input the same expected motion parameters, the actual motion parameters of the plurality of sub aircrafts are used as feedback quantities to form input quantities of respective PID control loops, respective control signals are formed after the PID control loops are resolved, the respective control signals are input into control surfaces of the corresponding sub aircrafts in the plurality of sub aircrafts to realize the stable control of the aircrafts,
E) setting PID parameters to realize stable control
And respectively and independently providing PID control parameters in the respective PID control loops of the plurality of sub aircrafts, and adjusting the sizes of the proportional control parameters, the integral control parameters and the differential control parameters so as to ensure that the flight platform is integrally stable and realize stable control.
Drawings
FIG. 1 is a schematic view of a wingtip articulated combined flying platform suitable for use with the present invention.
FIG. 2 is a schematic view of a sub-aircraft to which the present invention is applicable.
Fig. 3 is a schematic view of a trim state to which the present invention is applied.
FIG. 4 is a time domain response curve of roll angles of two sub-aircraft without the stability augmentation system.
FIG. 5 is a schematic diagram of an exemplary control scheme of the present invention.
FIG. 6 shows the time-domain response result of the roll angle after the stability augmentation control system is started by the sub-aircraft.
FIG. 7 shows the time-domain response result of the roll angular velocity after the sub-aircraft starts the stability augmentation control system.
Detailed Description
The stability augmentation control method for the wingtip articulated combined flying platform according to the invention is described below by taking a certain wingtip articulated combined flying platform as an example.
The wingtip hinged combined flying platform shown in fig. 1 is a combined example of two aircrafts, and each sub aircraft is regarded as a rigid body. Each sub-aircraft structure is shown in fig. 2. The sub-aircraft structure comprises 8 parts: the airplane comprises an airplane body 1, a right wing 2, a right wing aileron 3, a full-motion vertical tail 4 (which can be used as a rudder), a full-motion horizontal tail 5 (which can be used as an elevator), a left wing 6, a left wing aileron 7 and a middle wing 8. The two sub-aircrafts are connected by a one-way hinge and can freely rotate along the chordwise direction.
The stability augmentation control method for the wingtip articulated combined flying platform according to one embodiment of the invention comprises the following steps:
step 1: computing initialization, including:
given the relevant information for each sub-aircraft in the computational model: the design parameters are as follows:
design parameters Numerical value Design parameters Numerical value
Long wing span 3000mm Wing chord length 270mm
Vertical tail is long 240mm Vertical tail chord length 120mm
Flat tail length 960mm Length of flat tail chord 120mm
Distance of tail wing from gravity center 1020mm Quality of 8kg
Definition of OgxgygzgIs the geodetic coordinate system (inertial coordinate system), zgThe shaft is vertically downward directed to the center of the earth; definition of ObxbybzbFor a coordinate system of the body, x, fixedly connected to the sub-aircraftbThe axis being directed parallel to the aircraft axis in the plane of symmetry of the sub-aircraft toward the nose, zbPerpendicular to the plane of symmetry of the sub-aircraft and directed below the fuselage. The inertial coordinate system and the machine body coordinate system both meet the right-hand rule.
Let L be the lift, C be the lateral force, Mx, My, Mz be the aerodynamic moment of three directions respectively, p, q, r be the angular velocity of three directions in the organism coordinate system respectively, alpha be the angle of attack, beta be the sideslip angle, deltae,δa,δrRespectively an elevator, an aileron and a rudder deflection angle. The aerodynamic derivatives of the aircraft are as follows:
Figure BDA0002716395160000041
step 2: establishing a flight mechanics model, comprising:
and (3) assuming each sub aircraft as a rigid body, establishing a force and moment balance equation, and obtaining a dynamic model after a given articulation constraint. And (3) giving aerodynamic expression by using a proper aerodynamic model, and combining the aerodynamic model to obtain a flight mechanics model. In the embodiment, a dynamic equation of the sub-aircraft is established based on a Newton-Euler equation:
Figure BDA0002716395160000051
Figure BDA0002716395160000052
wherein r isi＝(xi,yi,zi)，xi,yi,ziIs 3 cartesian coordinates of the ith aircraft center of mass relative to the inertial frame,
Figure BDA0002716395160000053
θi,ψithe roll angle, the pitch angle and the yaw angle of the ith aircraft are respectively; m isiMass of the ith aircraft, Ji (i)Sub-aircraft inertia matrix, Fi (i)As vectors of external forces, Mi (i)For a moment vector (conforming to the general flight mechanics expression habit), the upper right corner (i) represents that the vector is expressed in the body coordinate system of the ith sub-aircraft, A(0,i)For transforming matrices to coordinate systems, DiIs a transformation matrix of angular coordinate derivative and body angular velocity vector.
The equations (3) and (4) can be written comprehensively as:
Figure BDA0002716395160000054
Figure BDA0002716395160000055
Figure BDA0002716395160000056
qi＝(xi,yi,zi,φi,θi,ψi) (8)
for a flight platform consisting of 2 aircraft, the system of equations is obtained:
Figure BDA0002716395160000057
A＝diag(A1,A2) (10)
Figure BDA0002716395160000058
Figure BDA0002716395160000059
the sub-aircrafts are connected through one-way hinges, and the constraint equation is as follows:
Figure BDA00027163951600000510
Figure BDA00027163951600000511
wherein, cij,cjjStarting from the center of mass to the hinge point OjBody hinge vector of, pjIs the direction vector of the hinge allowing rotation. Equations (5) (6) can be written collectively as:
Φ(q1,q2)＝0 (15)
introducing a Lagrange multiplier vector lambda to determine a constraint relation, and deriving a Lagrange equation of dynamics as follows:
Figure BDA00027163951600000512
wherein phiqTo constrain the Jacobian matrix of equations with respect to q, the flight mechanics equations can be obtained by integrating equations (9) (16):
Figure BDA00027163951600000513
Figure BDA0002716395160000061
the flight mechanics equation is consistent with the general form of equation (1).
And step 3: calculating the trim state, and performing linearization processing
Given flight conditions, the trim speed is 20m/s, and the atmospheric density is 1.225kg/m3And (3) requiring the flight platform to fly straight and flat, solving the flight mechanics equation in the step 1, and calculating the trim state as follows:
degree of freedom of balancing Trim result/rad
Attack angle of No. 1 machine 0.0821
Attack angle of No. 2 machine 0.0821
Elevator declination angle -0.0964
Rolling angle of No. 1 machine 0.0020
Rolling angle of No. 2 machine -0.0020
No. 1 aircraft aileron declination 0.0751
No. 2 aircraft aileron declination -0.0751
The schematic view of the trim state is shown in fig. 3, with machine No. 1 on the left and machine No. 2 on the right. The attack angle and the deflection angle of the elevator of the two sub-aircrafts are consistent, the rolling angles are the same, the directions are opposite, and the two sub-aircrafts are folded during trim flight.
And (4) performing dynamic linearization processing in the trim state, and establishing system equation simulation in Matlab. The time domain response curves of the roll angles of the two sub-aircrafts without the stability augmentation system are shown in figure 4. The time domain response of the roll angle signal is quickly diverged, which indicates that the flying platform can not stably fly without a control system.
And 4, step 4: establishing PID control loop
And establishing a PID (proportion-integral-derivative) control loop for each sub aircraft in the flight platform, wherein the PID loops are not coupled. The control concept is shown in fig. 5. And inputting the same expected motion parameters to each sub-aircraft, taking the actual motion parameters of each sub-aircraft as feedback quantities to form input quantities of each PID controller, and forming respective control signals to be input to each corresponding sub-aircraft after the input quantities are calculated by the PID controllers. Setting the roll angle of each sub aircraft for the example actual motion parameter, namely the feedback quantity; the feedback loop forms a control signal through the PID controller to operate the aileron control surface deflection so as to realize the stable control of the aircraft. During simulation, a corresponding control simulation loop can be established in Matlab
And 5: setting PID parameter to realize stable control
In a PID stability augmentation loop in each sub aircraft, PID control parameters are respectively and independently given, and the sizes of three types of control parameters, namely proportion (P), integral (I) and differential (D), are adjusted to enable the whole flying platform to be stable, so that stable control is realized. And the control systems of the sub aircrafts are not coupled.
After the sub-aircraft starts the stability augmentation control system, the time-domain response result of the roll angle is shown in fig. 6, and the time-domain response result of the roll angle speed is shown in fig. 7. The roll angle and roll angle speed signals of all the sub-aircrafts are converged to fixed values quickly, and the movement is stable. The flying platform can stably fly under the action of the control system.
Under the condition of a plurality of sub-aircrafts, the actual state of the roll angle is based on the actual trim calculation result, and the attack angle and the elevator are consistent. In particular, in the case of two aircraft, the roll angles are of the same magnitude and in opposite directions.
The invention provides a practical and effective control strategy scheme aiming at a wingtip hinged combined type flight platform comprising a plurality of sub-aircrafts.
The beneficial effects of the invention include: (1) the PID control law is simple and easy to implement, and the engineering applicability is very strong; (2) each sub aircraft control loop is independent, control parameters are not coupled, and debugging and engineering realization are facilitated; (3) the method is convenient for realizing time domain simulation, can adopt a ready-made Matlab.

Claims (4)

1. The control method of the wingtip hinged combined flying platform comprising a plurality of sub-aircrafts is characterized by comprising the following steps:
A) an initialization step, which comprises the steps of setting the design parameters of each sub-aircraft, analyzing a coordinate system and an aerodynamic derivative,
B) establishing a flight mechanics model, comprising:
each sub aircraft is regarded as a rigid body, the flight mechanics equation of the flight platform is determined,
C) calculating a trim state and performing linearization processing, wherein the method comprises the following steps:
under a given flight condition, solving the flight mechanics equation, and calculating a balancing result of each balancing degree of freedom, wherein the balancing degree of freedom comprises an attack angle of each of the plurality of sub aircrafts, an elevator deflection angle of each of the plurality of sub aircrafts, a roll angle of each of the plurality of sub aircrafts, and an aileron deflection angle of each of the plurality of sub aircrafts,
the dynamic linear processing is carried out under the balancing state,
D) establishing a PID control loop, comprising:
establishing respective PID control loops for the plurality of sub-aircraft, wherein:
there is no coupling between the PID control loops,
the plurality of sub aircrafts respectively input the same expected motion parameters, the actual motion parameters of the plurality of sub aircrafts are used as feedback quantities to form input quantities of respective PID control loops, respective control signals are formed after the PID control loops are resolved, the respective control signals are input to the control surfaces of the corresponding sub aircrafts in the plurality of sub aircrafts to realize the stable control of the aircrafts,
E) setting PID parameters to realize stable control
And respectively and independently providing PID control parameters in the respective PID control loops of the plurality of sub aircrafts, and adjusting the sizes of the proportional control parameters, the integral control parameters and the differential control parameters so as to ensure that the flight platform is integrally stable and realize stable control.
2. The control method according to claim 1, characterized in that:
further comprising establishing a simulation of system equations in matlab
The control simulation loop comprises a time domain response simulation part for judging and evaluating the stable control performance.
3. The control method according to claim 1, characterized in that:
in the step D), the actual motion parameter, i.e. the feedback quantity, is set to the respective roll angles of the plurality of sub-aircraft, and the feedback loop forms a control signal for operating the deflection of the aileron control surface via a PID control loop.
4. The control method according to claim 1, characterized in that:
the feedback quantity is set as the degree of freedom consistent with the hinge degree of freedom, and when the hinge connection releases the relative rolling degree of freedom of the sub aircrafts, the feedback quantity is the rolling angle of each sub aircraft; when the attack angle freedom degree of the sub aircraft is opened through hinged connection, the feedback quantity is the pitch angle of each sub aircraft,
the feedback loop forms a control signal through the PID controller, the control signal operates the control surface to realize the control and the control of the sub-aircraft,
the corresponding relationship between the feedback signal and the control plane is as follows: roll angle feedback-aileron control surface deflection, pitch angle feedback-elevator control surface deflection, yaw angle feedback-rudder control surface deflection.
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