CN112433473B - Robust decoupling control system and control method considering coupling problem of rotary aircraft - Google Patents
Robust decoupling control system and control method considering coupling problem of rotary aircraft Download PDFInfo
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Abstract
The invention discloses a robust decoupling control system and a robust decoupling control method considering the coupling problem of a rotary aircraft, wherein the system comprises a required overload receiving module for receiving required overload information transmitted by a guidance system in real time, an aircraft parameter measuring module for obtaining flight parameters of the aircraft in real time, and a robust decoupling control module for obtaining a steering engine response instruction, wherein an attack angle convergence error and a sideslip angle convergence error are obtained in real time through a convergence error resolving submodule; obtaining an attack angle convergence sliding mode surface and a side sliding angle convergence sliding mode surface in real time through a convergence sliding mode surface resolving submodule; and obtaining a pitching direction steering engine response instruction and a yawing direction steering engine response instruction through a steering engine response instruction resolving submodule.
Description
Technical Field
The invention relates to a control system and a control method of a rotary aircraft, in particular to a robust decoupling control system and a robust decoupling control method considering the coupling problem of the rotary aircraft.
Background
The rotary aircraft can bring a plurality of benefits by adopting a spinning mode, such as effectively reducing the influence of the structural design deviation of the aircraft on the trajectory of the aircraft, simplifying the design of a control system, omitting a rolling control mechanism and the like. However, the aircraft has a plurality of advantages and disadvantages. Because the aircraft can generate a larger rolling angular velocity after autorotation, the aircraft generates characteristics such as pneumatic coupling, inertial coupling, control coupling and the like, the pitching channel and the yawing channel are mutually coupled and crosslinked, and the accurate control on the pitching channel and the yawing channel of the aircraft is not facilitated. In addition, the rotating aircraft is subjected to the gyroscopic effect and magnus moment caused by its own rotation during flight. These problems present a significant challenge to the precise and stable control of a rotary aircraft control system;
in the prior art, in the process of actually controlling the rotary aircraft, the influence of the coupling is ignored, some rotary aircraft adopt a traditional decoupling algorithm based on a linear control theory, the algorithm is often poor in robustness and is easily influenced by external interference and internal noise, the accuracy requirement of the algorithm on a system model is extremely high, once disturbance change occurs, the operation of an aircraft control system is often disordered, and the aircraft cannot normally work due to instability;
for most of the rotating aircrafts in practical application, because the conventional autopilot does not consider the coupling problem, when a single channel of the aircraft is controlled, the response of another channel is often caused, so that the control system is disordered, the attitude of the aircraft is unstable, and unpredictable results are caused.
In order to solve the above problems, it is necessary to design a decoupling control system and method considering the coupling problem of the rotating aircraft, so as to ensure the accurate control of the pitch and yaw channels of the aircraft.
Disclosure of Invention
In order to overcome the problems, the inventor of the invention carries out intensive research and designs a robust decoupling control system and a robust decoupling control method considering the coupling problem of a rotary aircraft, wherein the system comprises an overload receiving module which is used for receiving overload information which is required and transmitted by a guidance system in real time, an aircraft parameter measuring module which is used for obtaining flight parameters of the aircraft in real time, and a robust decoupling control module which is used for obtaining a steering engine response instruction, wherein an attack angle convergence error and a sideslip angle convergence error are obtained in real time through a convergence error resolving submodule; obtaining an attack angle convergence sliding mode surface and a side sliding angle convergence sliding mode surface in real time through a convergence sliding mode surface resolving submodule; and obtaining a pitching direction steering engine response instruction and a yawing direction steering engine response instruction through a steering engine response instruction resolving submodule, thereby completing the invention.
In particular, it is an object of the invention to provide a robust decoupled control system taking into account the coupling problem of a rotary aircraft on which the system is mounted, the system comprising
The overload receiving module 1 is connected with a guidance system on the rotary aircraft and used for receiving the overload information which is transmitted by the guidance system in real time,
an aircraft parameter measurement module 2 for obtaining flight parameters of the aircraft in real time, and
and the robust decoupling control module 3 is used for obtaining a steering engine response instruction in real time according to the overload information required and the flight parameters of the aircraft.
Wherein the overload demand information includes an expected pitch overload and an expected yaw overload
The steering engine response command comprises a pitching direction steering engine response command and a yawing direction steering engine response command.
Wherein the aircraft parameter measurement module 2 comprises an accelerometer 21, an inertial gyro 22, an estimator 23 and an integrator 24;
wherein, the accelerometer 21 is used for measuring acceleration information of the aircraft in real time,
the inertial gyroscope 22 is used for measuring yaw rate information and pitch rate information of the aircraft in real time,
the estimator 23 is configured to estimate in real time an attack angle and a sideslip angle of the aircraft according to the triaxial acceleration information;
the integrator 24 is configured to perform real-time integration according to the triaxial acceleration information to obtain speed information of the aircraft, perform real-time integration according to the yaw rate information to obtain yaw angle information, and perform real-time integration according to the pitch angle rate information to obtain pitch angle information.
The robust decoupling control module 3 comprises a convergence error resolving submodule 31, a convergence sliding mode surface resolving submodule 32 and a steering engine response instruction resolving submodule 33;
the convergence error resolving submodule 31 is configured to obtain an attack angle convergence error and a sideslip angle convergence error in real time according to the required overload information and flight parameters of the aircraft;
the convergence sliding mode surface resolving submodule 32 is used for obtaining an attack angle convergence sliding mode surface according to flight parameters and an attack angle convergence error of the aircraft in real time and obtaining a sideslip angle convergence sliding mode surface according to the flight parameters and the sideslip angle convergence error of the aircraft in real time;
the steering engine response instruction resolving submodule 33 is used for obtaining a pitch direction steering engine response instruction in real time according to the overload information required, the flight parameters of the aircraft and the attack angle convergence sliding mode surface, and obtaining a yaw direction steering engine response instruction in real time according to the overload information required, the flight parameters of the aircraft and the side slip angle convergence sliding mode surface.
Wherein, the convergence error resolving submodule 31 obtains the convergence error of the attack angle and the convergence error of the sideslip angle in real time through the following formula (one),
wherein e is1Representing angle of attack convergence error, e2Denotes the sideslip angle convergence error, alpha denotes the angle of attack, beta denotes the sideslip angle, aycIndicating an expected pitch overload, azcIndicating an expected yaw overload, V indicating the speed of the aircraft, a34Representing the power coefficient of the rotating aircraft.
Wherein, the convergence sliding mode surface resolving submodule 32 obtains an attack angle convergence sliding mode surface and a side sliding angle convergence sliding mode surface in real time through the following formula (II),
wherein s is1Representing the convergent sliding-form surface of angle of attack, s2Representing the sideslip angle converging sliding mode face, c and d both representing gain factors, theta representing the pitch angle, psi representing the yaw angle.
7. The robust decoupling control system in consideration of rotating aerial vehicle coupling issues of claim 4,
the steering engine response instruction resolving submodule (33) obtains a pitching direction steering engine response instruction and a yawing direction steering engine response instruction in real time through the following formulas (three) and (four),
wherein, deltayIndicating steering engine response command in pitch direction, deltazIndicating the yaw direction steering engine response command,
the derivative of the representation of the derivative of xi,
is representative of xi2The derivative of (a) of (b),
a25、a24、a27、a22and a28Both represent the power coefficient of the rotating aircraft.
The invention also provides a robust decoupling control method considering the coupling problem of the rotary aircraft,
the method comprises the following steps of,
step 1, receiving overload information required to be used transmitted by a guidance system through an overload receiving module 1;
step 3, a steering engine response instruction is obtained through the robust decoupling control module 3 according to the overload information required and the flight parameters of the aircraft;
and 4, repeating the steps 1-3 in real time, so as to obtain the steering engine response instruction in real time.
Wherein the step 2 comprises the following sub-steps,
substep 2-1, obtaining acceleration information of the aircraft through real-time measurement of the accelerometer 21, obtaining yaw rate information and pitch rate information of the aircraft through real-time measurement of the inertial gyroscope 22,
in the substep 2-2, the attack angle and the sideslip angle of the aircraft are obtained by estimating in real time according to the triaxial acceleration information through the estimator 23;
and obtaining the speed information of the aircraft by integrating the triaxial acceleration information in real time through an integrator 24, obtaining the yaw angle information by integrating the yaw angle rate information in real time, and obtaining the pitch angle information by integrating the pitch angle rate information in real time.
Wherein the step 3 comprises the sub-steps of,
in the substep 3-1, the convergence error of the attack angle and the convergence error of the sideslip angle are obtained in real time according to the overload information required and the flight parameters of the aircraft through the convergence error calculation submodule 31;
in the substep 3-2, an attack angle convergence sliding mode surface is obtained in real time according to the flight parameters and the attack angle convergence error of the aircraft through a convergence sliding mode surface resolving submodule 32, and a sideslip angle convergence sliding mode surface is obtained in real time according to the flight parameters and the sideslip angle convergence error of the aircraft;
and in the substep 3-3, acquiring a pitch direction steering engine response instruction in real time according to the overload information required, the flight parameters of the aircraft and the attack angle convergence sliding mode surface through the steering engine response instruction resolving submodule 33, and acquiring a yaw direction steering engine response instruction in real time according to the overload information required, the flight parameters of the aircraft and the side slip angle convergence sliding mode surface.
The invention has the advantages that:
(1) the robust decoupling control system and the robust decoupling control method considering the coupling problem of the rotary aircraft adopt the design concept of sliding mode control, so that the overload response on the rotary aircraft can quickly and accurately track the overload instruction required to be used, and the control on a pitching channel cannot influence the overload response of a yawing channel;
(2) the robust decoupling control system and the robust decoupling control method considering the coupling problem of the rotary aircraft can provide more reasonable steering engine control instructions for the steering engine by combining the current flight condition of the rotary aircraft and considering the coupling condition on the basis of the received overload required, thereby enhancing the control effect of the rotary aircraft and improving the control precision of the rotary aircraft.
Drawings
FIG. 1 illustrates an overall logic diagram of a robust decoupled control system that takes into account the problem of rotating aircraft coupling, in accordance with a preferred embodiment of the present invention;
FIG. 2 shows the overload and response curve required in the pitch direction in a simulation experiment;
fig. 3 shows the required overload and response curve for the yaw direction in the simulation experiment.
The reference numbers illustrate:
1-overload receiving module
2-aircraft parameter measuring module
3-robust decoupling control module
21-accelerometer
22-inertia gyroscope
23-estimator
24-integrator
31-convergence error resolving submodule
32-convergence sliding mode surface resolving submodule
33-steering engine response instruction resolving submodule
Detailed Description
The invention is explained in more detail below with reference to the figures and examples. The features and advantages of the present invention will become more apparent from the description.
The word "exemplary" is used exclusively herein to mean "serving as an example, embodiment, or illustration. Any embodiment described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other embodiments. While the various aspects of the embodiments are presented in drawings, the drawings are not necessarily drawn to scale unless specifically indicated.
According to the robust decoupling control system considering the coupling problem of the rotary aircraft, as shown in FIG. 1, the system is installed on the rotary aircraft, and the rotary aircraft is preferably a high-dynamic rotary aircraft, namely a rotary aircraft with the rotating speed of more than 10 r/s; the coupling means that when the pitch direction and the yaw direction of the aircraft are controlled respectively, a control command in one direction influences and interferes with the other direction, and particularly when the pitch direction of the aircraft is controlled, due to rotation, acting force generated by a steering engine has a certain component force in the horizontal direction, and the component force can cause the aircraft to deflect in the yaw direction.
The system comprises a required overload receiving module 1, an aircraft parameter measuring module 2 and a robust decoupling control module 3;
wherein, the overload receiving module 1 is connected with a guidance system on the rotary aircraft and used for receiving the overload information which is required and transmitted by the guidance system in real time,
the guidance system is also installed on the aircraft, and can give the overload that needs to be used in real time according to aircraft self information and target information that sensing device obtained on the aircraft, and the steering engine is controlled according to this needs again to beat the rudder work under general condition, but in the scheme that this application provided, should need to use the overload and can not directly transmit for the steering engine, but transmit for need to use overload receiving module 1, transmit the steering engine response command that obtains processing for the steering engine after handling. Therefore, the steering work of the steering engine is more targeted, and the control effect on the rolling aircraft is better.
The guidance system is a guidance system existing in the field, and an existing guidance law such as a proportional guidance law, a gravity compensation guidance law and the like can be adopted.
The aircraft parameter measuring module 2 is used for obtaining flight parameters of an aircraft in real time, wherein the flight parameters comprise acceleration, speed, yaw rate, pitch rate, attack angle, sideslip angle, yaw angle and pitch angle; the aircraft parameter measuring module 2 can call the power coefficient related to the flight parameter in real time from a memory chip carried by the aircraft parameter measuring module.
The robust decoupling control module 3 is used for obtaining a steering engine response instruction according to overload information required and flight parameters of an aircraft in real time, transmitting the steering engine response instruction to a steering engine, and enabling the steering engine to steer according to the instruction.
In a preferred embodiment, the overload need information includes an expected pitch overload and an expected yaw overload; the expected pitch overload is the overload which is solved by the guidance system and needs to be provided in the pitch direction; the expected yaw overload is the overload that the guidance system solves for, and needs to provide in the yaw direction.
The steering engine response instruction comprises a pitching steering engine response instruction and a yawing steering engine response instruction; the pitch direction steering engine response instruction represents an instruction which is finally transmitted to the steering engine and is executed by the steering engine in the pitch direction; and the yaw direction steering engine response command represents a command which is finally transmitted to the steering engine and is executed by the steering engine in the yaw direction.
When a guidance system of an aircraft obtains overload needing to be used, if the overload needing to be used is directly transmitted to a steering engine, the steering engine inevitably solves a corresponding response instruction of the steering engine in a pitching direction according to expected pitching overload in the overload needing to be used, and solves a corresponding response instruction of the steering engine in a yawing direction according to expected yawing overload in the overload needing to be used, and the action effects of the two response instructions are inevitably coupled in the execution process of the steering engine, so that the deviation between the final steering result and the expected value is overlarge; after the overload is resolved by the robust decoupling control module, the influence caused by interference factors such as coupling and the like is considered in advance, so that the steering operation is carried out according to the finally obtained steering engine response instruction, the steering result is closer to the expected value, and the pitching coupling interference on at least the yaw direction is smaller.
In a preferred embodiment, the aircraft parameter measurement module 2 comprises an accelerometer 21, an inertial gyro 22, an estimator 23 and an integrator 24;
wherein, the accelerometer 21 is used for measuring acceleration information of the aircraft in real time,
the inertial gyroscope 22 is used for measuring yaw rate information and pitch rate information of the aircraft in real time,
the estimator 23 is configured to estimate in real time an attack angle and a sideslip angle of the aircraft according to the triaxial acceleration information;
the integrator 24 is configured to perform real-time integration according to the triaxial acceleration information to obtain speed information of the aircraft, perform real-time integration according to the yaw rate information to obtain yaw angle information, and perform real-time integration according to the pitch angle rate information to obtain pitch angle information.
Wherein, said accelerometer 21 is provided with a plurality of, preferably at least 3, at least one of them is located on the center of mass of the aircraft, and it is installed towards the aircraft traveling direction along the aircraft axis, so as to measure the acceleration of the aircraft along the axis direction, i.e. the acceleration of the aircraft itself, and the acceleration can be integrated to obtain the speed information of the aircraft;
in addition, two accelerometers are arranged on the axis of the aircraft and deviate from the center of mass by a certain distance, the installation directions of the two accelerometers are perpendicular to each other, the two accelerometers are connected with the estimator 23, the accelerometers can measure the acceleration value of the position where the accelerometers are located in real time, the velocity of the point can be obtained after integration, the angular velocity of the point can be obtained by multiplying the distance between the point and the center of mass, and the angle can be obtained by integration; preferably, the estimator is further connected with a geomagnetic sensor on the aircraft, and the geomagnetic sensor can acquire the roll angle of the aircraft in real time, so that the sideslip angle and the attack angle of the aircraft can be obtained through the two accelerometers and the roll angle information respectively. The distance between the accelerometer and the centroid is stored in the estimator, and integral calculation can be carried out in the estimator, so that the estimator can give sideslip angle information and attack angle information of the aircraft in real time.
In a preferred embodiment, the robust decoupling control module 3 includes a convergence error resolving submodule 31, a convergence sliding mode surface resolving submodule 32 and a steering engine response instruction resolving submodule 33;
the convergence error resolving submodule 31 is configured to obtain an attack angle convergence error and a sideslip angle convergence error in real time according to the required overload information and flight parameters of the aircraft;
the convergence sliding mode surface resolving submodule 32 is used for obtaining an attack angle convergence sliding mode surface according to flight parameters and an attack angle convergence error of the aircraft in real time and obtaining a sideslip angle convergence sliding mode surface according to the flight parameters and the sideslip angle convergence error of the aircraft in real time;
the steering engine response instruction resolving submodule 33 is used for obtaining a pitch direction steering engine response instruction in real time according to the overload information required, the flight parameters of the aircraft and the attack angle convergence sliding mode surface, and obtaining a yaw direction steering engine response instruction in real time according to the overload information required, the flight parameters of the aircraft and the side slip angle convergence sliding mode surface.
Preferably, the convergence error solution submodule 31 obtains the convergence error of the attack angle and the convergence error of the sideslip angle in real time by the following formula (one),
wherein e is1Representing angle of attack convergence error, e2Denotes the sideslip angle convergence error, alpha denotes the angle of attack, beta denotes the sideslip angle, aycIndicating an expected pitch overload, azcIndicating an expected yaw overload, V indicating the speed of the aircraft, a34Representing the power coefficient of the rotary aircraft; the power coefficient of the rotary aircraft is known data preinstalled in the aircraft, and is generally obtained by calculation in wind tunnel experiments and other ways before the aircraft leaves a factory, and the data can be called at any time in the flight process of the aircraft.
Preferably, the convergence sliding mode surface resolving submodule 32 obtains an attack angle convergence sliding mode surface and a side sliding angle convergence sliding mode surface in real time through the following formula (two),
wherein s is1Representing the convergent sliding-form surface of angle of attack, s2Representing the sideslip angle converging sliding mode face, c and d both representing gain factors, theta representing the pitch angle, psi representing the yaw angle.
Preferably, the value of c is 230-150, more preferably 200 in the application, and the value of d is 0.5-2, more preferably 1 in the application.
Preferably, the steering engine response command resolving submodule 33 obtains the pitch direction steering engine response command and the yaw direction steering engine response command in real time through the following formulas (three) and (four),
wherein, deltayIndicating steering engine response command in pitch direction, deltazIndicating the yaw direction steering engine response command,
f1、f2、
ξ、
ξ2all are intermediate variables used in the resolving process, and have no practical physical significance;
a25、a24、a27、a22And a28The dynamic coefficient of the rotary aircraft is represented, and when the dynamic coefficient of the rotary aircraft is calculated, specific corresponding values can be directly taken from the aircraft.
The invention also provides a robust decoupling control method considering the coupling problem of the rotary aircraft, which is realized by the robust decoupling control system considering the coupling problem of the rotary aircraft,
the method comprises the following steps of,
step 1, receiving overload information required to be used transmitted by a guidance system through an overload receiving module 1;
step 3, a steering engine response instruction is obtained through the robust decoupling control module 3 according to the overload information required and the flight parameters of the aircraft;
and 4, repeating the steps 1-3 in real time, so as to obtain the steering engine response instruction in real time.
Preferably, said step 2 comprises the sub-steps of,
substep 2-1, obtaining acceleration information of the aircraft through real-time measurement of the accelerometer 21, obtaining yaw rate information and pitch rate information of the aircraft through real-time measurement of the inertial gyroscope 22,
in the substep 2-2, the attack angle and the sideslip angle of the aircraft are obtained by estimating in real time according to the triaxial acceleration information through the estimator 23;
and obtaining the speed information of the aircraft by integrating the triaxial acceleration information in real time through an integrator 24, obtaining the yaw angle information by integrating the yaw angle rate information in real time, and obtaining the pitch angle information by integrating the pitch angle rate information in real time.
Preferably, said step 3 comprises the sub-steps of,
in the substep 3-1, the convergence error of the attack angle and the convergence error of the sideslip angle are obtained in real time according to the overload information required and the flight parameters of the aircraft through the convergence error calculation submodule 31;
in the substep 3-2, an attack angle convergence sliding mode surface is obtained in real time according to the flight parameters and the attack angle convergence error of the aircraft through a convergence sliding mode surface resolving submodule 32, and a sideslip angle convergence sliding mode surface is obtained in real time according to the flight parameters and the sideslip angle convergence error of the aircraft;
and in the substep 3-3, acquiring a pitch direction steering engine response instruction in real time according to the overload information required, the flight parameters of the aircraft and the attack angle convergence sliding mode surface through the steering engine response instruction resolving submodule 33, and acquiring a yaw direction steering engine response instruction in real time according to the overload information required, the flight parameters of the aircraft and the side slip angle convergence sliding mode surface.
Simulation experiment:
carrying out simulation experiment of the rotary aircraft through a computer, wherein the simulation conditions of the rotary aircraft are as follows: the flying speed of the rotary aircraft is 580m/s, and the rotating speed is 11.6 r/s;
the guidance system and the steering engine system of the rotary aircraft can be directly simulated through a computer, the guidance system can give guidance instructions in real time, namely overload is required, specifically, the overload is required to be used and comprises expected pitch overload and expected yaw overload, and the time-varying track of the overload is shown as a solid line in fig. 2 and fig. 3; the steering engine system can control the steering engine to steer according to a guidance instruction or overload, and directly give an overload condition which can be actually provided for the rotary aircraft after the steering engine is controlled to work according to the overload;
in an experimental example, intercepting overload which is required and given by a guidance system in a computer, transmitting the overload which is required and given to a steering engine system, and transmitting the overload which is required and given to a robust decoupling control system which is provided by the application and considers the coupling problem of a rotary aircraft, obtaining a steering engine response instruction by the robust decoupling control method which is provided by the application and considers the coupling problem of the rotary aircraft, transmitting the steering engine response instruction to the steering engine system, controlling the steering engine to work according to the steering engine, and obtaining the overload condition which can be actually provided for the rotary aircraft after the steering engine works;
wherein, the overload information is received by the overload receiving module, i.e. the expected pitch overload ayc10 × square (t), expected yaw overload azc＝0；
Real-time providing flight parameters of the simulated aircraft through a computer, wherein the flight parameters comprise an attack angle, a sideslip angle, a speed, a pitch angle and a yaw angle; and the power coefficient of the rotary aircraft is given as follows:
resolving through the following formulas (I), (II), (III) and (IV) to obtain a pitching direction steering engine response instruction and a yawing direction steering engine response instruction;
aycindicating an expected pitch overload, azcRepresenting an expected yaw overload, alpha representing an angle of attack, beta representing a sideslip angle, V representing a speed of the aircraft, theta representing a pitch angle, psi representing a yaw angle, c representing 200, d representing 1, deltayIndicating steering engine response command in pitch direction, deltazIndicating yaw direction steering engine responseAnd (5) instructions.
After the steering engine response instruction is transmitted to a steering engine system in a computer, the overload condition which can be actually provided for the rotary aircraft after the steering engine works is obtained through simulation and is shown as a dotted line 'robust decoupling control response' in fig. 2 and 3.
In the comparative example, the overload demand given by the guidance system in the computer is intercepted, and is not directly transmitted to the steering engine system, and the overload demand is transmitted to the traditional three-loop autopilot used in the rotary aircraft in the prior art, and the traditional three-loop autopilot responds to the overload demand to obtain a steering engine control command, and controls the steering engine accordingly, and finally obtains the overload condition provided by the steering engine, as shown by a chain line 'traditional three-loop autopilot response' in fig. 2 and 3.
As can be seen from fig. 2 and 3, the conventional three-loop autopilot has a coupling effect, when an overload is expected to be adjusted in the pitching direction, the fluctuation of the overload in the yawing direction is large, and the convergence rate of the overload in the yawing direction is slow, which inevitably has an adverse effect on the attitude and the hit precision of the aircraft; the robust decoupling control system considering the coupling problem of the rotary aircraft, which is provided by the application, not only eliminates the coupling influence on yaw control during pitch control, but also has high convergence speed and obvious advantages, and as shown in fig. 3, the robust decoupling control response basically coincides with the expected yaw overload.
The present invention has been described above in connection with preferred embodiments, but these embodiments are merely exemplary and merely illustrative. On the basis of the above, the invention can be subjected to various substitutions and modifications, and the substitutions and the modifications are all within the protection scope of the invention.
Claims (9)
1. A robust decoupled control system that takes into account the coupling problem of a rotating aircraft, the system being mounted on the rotating aircraft, the system comprising
The overload receiving module (1) is connected with a guidance system on the rotary aircraft and used for receiving the overload information which is transmitted by the guidance system in real time,
an aircraft parameter measurement module (2) for obtaining flight parameters of the aircraft in real time, and
the robust decoupling control module (3) is used for obtaining a steering engine response instruction in real time according to the overload information required and flight parameters of the aircraft;
the robust decoupling control module (3) comprises a convergence error resolving submodule (31), a convergence sliding mode surface resolving submodule (32) and a steering engine response instruction resolving submodule (33);
the convergence error resolving submodule (31) is used for obtaining an attack angle convergence error and a sideslip angle convergence error in real time according to overload information required and flight parameters of an aircraft;
the convergence sliding mode surface resolving submodule (32) is used for obtaining an attack angle convergence sliding mode surface in real time according to flight parameters and an attack angle convergence error of the aircraft and obtaining a sideslip angle convergence sliding mode surface in real time according to the flight parameters and the sideslip angle convergence error of the aircraft;
the steering engine response instruction resolving submodule (33) is used for obtaining a pitch direction steering engine response instruction in real time according to the overload information, the flight parameters of the aircraft and the attack angle convergence sliding mode surface, and obtaining a yaw direction steering engine response instruction in real time according to the overload information, the flight parameters of the aircraft and the side slip angle convergence sliding mode surface.
2. The robust decoupling control system in view of rotary aerial vehicle coupling issues of claim 1,
the required overload information comprises expected pitch overload and expected yaw overload;
the steering engine response command comprises a pitching direction steering engine response command and a yawing direction steering engine response command.
3. The robust decoupling control system in view of rotary aerial vehicle coupling issues of claim 1,
the aircraft parameter measurement module (2) comprises an accelerometer (21), an inertial gyro (22), an estimator (23) and an integrator (24);
wherein the accelerometer (21) is used for measuring acceleration information of the aircraft in real time,
the inertial gyroscope (22) is used for measuring yaw rate information and pitch rate information of the aircraft in real time,
the estimator (23) is used for estimating in real time according to the triaxial acceleration information to obtain attack angle information and sideslip angle information of the aircraft;
the integrator (24) is used for integrating in real time according to the triaxial acceleration information to obtain speed information of the aircraft, integrating in real time according to the yaw angle rate information to obtain yaw angle information, and integrating in real time according to the pitch angle rate information to obtain pitch angle information.
4. The robust decoupling control system in view of rotary aerial vehicle coupling issues of claim 1,
the convergence error resolving submodule (31) obtains an attack angle convergence error and a sideslip angle convergence error in real time through the following formula (I),
wherein e is1Representing angle of attack convergence error, e2Denotes the sideslip angle convergence error, alpha denotes the angle of attack, beta denotes the sideslip angle, aycIndicating an expected pitch overload, azcIndicating an expected yaw overload, V indicating the speed of the aircraft, a34Representing the power coefficient of the rotating aircraft.
5. The robust decoupling control system in view of rotary aerial vehicle coupling issues of claim 1,
the convergence sliding mode surface resolving submodule (32) obtains an attack angle convergence sliding mode surface and a side sliding angle convergence sliding mode surface in real time through the following formula (II),
wherein s is1Representing the convergent sliding-form surface of angle of attack, s2Representing the sideslip angle converging sliding mode face, c and d both representing gain factors, theta representing the pitch angle, psi representing the yaw angle.
6. The robust decoupling control system in view of rotary aerial vehicle coupling issues of claim 1,
the steering engine response instruction resolving submodule (33) obtains a pitching direction steering engine response instruction and a yawing direction steering engine response instruction in real time through the following formulas (three) and (four),
wherein, deltayIndicating steering engine response command in pitch direction, deltazIndicating yaw direction steering engine response command, f1、f2、
ξ、
ξ2All are intermediate variables used in the resolving process, have no practical physical significance,
the derivative of the representation of the derivative of xi,
is representative of xi2The derivative of (a) of (b),
a25、a24、a27、a22and a28All represent rotating fliesThe power coefficient of the travelling device.
7. A robust decoupling control method taking into account the coupling problem of a rotary aircraft, implemented by a robust decoupling control system taking into account the coupling problem of a rotary aircraft according to one of claims 1 to 6,
the method comprises the following steps of,
step 1, receiving overload information required to be used transmitted by a guidance system through an overload receiving module (1);
step 2, obtaining flight parameters of the aircraft through the aircraft parameter measuring module (2);
step 3, a steering engine response instruction is obtained through the robust decoupling control module (3) according to the overload information required and the flight parameters of the aircraft;
and 4, repeating the steps 1-3 in real time, so as to obtain the steering engine response instruction in real time.
8. The robust decoupling control method in view of rotating aerial vehicle coupling issues of claim 7,
said step 2 comprises the sub-steps of,
substep 2-1, obtaining acceleration information of the aircraft through real-time measurement of an accelerometer (21), obtaining yaw rate information and pitch rate information of the aircraft through real-time measurement of an inertial gyroscope (22),
in the substep 2-2, the attack angle and the sideslip angle of the aircraft are obtained by estimating in real time according to the triaxial acceleration information through an estimator (23);
and integrating in real time according to the triaxial acceleration information through an integrator (24) to obtain speed information of the aircraft, integrating in real time according to the yaw angle rate information to obtain yaw angle information, and integrating in real time according to the pitch angle rate information to obtain pitch angle information.
9. The robust decoupling control method in view of rotating aerial vehicle coupling issues of claim 7,
said step 3 comprises the sub-steps of,
in the substep 3-1, the convergence error of the attack angle and the convergence error of the sideslip angle are obtained in real time according to the overload information required and the flight parameters of the aircraft through a convergence error resolving submodule (31);
in the substep 3-2, an attack angle convergence sliding mode surface is obtained in real time according to flight parameters and an attack angle convergence error of the aircraft through a convergence sliding mode surface resolving submodule (32), and a sideslip angle convergence sliding mode surface is obtained in real time according to the flight parameters and the sideslip angle convergence error of the aircraft;
and in the substep 3-3, a steering engine response instruction in the pitching direction is obtained in real time according to the overload information required, the flight parameters of the aircraft and the attack angle convergence sliding mode surface through a steering engine response instruction resolving submodule (33), and a yaw direction steering engine response instruction is obtained in real time according to the overload information required, the flight parameters of the aircraft and the side sliding angle convergence sliding mode surface.
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