CN109062047B - Method and system for resolving slow loop control instruction based on addition information in dynamic inverse control - Google Patents
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Abstract
The method and the system for resolving the slow loop control instruction based on the metering information in the dynamic inverse control acquire the navigation information output by an aircraft navigation system, the measurement information output by the triaxial metering and an expected attitude control instruction output by a guidance system; filtering the measurement information output by the triaxial adding meter; establishing a slow loop inverse model according to the navigation information and the filtered triaxial accelerometer measurement information; generating a pseudo control instruction according to navigation information output by a navigation system of the aircraft and an expected attitude control instruction output by a guidance system; and finally, a slow loop control instruction is generated, so that the resolving of a complex inverse model of the slow loop and the online compensation of errors are avoided, the interference except the navigation system errors (including gravity calculation errors) can be resisted, and the slow loop control law resolving with low time consumption and high precision is realized.
Description
Technical Field
The invention relates to a method and a system for resolving a slow loop control instruction in dynamic inverse control, in particular to a method for resolving the slow loop control instruction based on addition information in the dynamic inverse control, and belongs to the field of nonlinear attitude control of hypersonic aircrafts.
Background
The hypersonic aircraft has large flight airspace/speed domain span, violent change of kinetic characteristics and model parameters, strong dynamic nonlinearity, coupling and uncertainty and serious internal and external interference. In the traditional attitude control design based on the PID control law, various means such as root tracks, frequency domain graphs and time domain response curves are adopted for auxiliary design, and meanwhile, in order to adapt to a large-change working range and a complex control mode, a large number of characteristic points are required to be selected for designing control parameters. Therefore, the control system design of the hypersonic aircraft faces the problems of large design task amount, low efficiency, long time consumption and the like.
The dynamic inverse control law is very suitable for nonlinear, strong-coupling, multivariable and parameter time-varying controlled objects, the design process is simple, the design efficiency is high, and the method is a potential method capable of effectively solving the attitude control problem of the hypersonic aircraft. However, the method depends on an accurate object model, uncertain factors obviously influence the inverse error of the object model, so that the control precision is reduced, and the designed control system has poor robustness.
In order to improve the anti-interference capability and control precision of dynamic inverse control, a more common method in domestic and foreign researches is to perform online compensation on model inverse errors by adopting a neural network so as to achieve the purpose of building an inverse model with high precision. However, because of the complex dynamics of the aircraft, the prior art needs to introduce a neural network of multiple layers or multiple neurons to perform online identification and compensation on the inverse error of the fast/slow loop, so as to obtain higher control precision. Because the external force calculation needs interpolation calculation, the calculation of the nominal inverse model is complex, and the inverse error compensation further causes low calculation efficiency and poor real-time performance of a control algorithm, thereby influencing engineering application.
At present, documents published at home and abroad pay little attention to the resolving efficiency of the dynamic inverse control law, and no article exists for researching the technology and the method for resolving the slow loop control instruction based on the addition information.
Disclosure of Invention
The technical problem solved by the invention is as follows: the method and the system for solving the slow loop control instruction based on the addition information are provided for overcoming the defects of the prior art and solving the problems of low solving efficiency and poor algorithm real-time property caused by the adoption of a dynamic inverse control law of the air-breathing hypersonic speed aircraft. The interference except the navigation system error (including the gravity calculation error) can be resisted, and the high-precision slow loop control law calculation can be realized. The calculated slow loop model is simple, a complex slow loop nominal inverse model does not need to be calculated, an inverse error does not need to be compensated, the calculation efficiency and the real-time performance are improved, and the problems of low technical efficiency and poor real-time performance of the existing method are solved.
The technical scheme of the invention is as follows: the method for resolving the slow loop control instruction based on the addition information in the dynamic inverse control comprises the following steps:
(1) acquiring navigation information output by a navigation system of an aircraft, triaxial accelerometer measurement information and an expected attitude control instruction output by the navigation system;
(2) filtering the measurement information output by the triaxial accelerometer in the step (1) by adopting a first-order digital filter to obtain filtered measurement information of the triaxial accelerometer;
(3) establishing a slow loop inverse model according to the navigation information output by the navigation system of the aircraft in the step (1) and the filtered three-axis measuring information of the accelerometer in the step (2);
(4) generating a pseudo control instruction according to the navigation information output by the navigation system of the aircraft in the step (1) and the expected attitude control instruction output by the guidance system;
(5) generating a control matrix according to the navigation information output by the navigation system of the aircraft in the step (1);
(6) and (4) generating a slow loop control instruction according to the slow loop inverse model in the step (3), the pseudo control instruction in the step (4) and the control matrix in the step (5).
The specific requirements of the aircraft in the step (1) are as follows: at least comprises the following steps: a navigation system, a three-axis metering and guidance system; the aircraft adopts a nonlinear dynamic inverse control law, and a controller is designed based on a fast loop and a slow loop, wherein the slow loop controller outputs an expected triaxial attitude angular rate control instruction, and the fast loop controller outputs an expected execution mechanism control instruction. The navigation system outputs navigation information required for resolving by a guidance control law, the three-axis accelerometer measures the apparent acceleration of the aircraft, and the guidance system operates the guidance law and outputs an expected attitude control instruction.
Navigation information, including: attack angle alpha, sideslip angle beta, roll angle sigma, airspeed V of aircraft, and triaxial gravity acceleration component g under body coordinate systemx、gy、gzAnd the three-axis angular rate component omegax、ωy、ωz。
The body coordinate system is defined as follows: the origin O is the center of mass of the aircraft, the positive direction of the X axis is the head pointed by the longitudinal symmetric axis of the aircraft, the positive direction of the Y axis is the upward pointed direction perpendicular to the X axis, and the Z axis is determined by the right-hand rule.
The three-axis and design union is installed on the aircraft body, and the sensitive axis of the three-axis and design union is parallel to the coordinate axis of the body coordinate system.
The expected attitude control command output by the guidance system is an expected attack angle alphacDesired side slip angle betacDesired roll angle σc。
Filtering the measurement information output by the triaxial accelerometer in the step (1), and resolving by adopting a first-order digital filter to obtain filtered apparent acceleration information axl、ayl、azl。
Step (3) establishing a slow loop inverse model f according to the navigation information output by the navigation system of the aircraft in the step (1) and the three-axis apparent acceleration information in the step (2)m＝f(axl,ayl,azl,gx,gy,gzα, β, V), the solution is completed.
Step (4) generating a pseudo control instruction U according to the navigation information output by the navigation system of the aircraft in the step (1) and the expected attitude control instruction output by the guidance systemdThe method comprises the following steps:
generating pseudo control commands using proportional plus derivative controllers
KpIs a proportionality coefficient, KdIs a differential coefficient.
Step (5) according to the navigation information output by the navigation system of the aircraft in the step (1),generating a control matrix GmThe method comprises the following steps:
and (6) generating a slow loop control command omega based on a nonlinear dynamic inverse control principle according to the slow loop inverse model in the step (3), the pseudo control command in the step (4) and the control matrix in the step (5)c，
The system for resolving the slow loop control instruction based on the addition information in the dynamic inverse control comprises the following steps: the device comprises an acquisition module, a filtering processing module, a model establishing module, a first instruction generating module, a matrix generating module and a second instruction generating module;
the acquisition module acquires navigation information output by a navigation system of the aircraft, triaxial accelerometer measurement information and an expected attitude control instruction output by the guidance system;
the filtering processing module is used for filtering the measurement information output by the triaxial accelerometer of the acquisition module by adopting a first-order digital filter to obtain filtered measurement information of the triaxial accelerometer;
the model building module is used for building a slow loop inverse model according to the navigation information output by the navigation system of the aircraft of the obtaining module and the three-axis measuring information filtered by the filtering processing module;
the first instruction generation module is used for generating a pseudo control instruction according to the navigation information output by the navigation system of the aircraft of the acquisition module and the expected attitude control instruction output by the guidance system;
the matrix generation module is used for generating a control matrix according to the navigation information output by the navigation system of the aircraft of the acquisition module;
and the second instruction generation module is used for generating a slow loop control instruction according to the slow loop inverse model of the model establishment module, the pseudo control instruction of the first instruction generation module and the control matrix of the second instruction generation module.
Compared with the prior art, the invention has the advantages that:
(1) the method for calculating the slow loop control instruction based on the addition information in the dynamic inverse control utilizes the addition information to construct a slow loop inverse model with a simple function form and complete the calculation of the slow loop control instruction, and improves the calculation efficiency, the real-time performance and the calculation precision of the slow loop control instruction calculation.
(2) The slow loop inverse model which is constructed by the method and takes the addition information as the variable is simple, does not contain a thrust and aerodynamic force calculation model or an error online compensation link, and has high calculation efficiency.
(3) The slow loop inverse model with the addition information as a variable can resist the interference except the navigation system error, and the high-precision calculation of the slow loop inverse model can be realized without error online compensation.
(4) The pseudo control instruction resolving adopts a proportion + differentiation control law, so that the resolving of the whole slow loop control instruction is simpler.
(5) The method realizes the slow loop control law calculation with high precision and low time consumption, can promote the fast loop to realize the high-precision fast loop inverse error compensation only by using a neural network with smaller scale, can ensure the real-time performance of the whole robust dynamic inverse control law, and has stronger engineering application capability.
Drawings
FIG. 1 is a flow chart of the method of the present invention;
FIG. 2 is a graph of CPU time consumption versus slow loop control command resolution;
FIG. 3 is a graph of angle of attack tracking;
FIG. 4 is a graph of a side slip angle tracking.
Detailed Description
The invention is described in further detail below with reference to the figures and specific embodiments.
The method and the system for resolving the slow loop control instruction based on the metering information in the dynamic inverse control acquire the navigation information output by an aircraft navigation system, the measurement information output by the triaxial metering and an expected attitude control instruction output by a guidance system; filtering the measurement information output by the triaxial adding meter; establishing a slow loop inverse model according to the navigation information and the filtered triaxial accelerometer measurement information; generating a pseudo control instruction according to navigation information output by a navigation system of the aircraft and an expected attitude control instruction output by a guidance system; and finally, a slow loop control instruction is generated, so that the resolving of a complex inverse model of the slow loop and the online compensation of errors are avoided, the interference except the navigation system errors (including gravity calculation errors) can be resisted, and the slow loop control law resolving with low time consumption and high precision is realized.
The invention relates to a system for resolving a slow loop control instruction based on addition information in dynamic inverse control, which comprises the following steps: the device comprises an acquisition module, a filtering processing module, a model establishing module, a first instruction generating module, a matrix generating module and a second instruction generating module;
the acquisition module acquires navigation information output by a navigation system of the aircraft, triaxial accelerometer measurement information and an expected attitude control instruction output by the guidance system;
the filtering processing module is used for filtering the measurement information output by the triaxial accelerometer of the acquisition module by adopting a first-order digital filter to obtain filtered measurement information of the triaxial accelerometer;
the model building module is used for building a slow loop inverse model according to the navigation information output by the navigation system of the aircraft of the obtaining module and the three-axis measuring information filtered by the filtering processing module;
the first instruction generation module is used for generating a pseudo control instruction according to the navigation information output by the navigation system of the aircraft of the acquisition module and the expected attitude control instruction output by the guidance system;
the matrix generation module is used for generating a control matrix according to the navigation information output by the navigation system of the aircraft of the acquisition module;
and the second instruction generation module is used for generating a slow loop control instruction according to the slow loop inverse model of the model establishment module, the pseudo control instruction of the first instruction generation module and the control matrix of the second instruction generation module.
As shown in FIG. 1, the invention is used for solving the slow loop control command in the nonlinear dynamic inverse control of the aircraft, and the expected triaxial attitude angular rate control command is output as the input of the fast loop controller. The method for resolving the slow loop control instruction based on the addition information in the dynamic inverse control comprises the following specific steps:
(1) and acquiring navigation information output by a navigation system of the aircraft, triaxial accelerometer measurement information and an expected attitude control instruction output by the guidance system. The navigation information includes: attack angle alpha, sideslip angle beta, roll angle sigma, airspeed V of aircraft, and triaxial gravity acceleration component g under body coordinate systemx、gy、gzAnd the three-axis angular rate component omegax、ωy、ωz. The three-axis accelerometer measurement information is apparent acceleration information ax、ay、az. The desired attitude control command is a desired angle of attack αcDesired side slip angle betacDesired roll angle σc。
(2) Filtering the measurement information output by the triaxial accelerometer in the step (1) by adopting a first-order digital filter, removing measurement noise, and obtaining filtered triaxial apparent acceleration information axl、ayl、azl. The calculation formula is as follows:
t is the time constant, dt is the sampling period, and the subscript i represents the coordinate axes x, y, z, ai(n)For this sampled value, ail(n-1)For the last filtered output value, ail(n)Is the output value of the filtering.
(3) Establishing a slow loop inverse model f based on the dynamic characteristics of the slow loop inverse model and the model stress information represented by the measurement information according to the navigation information output by the navigation system of the aircraft in the step (1) and the filtered triaxial measurement information in the step (2)m(x) And completing model calculation. The model is as follows:
(4) according to the three-axis angular rate component omega output by the navigation system of the aircraft in the step (1)x、ωy、ωzAnd the expected attitude control command alpha output by the guidance systemc、βc、σcGenerating a pseudo control instruction Ud. In order to improve the dynamic quality of attitude tracking, a proportion + differential control law is adopted. U shapedThe solution formula is as follows:
Kpthe value range is preferably [ -100, 100 ] for the proportionality coefficient],KdFor differential coefficient, the value range is preferably [ -100, 100 [ -100 [ ]]
(5) Generating a control matrix G according to the navigation information output by the navigation system of the aircraft in the step (1)m. The solving formula is as follows:
(6) slow loop inverse model f according to step (3)mAnd (4) pseudo control instruction UdAnd the control matrix G of step (5)mGenerating a slow loop control command omega based on a nonlinear dynamic inverse control principlec：
The preferred protocol is as follows (one-step running calculation):
inputting: navigation information alpha is 0.017, sideslip angle beta is 0, roll angle sigma is 0, airspeed V is 1000, and three-axis gravity acceleration component gx＝-3.828、gy＝-8.817、gz0.136, three-axis angular rate component ωx＝0、ωy＝-0.004、ωz-0.0255; three-axis accelerometer measurement information ax＝4.106、ay＝7.619、az-0.005; desired postureState control command, desired angle of attack alphac0.035, desired sideslip angle βcDesired roll angle σ of 0c＝0.087。
Outputting by a filter in the last step: a isxl-1＝4.106、ay-1＝7.619、az-1＝-0.005。
Parameters are as follows: the time constant T is 0.006, and the sampling period dt of the system is 0.01; coefficient of proportionality Kp5.0, differential coefficient Kd＝-0.4。
Calculating to obtain a slow loop control instruction: omegac＝[0.4363 0.0018 0.0963]T。
The invention is subjected to digital comparative simulation test. The case is to control the hypersonic flight vehicle to fly for 10s, the attack angle of the vehicle is required to be stabilized at 2 degrees, and the sideslip angle and the roll angle are required to be 0 degree. The attack angle, the sideslip angle and the roll angle of the aircraft at the starting moment are all 0 degree.
Carrying out comparison simulation of solving efficiency: the calculation of the slow loop inverse model in the traditional method adopts a nominal inverse model and neural network compensation, the average value of the CPU time consumption calculated by the slow loop control instruction in each control period is 80 mus, the average value of the CPU time consumption is 2us, and the calculation efficiency and the real-time performance are greatly improved. The slow loop control instruction resolved CPU time consumption versus curve is shown in fig. 2. FIG. 2 is a comparison graph of CPU time consumption for slow loop control instruction resolution, which shows that the method of the present invention has higher resolution efficiency and real-time performance than the conventional method; FIG. 3 is a graph of angle of attack tracking, showing that the present invention achieves high accuracy slow loop inverse model solution; FIG. 4 is a graph of a sideslip angle tracking curve, which shows that the invention realizes high-precision slow-loop inverse model solution.
As shown in fig. 2, the slow loop control instruction solves a CPU time consumption comparison curve, the traditional slow loop inverse model uses a nominal inverse model + neural network compensation for the solution, the slow loop control instruction solves a CPU time consumption average value of 80 μ s in each control cycle, and the CPU time consumption average value of the invention is 2 us;
as shown in FIG. 3, in an attack angle tracking curve, the thrust deviation is 10%, the thrust line transversely moves by 0.05m, when the normal inverse model is not compensated, the slow loop inverse model is solved, the steady-state error of 0.2 degree exists in the attack angle, and the steady-state error does not exist in the attack angle.
As shown in FIG. 4, the sideslip angle tracking curve has thrust deviation of 10% and transverse thrust line of 0.05m, when uncompensated, a nominal inverse model is adopted to solve a slow loop inverse model, the sideslip angle still has a tracking error of 0.25 degrees at the time of 10s, and the tracking error of the sideslip angle is less than 0.01 degrees after 2s of the method is adopted
Carrying out comparison simulation of solving precision: assuming that the thrust deviation of the aircraft is 10% and the thrust line transversely moves by 0.05m, the attack angle tracking curve is shown in FIG. 3, and the sideslip angle tracking curve is shown in FIG. 4. And when the compensation is not carried out, a nominal inverse model is adopted to solve a slow loop inverse model. As can be seen from fig. 3, when the angle of attack is not compensated, a steady-state error of 0.2 degrees exists, and the angle of attack has no steady-state error by using the method of the present invention. As can be seen from fig. 4, when the compensation is not performed, the tracking error of the sideslip angle still exists at the time of 10s, and the tracking error of the sideslip angle is less than 0.01 degrees after the invention is adopted for 2 s. The higher attitude control precision shows that the invention realizes high-precision slow loop inverse model calculation and can complete high-precision attitude control.
According to the method for calculating the slow loop control instruction based on the addition information in the dynamic inverse control, the addition information is utilized to construct the slow loop inverse model with a simple function form and complete the calculation of the slow loop control instruction, so that the calculation efficiency, the real-time performance and the calculation precision of the slow loop control instruction calculation are improved. The slow loop inverse model which is constructed by the method and takes the addition information as the variable is simple, does not contain a thrust and aerodynamic force calculation model or an error online compensation link, and has high calculation efficiency.
The slow loop inverse model with the addition information as the variable can resist the interference except the navigation system error, and the high-precision calculation of the slow loop inverse model can be realized without error on-line compensation. The pseudo control instruction resolving adopts a proportion + differentiation control law, so that the resolving of the whole slow loop control instruction is simpler.
The method realizes the slow loop control law calculation with high precision and low time consumption, can promote the fast loop to realize the high-precision fast loop inverse error compensation only by using a neural network with smaller scale, can ensure the real-time performance of the whole robust dynamic inverse control law, and has stronger engineering application capability.
The invention has not been described in detail in part of the common general knowledge of those skilled in the art.
Claims (6)
1. The method for resolving the slow loop control instruction based on the addition information in the dynamic inverse control is characterized by comprising the following steps of:
(1) acquiring navigation information output by a navigation system of an aircraft, triaxial accelerometer measurement information and an expected attitude control instruction output by the guidance system;
(2) filtering the measurement information output by the triaxial accelerometer in the step (1) by adopting a first-order digital filter to obtain filtered measurement information of the triaxial accelerometer;
(3) establishing a slow loop inverse model according to the navigation information output by the navigation system of the aircraft in the step (1) and the filtered three-axis measuring information of the accelerometer in the step (2);
slow loop inverse model
In the formula, the attack angle of the aircraft is alpha, the sideslip angle is beta, the airspeed is V, and the three-axis gravity acceleration components under the body coordinate system are g respectivelyx、gy、gz；
Filtering the measurement information output by the triaxial accelerometer in the step (1), and resolving by adopting a first-order digital filter to obtain filtered apparent acceleration information axl、ayl、azl；
(4) Generating a pseudo control instruction U according to the navigation information output by the navigation system of the aircraft in the step (1) and the expected attitude control instruction output by the guidance systemd；
(5) Generating a control matrix G according to the navigation information output by the navigation system of the aircraft in the step (1)m；
(6) Slow loop inverse model f according to step (3)mAnd (4) pseudo control instruction UdAnd the control matrix G of step (5)mGenerating a slow loop control command omegac：
2. The method for resolving the slow loop control instruction based on the addition information in the dynamic inverse control according to claim 1, wherein: the specific requirements of the aircraft in the step (1) are as follows: at least comprises the following steps: a navigation system, a three-axis metering and guidance system; the aircraft adopts a nonlinear dynamic inverse control law, the controller is designed based on a fast loop and a slow loop, the slow loop controller outputs an expected three-axis attitude angular rate control instruction, the fast loop controller outputs an expected execution mechanism control instruction, the navigation system outputs navigation information required for resolving a guidance control law, the three-axis accelerometer measures the apparent acceleration of the aircraft, and the guidance system operates the guidance law and outputs an expected attitude control instruction.
3. The method for resolving the slow loop control instruction based on the addition information in the dynamic inverse control according to claim 1, wherein: the body coordinate system is defined as follows: the origin O is the center of mass of the aircraft, the positive direction of the X axis is the head pointed by the longitudinal symmetric axis of the aircraft, the positive direction of the Y axis is the upward pointed direction perpendicular to the X axis, and the Z axis is determined by the right-hand rule.
4. The method for resolving the slow loop control instruction based on the addition information in the dynamic inverse control according to claim 1, wherein: step (4) generating a pseudo control instruction U according to the navigation information output by the navigation system of the aircraft in the step (1) and the expected attitude control instruction output by the guidance systemdThe method comprises the following steps:
generating pseudo control commands using proportional plus derivative controllers
Wherein the expected attitude control command output by the guidance system is expectedAngle of attack alphacDesired side slip angle betacDesired roll angle σcThe roll angle is sigma, and the triaxial angular rate component is omegax、ωy、ωz，KpIs a proportionality coefficient, KdIs a differential coefficient.
5. The method for resolving the slow loop control instruction based on the addition information in the dynamic inverse control according to claim 1, wherein: step (5) generating a control matrix G according to the navigation information output by the navigation system of the aircraft in the step (1)mThe method comprises the following steps:
6. the system for resolving the slow loop control instruction based on the addition information in the dynamic inverse control is characterized by comprising the following steps: the device comprises an acquisition module, a filtering processing module, a model establishing module, a first instruction generating module, a matrix generating module and a second instruction generating module;
the acquisition module acquires navigation information output by a navigation system of the aircraft, triaxial accelerometer measurement information and an expected attitude control instruction output by the guidance system;
the filtering processing module is used for filtering the measurement information output by the triaxial accelerometer of the acquisition module by adopting a first-order digital filter to obtain filtered measurement information of the triaxial accelerometer;
the model building module is used for building a slow loop inverse model according to the navigation information output by the navigation system of the aircraft of the obtaining module and the three-axis measuring information filtered by the filtering processing module;
slow loop inverse model
In the formula, the attack angle of the aircraft is alpha, the sideslip angle is beta, the airspeed is V, and the three-axis gravity acceleration components under the body coordinate system are g respectivelyx、gy、gz；
Filtering the measurement information output by the three-axis accelerometer, and resolving by using a first-order digital filter to obtain filtered apparent acceleration information axl、ayl、azl；
The first instruction generation module generates a pseudo control instruction U according to the navigation information output by the navigation system of the aircraft of the acquisition module and the expected attitude control instruction output by the guidance systemd；
A matrix generation module for generating a control matrix G according to the navigation information output by the navigation system of the aircraft of the acquisition modulem；
A second instruction generation module for establishing a slow loop inverse model f of the module according to the modelmPseudo control instruction U of first instruction generation moduledAnd a control matrix G of a second instruction generation modulemGenerating a slow loop control command omegac；
Wherein, ω isx、ωy、ωzIs omegacThe three-axis angular rate component of (a).
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