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CN119781518A - Aircraft self-adaptive thrust steering control method based on attitude measurement - Google Patents

Aircraft self-adaptive thrust steering control method based on attitude measurement Download PDF

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Publication number
CN119781518A
CN119781518A CN202411952141.3A CN202411952141A CN119781518A CN 119781518 A CN119781518 A CN 119781518A CN 202411952141 A CN202411952141 A CN 202411952141A CN 119781518 A CN119781518 A CN 119781518A
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pitch angle
error
rate
aircraft
pitch
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CN119781518B (en
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王赫
李潇
麻彦轩
林振强
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Beijing Great Wall Aviation Measurement And Control Technology Research Institute Co ltd
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Beijing Great Wall Aviation Measurement And Control Technology Research Institute Co ltd
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    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
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    • Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation

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Abstract

本发明属于飞行器控制技术领域，提供一种基于姿态测量的飞行器自适应推力转向控制方法，其步骤包括：S1、测量飞行器的俯仰角速率、俯仰角和速度俯仰倾角；S2、求解俯仰角误差信号和俯仰角误差积分信号；S3、求解误差相关项自适应信号；S4、求解角速率相关项自适应信号：S5、根据飞行器推力转向俯仰总控制信号实现转向控制。本发明通过空气动力学系数构建等效控制项模型，结合各总信号组成推力转向总控制信号，实现飞行器对期望俯仰角跟踪与调控，通过对速度俯仰倾角、俯仰角飞行控制参数进行实时补偿，调整控制增益参数使飞行器适应环境和气动参数变化，从而使飞行器控制系统能实时动态补偿干扰因素，提高飞行器的控制精度和稳定性。

The present invention belongs to the field of aircraft control technology, and provides an aircraft adaptive thrust steering control method based on attitude measurement, and the steps include: S1, measuring the pitch rate, pitch angle and speed pitch inclination of the aircraft; S2, solving the pitch angle error signal and the pitch angle error integral signal; S3, solving the error related item adaptive signal; S4, solving the angular rate related item adaptive signal; S5, realizing steering control according to the aircraft thrust steering pitch total control signal. The present invention constructs an equivalent control item model through aerodynamic coefficients, combines each total signal to form a thrust steering total control signal, realizes the aircraft tracking and regulation of the desired pitch angle, and adjusts the control gain parameters to make the aircraft adapt to the environment and aerodynamic parameter changes by real-time compensation of the speed pitch inclination and pitch angle flight control parameters, so that the aircraft control system can dynamically compensate for interference factors in real time, and improve the control accuracy and stability of the aircraft.

Description

Aircraft self-adaptive thrust steering control method based on attitude measurement
Technical Field
The invention belongs to the technical field of attitude control of aircrafts, and particularly relates to an aircraft self-adaptive thrust steering control method based on attitude measurement.
Background
Aircraft, particularly shaped aircraft, are paid attention to because of their unique aerodynamic profile, control is also more difficult than conventional aircraft, and particularly elliptical and disc-shaped aircraft are often controlled by redundant control arrangements, which on the one hand reduce the difficulty of control and on the other hand can remain stable when a single control is unstable.
The control concept of the existing aircraft is based on the consideration of two dimensions from rotation moment stabilization to vertical stress stabilization of a pitching channel, certain complexity exists, the traditional aircraft generally adopts composite control of a thrust steering system and a mass distance dual control system, one control system is responsible for stabilizing balance of force, the other control system is responsible for balancing mutual matching of moment, but under special conditions, such as the condition that the mass distance control system fails, the independent thrust steering control can also support basic stable flight of the aircraft. It is therefore necessary to propose an aircraft adaptive thrust steering control method based on attitude measurements.
Disclosure of Invention
Aiming at the problems existing in the prior art, the self-adaptive thrust steering control method of the aircraft based on attitude measurement is characterized in that the self-adaptive compensation is carried out on the speed pitching dip angle, the pitch angle rate and the uncertainty related to a pitch angle error hinge according to the pitch angle error and the pitch angle rate by measuring the speed pitching dip angle, the pitch angle and the pitch angle rate, and then an equivalent control item is constructed through an aerodynamic coefficient of a model, and the total thrust steering control quantity is formed by combining the pitch angle rate, the pitch angle error integral and the self-adaptive signal, so that the tracking and the regulation of the expected pitch angle of the aircraft are realized, and the control precision and the stability of the aircraft are improved.
The invention provides an aircraft self-adaptive thrust steering control method based on attitude measurement, which comprises the following steps:
S1, measuring pitch angle rate omega z and pitch angle of an aircraft And a speed pitch angle θ;
S2, mounting a thrust steering device at the axial tail part of the flying direction of the aircraft, wherein the thrust pitching swing angle of the thrust steering device is xi, and the expected pitch angle of the aircraft is set as Comparing with the pitch angle omega z of the aircraft to obtain a pitch angle error e, and integrating to obtain a pitch angle error integral s 1 of the aircraft, specifically:
s3, solving an error related item self-adaptive signal;
S4, solving an angular rate related term self-adaptive signal:
s5, solving a total control signal of thrust steering and pitching of the aircraft to realize steering control, wherein the substeps comprise the following steps:
S51, pitching the dip angle theta and the pitch angle according to the speed of the aircraft And aerodynamic parameters of the aircraft, designing a thrust steering equivalent control signal f e of the aircraft:
wherein a 24、az is the aerodynamic parameter of the aircraft and is constant;
The expression of the dynamics control model in the aircraft pitching channel control is as follows:
Δu=K1Δα+K1Δq
Wherein Deltau is a control input signal of the elevator, deltaalpha is pitch angle deviation, deltaq is pitch angle speed deviation, and K 1、K1 is control gain;
After receiving the pitch angle error e and the pitch angle rate omega z, the pitch angle theta and the pitch angle of the speed are adjusted The flight control parameters are compensated, and the aircraft is adapted to the environmental and aerodynamic parameter changes by adjusting the control gain parameters, so that the aircraft control system can dynamically compensate the interference factors in real time;
S52, superposing a pitch angle error e, a pitch angle error integral S 1, a pitch angle rate omega z, an error related item self-adaptive signal f c and an angular rate related item self-adaptive signal f d to form a thrust steering pitch total control signal of the aircraft, and transmitting the thrust steering pitch total control signal to a thrust steering device to control the thrust pitch swing angle of an engine to be equal to the thrust steering pitch total control signal of the aircraft, wherein the method specifically comprises the following steps of:
ξa＝-l1ωz-l2e-l3s1+fc+fd-fe;
Wherein, xi a is the total control signal of thrust steering and pitching, l 1、l2、l3 is the control parameter, and both are constants;
And constructing an equivalent control item through the model aerodynamic coefficient, and combining a pitch angle rate, a pitch angle error integral and a self-adaptive signal to form a thrust steering total control signal so as to realize the regulation and control of the aircraft on the expected pitch angle.
Preferably, the step S3 includes the following substeps:
S31, solving an error and speed pitching inclination angle hinge factor self-adaptive growth rate c 1d according to a pitching angle error e and a speed pitching inclination angle theta, and integrating to obtain an error and speed pitching inclination angle hinge self-adaptive signal c 1;
c1(n+1)＝c1(n)+c1dT
wherein k 1、k2 is a constant parameter used for adjusting the self-adaptive signal growth rate of the error and the speed, pitch and inclination angle hinge, epsilon 1 is a constant parameter used for softening the angle signal, and T is a constant integral parameter;
S32, according to the pitch angle error e and the pitch angle Solving the self-adaptive increase rate c 2d of the error and pitch angle hinge factor, and integrating to obtain an error and pitch angle hinge self-adaptive signal c 2:
c2(n+1)＝c2(n)+c2dT
Wherein, k 3 and k 4 are constant parameters for adjusting the self-adaptive signal growth rate of the error and pitch angle hinge;
S33, solving the adaptive growth rate c 3d of the error and pitch rate hinge factor according to the pitch angle error e and the pitch rate omega z, and integrating to obtain an adaptive signal c 3 of the error and the pitch rate hinge:
c3(n+1)＝c3(n)+c3dT
Wherein, k 5 and k 6 are constant parameters, which are used for adjusting the self-adaptive signal growth rate of the error and pitch angle rate hinge, and epsilon 2 is a constant parameter;
S34, according to the pitch angle error e, solving an error factor self-adaptive growth rate c 4d, and integrating to obtain an error factor self-adaptive signal c 4;
c4(n+1)＝c4(n)+c4dT
wherein, k 7 and k 8 are constant parameters for adjusting the error factor adaptive signal growth rate;
S35, signal superposition and comprehensive summarization are carried out, and error related item self-adaptive signals f c are obtained:
Preferably, the step S4 includes the following substeps:
s41, solving the self-adaptive growth rate c 5d of the angle rate and the speed pitching inclination angle hinge factor according to the pitch angle rate omega z and the speed pitching inclination angle theta, and integrating to obtain an angle rate and speed pitching inclination angle hinge self-adaptive signal c 5:
c5(n+1)＝c5(n)+c5dT
Wherein d 1 and d 2 are constant parameters for adjusting the rate of increase of the self-adaptive signals of the angular rate and the speed pitch tilt angle hinge;
s42, according to the pitch angle rate omega z and the pitch angle Solving the self-adaptive growth rate c 6d of the angle rate and pitch angle hinge factor, and integrating to obtain an angle rate and pitch angle hinge self-adaptive signal c 6:
c6(n+1)＝c6(n)+c6dT;
Wherein d 3 and d 4 are constant parameters for adjusting the rate of increase of the self-adaptive signal of the angle rate and pitch angle hinge;
S43, solving an angular rate factor self-adaptive growth rate c 7d according to the pitch angle rate omega z, and integrating to obtain an angular rate factor self-adaptive signal c 7:
c7(n+1)＝c7(n)+c7dT
Wherein d 5 and d 6 are constant parameters for adjusting the pitch rate adaptive signal growth rate;
S44, solving the self-adaptive growth rate c 8d of the angle rate and pitch angle error hinge according to the angle rate omega z and the pitch angle error c 1d, and integrating to obtain an angle rate and pitch angle error hinge self-adaptive signal c 8:
c8(n+1)＝c8(n)+c8dT
wherein d 7 and d 8 are constant parameters for adjusting the rate of increase of the self-adaptive signal of the angle rate and pitch angle error hinge;
S45, signal superposition and comprehensive summarization are carried out, and an angular rate related item self-adaptive signal f d is obtained:
Preferably, the pitch angle θ and the pitch angle are measured by measuring the speed And the pitch angle rate omega z, then, the speed pitch inclination angle theta, the pitch angle rate omega z and the uncertainty related to a pitch angle error e hinge are respectively subjected to self-adaptive compensation according to the pitch angle error e and the pitch angle rate omega z, self-adaptive signals related to the angle and the angle rate are respectively obtained, then, an equivalent control item is constructed through a model aerodynamic coefficient, and then, the dynamic characteristics of the pitch angle rate omega z are subjected to inversion design, and the pitch angle rate omega z, the pitch angle error e, the pitch angle error integral s 1 and the self-adaptive signals f d are combined to form a thrust steering total control quantity, so that the aircraft tracks and regulates the expected pitch angle.
Preferably, the pitch angle is measured by measuring the pitch angle rate omega z of the aircraftAnd the speed pitching inclination angle theta realize the self-adaptive control of the thrust steering of the aircraft, and specifically comprises the following steps:
the error and speed pitch tilt hinge adaptation signal c 1 is:
c1＝∫(k1 e+k2 e)dt
The error and pitch angle hinge adaptive signal c 2 is:
c2＝∫(k3 e+k4θ)dt
the angular rate and velocity pitch angle hinge adaptation signal c 5 is:
c5＝∫(k5θ+k6θv)dt
The thrust steering pitch total control signal ζ a is:
Xi a＝k7e+k8∫edt+k9 e+ adaptive signal
Where k 1、k2、k3、k4、k5 and k 6 are tuning parameters and k 7、k8、k9 is a control gain parameter.
Compared with the prior art, the invention has the following advantages:
1. the invention relates to an aircraft self-adaptive thrust steering control method based on attitude measurement,
1. And carrying out self-adaptive compensation on the disturbance and uncertainty related to various hinges of the speed pitch inclination angle, the pitch angle rate and the pitch angle error according to the pitch angle error signal and the pitch angle rate signal, and respectively obtaining self-adaptive signals related to the angle and the angular rate, thereby enabling the whole control system to be capable of self-adapting to the change of the environment and the aerodynamic parameters.
2. According to the attitude measurement-based aircraft self-adaptive thrust steering control method, the equivalent control item is constructed through the model aerodynamic coefficient, the known structural information of the model is reasonably and effectively utilized, and the unknown part and the interference part can be self-adaptively compensated.
3. According to the attitude measurement-based self-adaptive thrust steering control method for the aircraft, disclosed by the invention, the mode of matching error integration, proportion and pitch angle rate is adopted, the assistance of quality distance control is not needed, and the stable flight of the aircraft can be independently ensured when a quality distance control system fails.
4. According to the attitude measurement-based aircraft self-adaptive thrust steering control method, the problems of dynamic compensation and stabilization of stress stabilization in the vertical direction of pitching and the pitching inclination angle of speed do not need to be considered, so that the analysis and control logic and the process of the whole system are simple and effective.
Drawings
FIG. 1 is a flow chart of an aircraft adaptive thrust steering control method based on attitude measurements of the present invention;
FIG. 2 is a schematic diagram of a thrust steering apparatus according to the thrust steering control method of the present invention;
FIG. 3 is a graph of an aircraft pitch rate signal for an embodiment of a thrust steering control method of the present invention;
FIG. 4 is a graph of aircraft pitch angle for an embodiment of a thrust steering control method of the present invention;
FIG. 5 is a graph of aircraft speed pitch angle for an embodiment of a thrust steering control method of the present invention;
FIG. 6 is a graph of an aircraft pitch angle error signal for an embodiment of a thrust steering control method of the present invention;
FIG. 7 is a graph of integrated signal of pitch angle error of an aircraft for an embodiment of a thrust steering control method of the present invention;
FIG. 8 is a graph of an aircraft thrust vectoring total control signal for an embodiment of a thrust vectoring control method of the present invention.
Detailed Description
The present invention is described in detail with reference to the drawings, wherein the invention is described in detail.
The invention relates to an aircraft self-adaptive thrust steering control method based on attitude measurement, which is shown in fig. 1 and comprises the following steps:
S1, measuring pitch angle rate omega z and pitch angle of an aircraft The change curve of the speed pitching angle theta is shown in fig. 3-5.
As shown in fig. 4, the pitch angle of the aircraftRise to the desired value of 10 deg. around 1 s.
As shown in fig. 5, the pitch angle error e converges to 0 at around 1 s;
S2, mounting a thrust steering device at the axial tail part of the flying direction of the aircraft, wherein as shown in FIG. 2, the embodiment takes a circular aircraft as an example, the method provided by the invention does not require that the shape of the aircraft is limited to a circular shape, a dish shape or an oval shape, the thrust pitching swing angle is xi, and the expected pitch angle of the aircraft is set as xi according to the flying task In this embodiment, the expected pitch angle signal of the aircraft is 10 °, and compared with the pitch angle ω z of the aircraft, a pitch angle error e is obtained, the pitch angle error e curve is shown in fig. 6, and the pitch angle error integral s 1 is obtained by integration, specifically:
S3, solving an error related term self-adaptive signal, which comprises the following substeps:
S31, solving an error and speed pitching inclination angle hinge factor self-adaptive growth rate c 1d according to a pitching angle error e and a speed pitching inclination angle theta, and integrating to obtain an error and speed pitching inclination angle hinge self-adaptive signal c 1;
c1(n+1)＝c1(n)+c1dT
Wherein k 1、k2 is a constant parameter for adjusting the error and the rate of growth of the self-adaptive signal of the speed-pitching inclination angle hinge, k 1＝0.1,k2＝0.15,ε1 is a constant parameter for softening the angle signal, and T is a constant integral parameter;
S32, according to the pitch angle error e and the pitch angle Solving the self-adaptive increase rate c 2d of the error and pitch angle hinge factor, and integrating to obtain an error and pitch angle hinge self-adaptive signal c 2:
c2(n+1)＝c2(n)+c2dT
Wherein k 3 and k 4 are constant parameters for adjusting the error and pitch angle hinge adaptive signal growth rate, k 3＝0.18,k4 =0.25 in the present embodiment;
S33, solving the adaptive growth rate c 3d of the error and pitch rate hinge factor according to the pitch angle error e and the pitch rate omega z, and integrating to obtain an adaptive signal c 3 of the error and the pitch rate hinge:
c3(n+1)＝c3(n)+c3dT
Wherein k 5 and k 6 are constant parameters for adjusting the rate of increase of the adaptive signal of the error and pitch rate hinge, and k 5＝0.35,k6＝0.05,ε2 is a constant parameter in the embodiment;
S34, according to the pitch angle error e, solving an error factor self-adaptive growth rate c 4d, and integrating to obtain an error factor self-adaptive signal c 4;
c4(n+1)＝c4(n)+c4dT
where k 7 and k 8 are constant parameters for adjusting the rate of increase of the error factor adaptive signal, k 7＝0.03,k8 =0.08 in this embodiment.
S35, signal superposition and comprehensive summarization are carried out, and error related item self-adaptive signals f c are obtained:
s4, solving an angular rate related term self-adaptive signal, which comprises the following substeps:
s41, solving the self-adaptive growth rate c 5d of the angle rate and the speed pitching inclination angle hinge factor according to the pitch angle rate omega z and the speed pitching inclination angle theta, and integrating to obtain an angle rate and speed pitching inclination angle hinge self-adaptive signal c 5:
c5(n+1)＝c5(n)+c5dT
Wherein d 1 and d 2 are constant parameters for adjusting the rate of increase of the adaptive signal of the angular rate and the rate of pitch angle hinge, and d 1＝0.25,d2 =0.13 in the present embodiment.
S42, according to the pitch angle rate omega z and the pitch angleSolving the self-adaptive growth rate c 6d of the angle rate and pitch angle hinge factor, and integrating to obtain an angle rate and pitch angle hinge self-adaptive signal c 6:
c6(n+1)＝c6(n)+c6dT;
Wherein d 3 and d 4 are constant parameters for adjusting the rate of increase of the adaptive signal of the angular rate and the pitch angle hinge, and d 3＝0.6,d4 =0.04 is selected in this embodiment.
S43, solving an angular rate factor self-adaptive growth rate c 7d according to the pitch angle rate omega z, and integrating to obtain an angular rate factor self-adaptive signal c 7:
c7(n+1)＝c7(n)+c7dT
Where d 5 and d 6 are constant parameters for adjusting the pitch rate adaptive signal growth rate, d 5＝0.33,d6 =0.02 in this embodiment.
S44, solving the self-adaptive growth rate c 8d of the angle rate and pitch angle error hinge according to the angle rate omega z and the pitch angle error c 1d, and integrating to obtain an angle rate and pitch angle error hinge self-adaptive signal c 8:
c8(n+1)＝c8(n)+c8dT
Where d 7 and d 8 are constant parameters for adjusting the rate of increase of the adaptive signal of the angular rate and pitch angle error hinge, d 7＝0.05,d8 =0.1 in this embodiment.
S45, signal superposition and comprehensive summarization are carried out, and an angular rate related item self-adaptive signal f d is obtained:
s5, solving a total control signal of thrust steering and pitching of the aircraft to realize steering control, wherein the substeps comprise the following steps:
S51, pitching the dip angle theta and the pitch angle according to the speed of the aircraft And aerodynamic parameters of the aircraft, designing a thrust steering equivalent control signal f e of the aircraft:
Where a 24、az is the aerodynamic parameter of the aircraft, and in this embodiment, a 24＝829,az = -12.89 is obtained according to the wind tunnel test result of its external dimension.
The expression of the dynamics control model in the aircraft pitching channel control is as follows:
Δu=K1Δα+K1Δq
Where Δu is the elevator control input signal, Δα is the pitch angle deviation, Δq is the pitch angle rate deviation, and K 1、K1 is the control gain.
After receiving the pitch angle error e and the pitch angle rate omega z in the step S5, the pitch angle theta and the pitch angle of the speed are adjustedThe flight control parameters are compensated, and the aircraft is adapted to the environmental and aerodynamic parameter changes by adjusting the gain parameters, so that the aircraft control system can dynamically compensate the interference factors in real time.
S52, superposing a pitch angle error e, a pitch angle error integral S 1, a pitch angle rate omega z, an error related item self-adaptive signal f c and an angular rate related item self-adaptive signal f d to form a thrust steering pitch total control signal of the aircraft, and transmitting the thrust steering pitch total control signal to a thrust steering device to control the thrust pitch swing angle of an engine to be equal to the thrust steering pitch total control signal of the aircraft, wherein the method specifically comprises the following steps of:
ξa＝-l1ωz-l2e-l3s1+fc+fd-fe;
Where ζ a is the total thrust steering pitch control signal, l 1、l2、l3 is the control parameter, which are both constants, and in this embodiment l 1＝0.5,l1＝1.5,l3 =0.1.
And obtaining a total control signal of thrust steering and pitching, realizing the tracking of the expected pitch angle by the aircraft, and completing the pitch channel control task.
By measuring pitch angle rate omega z, pitch angle of aircraftAnd the speed pitching inclination angle theta realize the self-adaptive control of the thrust steering of the aircraft, and specifically comprises the following steps:
the error and speed pitch tilt hinge adaptation signal c 1 is:
c1＝∫(k1 e+k2 e)dt
The error and pitch angle hinge adaptive signal c 2 is:
c2＝∫(k3e+k4θ)dt
the angular rate and velocity pitch angle hinge adaptation signal c 5 is:
c5＝∫(k5θ+k6θv)dt
The thrust steering pitch total control signal ζ a is:
Xi a＝k7e+k8∫edt+k9 e+ adaptive signal
Where k 1、k2、k3、k4、k5 and k 6 are tuning parameters and k 7、k8、k9 is a control gain parameter.
First, by measuring the speed pitch angle θ, pitch angleAnd the pitch angle rate omega z, then pitch the tilt angle theta and the pitch angle according to the pitch angle error e and the pitch angle rate omega z The method comprises the steps of adaptively compensating uncertainty related to a pitch angle rate omega z and a pitch angle error e hinge, respectively obtaining adaptive signals related to the angle and the angle rate, constructing an equivalent control item through an aerodynamic coefficient of a model, adopting inversion design through dynamic characteristics of the pitch angle rate omega z, combining the pitch angle rate omega z, the pitch angle error e, a pitch angle error integral s 1 and an adaptive signal f d to form a thrust steering total control quantity, tracking the expected pitch angle of the aircraft, and completing a pitch channel control task.
As shown in fig. 8, the total control signal of thrust steering pitch is smoothly stabilized at 0, the maximum value of the whole process is not more than 4, no obvious flutter and shake are generated, the control signal can meet engineering requirements, and the whole system response has very good stability and rapidity.
The invention discloses an aircraft self-adaptive thrust steering control method based on attitude measurement, which can adaptively compensate the pitching angle, the angular rate and the pitching angle error of a speed according to a pitching angle error signal and a pitching angle rate signal, construct an equivalent control item through a model aerodynamic coefficient, and combine the pitching angle rate, the pitching angle error, a pitching angle error integral and a self-adaptive signal to form a thrust steering total control signal so as to realize the regulation and control of an aircraft on an expected pitching angle.
In addition, according to the specific characteristics and flight task requirements of the aircraft, the self-adaptive control is optimized by adjusting the flight control parameters so as to adapt to different flight requirements of each stage, and error information can be accumulated by the control signals, so that the control strategy is gradually adjusted in the long-term flight process so as to adapt to the change trend of the attitude of the aircraft, and the integral term can continuously act when facing external interference, so that the aircraft is restored to the expected flight state.
The above examples are only illustrative of the preferred embodiments of the present invention and are not intended to limit the scope of the present invention, and various modifications and improvements made by those skilled in the art to the technical solution of the present invention should fall within the scope of protection defined by the claims of the present invention without departing from the spirit of the present invention.

Claims (5)

1.一种基于姿态测量的飞行器自适应推力转向控制方法，其特征在于，其包括以下步骤：1. An aircraft adaptive thrust steering control method based on attitude measurement, characterized in that it comprises the following steps: S1、测量飞行器的俯仰角速率ωz、俯仰角和速度俯仰倾角θ；S1. Measure the pitch rate ω z and pitch angle of the aircraft and velocity pitch angle θ; S2、在飞行器飞行方向的轴向尾部安装推力转向装置，推力转向装置的推力俯仰摆角为ξ，设置飞行器的期望俯仰角为与飞行器的俯仰角ωz比较，得到俯仰角误差e，并积分得到飞行器的的俯仰角误差积分s1，具体为：S2. Install a thrust steering device at the axial tail of the aircraft in the flight direction. The thrust pitch swing angle of the thrust steering device is ξ. Set the desired pitch angle of the aircraft to Compared with the pitch angle ω z of the aircraft, the pitch angle error e is obtained, and the pitch angle error integral s 1 of the aircraft is obtained by integration, which is specifically: S3、求解误差相关项自适应信号；S3, solving the error-related adaptive signal; S4、求解角速率相关项自适应信号：S4. Solve the adaptive signal of angular rate related items: S5、求解飞行器推力转向俯仰总控制信号，实现转向控制，其子步骤包括：S5, solving the total control signal of the thrust steering pitch of the aircraft to achieve steering control, the sub-steps of which include: S51、根据飞行器的速度俯仰倾角θ与俯仰角以及飞行器的空气动力学参数，设计飞行器的推力转向等效控制信号fe：S51, according to the speed of the aircraft pitch angle θ and the pitch angle As well as the aerodynamic parameters of the aircraft, the thrust steering equivalent control signal fe of the aircraft is designed: 式中，a24、az分别为飞行器的空气动力学参数，均为常数；Where a 24 and a z are the aerodynamic parameters of the aircraft, both of which are constants; 飞行器俯仰通道控制中动力学控制模型的表达式为：The expression of the dynamic control model in the pitch channel control of the aircraft is: Δu＝K1Δα+K1ΔqΔu＝K 1 Δα+K 1 Δq 式中，Δu为升降舵的控制输入信号，Δα为俯仰角偏差，Δq为俯仰角速度偏差，K1、K1为控制增益；Where, Δu is the control input signal of the elevator, Δα is the pitch angle deviation, Δq is the pitch angle velocity deviation, K 1 and K 1 are control gains; 接收到俯仰角误差e和俯仰角速率ωz后，对速度俯仰倾角θ、俯仰角飞行控制参数进行补偿，通过调整控制增益参数使飞行器适应环境和气动参数变化，从而使飞行器控制系统能实时动态补偿干扰因素；After receiving the pitch angle error e and the pitch angle rate ω z , the velocity pitch angle θ and the pitch angle Flight control parameters are compensated, and the aircraft is adapted to the environment and aerodynamic parameter changes by adjusting the control gain parameters, so that the aircraft control system can dynamically compensate for interference factors in real time; S52、叠加俯仰角误差e、俯仰角误差积分s1、以及俯仰角速率ωz、误差相关项自适应信号fc与角速率相关项自适应信号fd，形成飞行器的推力转向俯仰总控制信号，并传送至推力转向装置，控制发动机的推力俯仰摆角等于飞行器的推力转向俯仰总控制信号ξa，具体为：S52, superimposing the pitch angle error e, the pitch angle error integral s 1 , the pitch angle rate ω z , the error-related adaptive signal f c and the angular rate-related adaptive signal f d , forming a thrust steering pitch total control signal of the aircraft, and transmitting it to the thrust steering device to control the thrust pitch swing angle of the engine to be equal to the thrust steering pitch total control signal ξ a of the aircraft, specifically: ξa＝-l1ωz-l2e-l3s1+fc+fd-fe；ξ a =-l 1 ω z -l 2 el 3 s 1 +f c +f d -f e ; 式中，l1、l2、l3为控制参数，均为常数；In the formula, l 1 , l 2 , l 3 are control parameters, all of which are constants; 通过模型空气动力学系数构建等效控制项，结合俯仰角速率及俯仰角误差、俯仰角误差积分、自适应信号组成推力转向总控制信号，实现飞行器对期望俯仰角的调控。The equivalent control item is constructed by the model aerodynamic coefficient, and the total thrust steering control signal is composed of the pitch angle rate and pitch angle error, the pitch angle error integral and the adaptive signal to realize the control of the aircraft to the desired pitch angle. 2.根据权利要求1所述的基于姿态测量的飞行器自适应推力转向控制方法，其特征在于，所述步骤S3包括以下子步骤：2. The method for controlling an aircraft's adaptive thrust steering based on attitude measurement according to claim 1, wherein step S3 comprises the following sub-steps: S31、根据俯仰角误差e与速度俯仰倾角θ，求解俯仰角误差与速度俯仰倾角铰链因子自适应增长速率c1d，并积分得到误差与速度俯仰倾角铰链自适应信号c1；S31, according to the pitch angle error e and the velocity pitch tilt angle θ, solve the pitch angle error and the velocity pitch tilt angle hinge factor adaptive growth rate c 1d , and integrate the error and velocity pitch tilt angle hinge adaptive signal c 1 ; c1(n+1)＝c1(n)+c1dTc 1 (n+1)＝c 1 (n)+c 1d T 式中，k1、k2为常值参数，用于调节误差与速度俯仰倾角铰链自适应信号增长速率，ε1为常值参数，用于对角度信号柔化处理，T为常值积分参数；Wherein, k 1 and k 2 are constant parameters used to adjust the error and velocity pitch angle hinge adaptive signal growth rate, ε 1 is a constant parameter used to soften the angle signal, and T is a constant integral parameter; S32、根据俯仰角误差e与俯仰角求解误差与俯仰角铰链因子自适应增长速率c2d，并积分得到误差与俯仰角铰链自适应信号c2：S32, according to the pitch angle error e and the pitch angle Solve the error and pitch angle hinge factor adaptive growth rate c 2d , and integrate the error and pitch angle hinge adaptive signal c 2 : c2(n+1)＝c2(n)+c2dTc 2 (n+1)＝c 2 (n)+c 2d T 式中，k3与k4为常值参数，用于调节误差与俯仰角铰链自适应信号增长速率；Where k 3 and k 4 are constant parameters, which are used to adjust the error and pitch angle hinge adaptive signal growth rate; S33、根据俯仰角误差e与俯仰角速率ωz，求解误差与俯仰角速率铰链因子自适应增长速率c3d，并积分得到误差与俯仰角速率铰链自适应信号c3：S33, according to the pitch angle error e and the pitch angle rate ω z , solve the error and pitch angle rate hinge factor adaptive growth rate c 3d , and integrate the error and pitch angle rate hinge adaptive signal c 3 : c3(n+1)＝c3(n)+c3dTc 3 (n+1)＝c 3 (n)+c 3d T 式中，k5与k6为常值参数，用于调节误差与俯仰角速率铰链自适应信号增长速率，ε2为常值参数；Where, k 5 and k 6 are constant parameters, which are used to adjust the error and pitch angle rate hinge adaptive signal growth rate, and ε 2 is a constant parameter; S34、根据俯仰角误差e，求解误差因子自适应增长速率c4d，并积分得到误差因子自适应信号c4；S34, according to the pitch angle error e, solving the error factor adaptive growth rate c 4d , and integrating to obtain the error factor adaptive signal c 4 ; c4(n+1)＝c4(n)+c4dTc 4 (n+1)＝c 4 (n)+c 4d T 式中，k7与k8为常值参数，用于调节误差因子自适应信号增长速率；Where k7 and k8 are constant parameters used to adjust the error factor adaptive signal growth rate; S35、信号叠加与综合汇总，得到误差相关项自适应信号fc：S35, signal superposition and comprehensive summary, to obtain the error-related adaptive signal f c : 3.根据权利要求1所述的基于姿态测量的飞行器自适应推力转向控制方法，其特征在于，所述步骤S4包括以下子步骤：3. The method for controlling an aircraft's adaptive thrust steering based on attitude measurement according to claim 1, wherein step S4 comprises the following sub-steps: S41、根据俯仰角速率ωz与速度俯仰倾角θ，求解角速率与速度俯仰倾角铰链因子自适应增长速率c5d，并积分得到角速率与速度俯仰倾角铰链自适应信号c5：S41, according to the pitch angular rate ω z and the velocity pitch tilt angle θ, solve the angular rate and velocity pitch tilt angle hinge factor adaptive growth rate c 5d , and integrate to obtain the angular rate and velocity pitch tilt angle hinge adaptive signal c 5 : c5(n+1)＝c5(n)+c5dTc 5 (n+1)＝c 5 (n)+c 5d T 式中：d1与d2为常值参数，用于调节角速率与速度俯仰倾角铰链自适应信号增长速率；Where: d1 and d2 are constant parameters, which are used to adjust the growth rate of the angular rate and velocity pitch angle hinge adaptive signal; S42、根据俯仰角速率ωz与俯仰角求解角速率与俯仰角铰链因子自适应增长速率c6d，并积分得到角速率与俯仰角铰链自适应信号c6：S42, according to the pitch angle rate ω z and the pitch angle Solve the angular rate and pitch angle hinge factor adaptive growth rate c 6d , and integrate to obtain the angular rate and pitch angle hinge adaptive signal c 6 : c6(n+1)＝c6(n)+c6dT； c6 (n+1)= c6 (n)+ c6dT ; 式中，d3与d4为常值参数，用于调节角速率与俯仰角铰链自适应信号增长速率；Where d3 and d4 are constant parameters, which are used to adjust the growth rate of the angular rate and pitch angle hinge adaptive signal; S43、根据俯仰角速率ωz，求解角速率因子自适应增长速率c7d，并积分得到角速率因子自适应信号c7：S43, according to the pitch angle rate ω z , solve the angular rate factor adaptive growth rate c 7d , and integrate to obtain the angular rate factor adaptive signal c 7 : c7(n+1)＝c7(n)+c7dTc 7 (n+1)＝c 7 (n)+c 7d T 式中，d5与d6为常值参数，用于调节俯仰角速率自适应信号增长速率；Wherein, d5 and d6 are constant parameters used to adjust the growth rate of the pitch angle rate adaptive signal; S44、根据俯仰角速率ωz与俯仰角误差c1d，求解角速率与俯仰角误差铰链自适应增长速率c8d，并积分得到角速率与俯仰角误差铰链自适应信号c8：S44, according to the pitch angle rate ω z and the pitch angle error c 1d , solve the angular rate and pitch angle error hinge adaptive growth rate c 8d , and integrate to obtain the angular rate and pitch angle error hinge adaptive signal c 8 : c8(n+1)＝c8(n)+c8dTc 8 (n+1)＝c 8 (n)+c 8d T 式中，d7与d8为常值参数，用于调节角速率与俯仰角误差铰链自适应信号增长速率；Where d7 and d8 are constant parameters, which are used to adjust the growth rate of the angular rate and pitch angle error hinge adaptive signal; S45、信号叠加与综合汇总，得到角速率相关项自适应信号fd：S45, superimpose and summarize the signals to obtain the angular rate related adaptive signal f d : 4.根据权利要求1所述的基于姿态测量的飞行器自适应推力转向控制方法，其特征在于，通过测量飞行器的俯仰角速率ωz、俯仰角和速度俯仰倾角θ实现对飞行器推力转向的自适应控制，具体为：4. The method for controlling the aircraft's adaptive thrust steering based on attitude measurement according to claim 1 is characterized in that the pitch angle rate ω z and pitch angle The adaptive control of the aircraft thrust steering is realized by using the velocity pitch angle θ, specifically: 误差与速度俯仰倾角铰链自适应信号c1为：The error and velocity pitch angle hinge adaptive signal c1 is: c1＝∫(k1e+k2e)dtc 1 =∫(k 1 e+k 2 e)dt 误差与俯仰角铰链自适应信号c2为：The error and pitch angle hinge adaptive signal c2 are: c2＝∫(k3e+k4θ)dtc 2 =∫(k 3 e+k 4 θ)dt 角速率与速度俯仰倾角铰链自适应信号c5为：The angular rate and velocity pitch angle hinge adaptive signal c5 is: c5＝∫(k5θ+k6θv)dtc 5 =∫(k 5 θ+k 6 θv)dt 推力转向俯仰总控制信号ξa为：The total thrust steering pitch control signal ξa is: ξa＝k7e+k8∫edt+k9e+自适应信号ξ a = k 7 e + k 8 ∫edt + k 9 e + adaptive signal 式中，k1、k2、k3、k4、k5和k6为调节参数，k7、k8、k9为控制增益参数。Wherein, k 1 , k 2 , k 3 , k 4 , k 5 and k 6 are adjustment parameters, and k 7 , k 8 and k 9 are control gain parameters. 5.根据权利要求4所述的基于姿态测量的飞行器自适应推力转向控制方法，其特征在于，通过测量速度俯仰倾角θ、俯仰角与俯仰角速率ωz，然后分别根据俯仰角误差e以及俯仰角速率ωz对速度俯仰倾角θ、俯仰角俯仰角速率ωz以及俯仰角误差e铰链相关的不确定性进行自适应补偿，分别得到角度与角速率相关的自适应信号，然后通过模型空气动力学系数构造等效控制项，再通过俯仰角速率ωz的动力学动态特性采用反演设计，并结合俯仰角速率ωz以及俯仰角误差e、俯仰角误差积分s1、自适应信号fd组成推力转向总控制量，实现飞行器对期望俯仰角的跟踪与调控。5. The method for adaptive thrust steering control of an aircraft based on attitude measurement according to claim 4 is characterized in that by measuring the velocity pitch angle θ, the pitch angle and pitch angle rate ω z , and then the velocity pitch angle θ and pitch angle are calculated according to the pitch angle error e and the pitch angle rate ω z. The uncertainties related to the hinge of pitch angle rate ωz and pitch angle error e are adaptively compensated to obtain adaptive signals related to angle and angular rate respectively. Then, the equivalent control term is constructed through the aerodynamic coefficient of the model. Then, the inversion design is adopted through the dynamic characteristics of the pitch angle rate ωz . The total thrust steering control quantity is composed of the pitch angle rate ωz , the pitch angle error e, the pitch angle error integral s1 and the adaptive signal fd to realize the tracking and regulation of the desired pitch angle of the aircraft.
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Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20070156316A1 (en) * 2005-12-06 2007-07-05 Kabushiki Kaisha Toyota Chuo Kenkyusho Vehicle physical quantity estimation apparatus and storage medium having stored thereon computer program relating to the apparatus
RU2015149473A (en) * 2015-11-18 2017-05-19 Акционерное общество "Концерн радиостроения "Вега" AUTOMATED MODAL CONTROL SYSTEM IN THE LONGITUDINAL CHANNEL OF AIRCRAFT
CN111459184A (en) * 2020-04-15 2020-07-28 烟台南山学院 Unmanned aerial vehicle automatic carrier landing control method adopting segmented attack angle instruction
CN116627152A (en) * 2023-05-19 2023-08-22 重庆人文科技学院 Thrust vector and mass distance inversion composite control method for disc-shaped aircraft

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20070156316A1 (en) * 2005-12-06 2007-07-05 Kabushiki Kaisha Toyota Chuo Kenkyusho Vehicle physical quantity estimation apparatus and storage medium having stored thereon computer program relating to the apparatus
RU2015149473A (en) * 2015-11-18 2017-05-19 Акционерное общество "Концерн радиостроения "Вега" AUTOMATED MODAL CONTROL SYSTEM IN THE LONGITUDINAL CHANNEL OF AIRCRAFT
CN111459184A (en) * 2020-04-15 2020-07-28 烟台南山学院 Unmanned aerial vehicle automatic carrier landing control method adopting segmented attack angle instruction
CN116627152A (en) * 2023-05-19 2023-08-22 重庆人文科技学院 Thrust vector and mass distance inversion composite control method for disc-shaped aircraft

Non-Patent Citations (2)

* Cited by examiner, † Cited by third party
Title
YUSEN WANG, MEIJU ZHANG: "Thin Film Heat Flux Sensor With Sandwich Structure for Extreme Temperature Environment", 《IEEE TRANSACTIONS ON ELECTRON DEVICES》, 4 September 2023 (2023-09-04) *
怀思然;甘志超;陈亭宇;王赫;阎炎: "毫米波收发组件微组装激光三维扫描技术研究", 《测控技术》, 24 August 2020 (2020-08-24) *

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