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CN113126494B - Low-altitude flight pneumatic identification control method with reference track dynamically corrected - Google Patents

Low-altitude flight pneumatic identification control method with reference track dynamically corrected Download PDF

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CN113126494B
CN113126494B CN202110402167.0A CN202110402167A CN113126494B CN 113126494 B CN113126494 B CN 113126494B CN 202110402167 A CN202110402167 A CN 202110402167A CN 113126494 B CN113126494 B CN 113126494B
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许斌
王霞
孙绍山
陶呈纲
胡逸雯
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AVIC Chengdu Aircraft Design and Research Institute
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Abstract

本发明涉及一种参考轨迹动态修正的低空飞行气动辨识控制方法，属于飞行器控制领域。该方法首先设计了自回归模型对海浪高度进行预测，给出了引入海浪高度动态预测信息的飞行高度参考指令。然后将飞行器纵向动力学模型转换为速度子系统和高度子系统，针对速度子系统设计动态逆控制；针对高度子系统设计反步法控制，其中考虑海浪引起的气动特性未知问题，将气动函数转换为已知状态向量和未知气动参数向量相乘的线性参数化形式，设计自适应估计律对未知气动参数进行估计。该发明通过考虑海浪影响设计参考轨迹动态修正的气动辨识控制方法，实现了高度精细控制，对超低空掠海突防具有重要意义。

Figure 202110402167

The invention relates to a low-altitude flight aerodynamic identification control method with reference trajectory dynamic correction, belonging to the field of aircraft control. In this method, an autoregressive model is first designed to predict the wave height, and a flight height reference instruction that introduces the dynamic prediction information of the wave height is given. Then, the longitudinal dynamics model of the aircraft is converted into the velocity subsystem and the altitude subsystem, and the dynamic inverse control is designed for the velocity subsystem; the backstepping control is designed for the altitude subsystem, in which the unknown aerodynamic characteristics caused by ocean waves are considered, and the aerodynamic function is converted into An adaptive estimation law is designed to estimate the unknown aerodynamic parameters, which is a linear parameterized form in which the known state vector and the unknown aerodynamic parameter vector are multiplied. The invention realizes highly precise control by considering the influence of ocean waves to design an aerodynamic identification control method for dynamic correction of reference trajectories, which is of great significance for ultra-low-altitude sea skimming penetration.

Figure 202110402167

Description

参考轨迹动态修正的低空飞行气动辨识控制方法Low-altitude flight aerodynamic identification control method based on dynamic correction of reference trajectory

技术领域technical field

本发明涉及一种飞行器控制方法，特别是涉及一种参考轨迹动态修正的低空飞行气动辨识控制方法，属于飞行器控制领域。The invention relates to an aircraft control method, in particular to a low-altitude flight aerodynamic identification control method with dynamic correction of reference trajectory, belonging to the field of aircraft control.

背景技术Background technique

飞行器超低空掠海飞行可以借助海面背景掩盖自身散射特征，有效避开船舰和机载雷达，实现对敌方海上船舰目标实施突袭，大大增强其生存和突防能力。海浪是影响超低空掠海飞行的主要环境因素，它一方面直接通过波浪频繁波动干扰飞行高度，另一方面通过影响气动特性影响高度控制精度，进而降低生存能力和低空突防效能。因此，对海浪引起的效应进行稳定补偿是实现飞行器稳定不击水掠海飞行的关键。The ultra-low-altitude sea-skimming flight of the aircraft can cover its own scattering characteristics with the help of the sea surface background, effectively avoid ships and airborne radars, and realize surprise attacks on enemy ships and ships at sea, greatly enhancing their survival and penetration capabilities. Ocean waves are the main environmental factors affecting ultra-low-altitude sea-skimming flight. On the one hand, it directly interferes with the flight altitude through frequent wave fluctuations, and on the other hand, it affects the altitude control accuracy by affecting the aerodynamic characteristics, thereby reducing the survivability and low-altitude penetration efficiency. Therefore, stably compensating for the effects caused by ocean waves is the key to realize the stable flight of the aircraft without hitting the water and skimming the sea.

《低空掠海飞行拖靶自抗扰高度控制律设计》(方晓星,王勇,王英勋,南京理工大学学报,2012,36(05):835-839+845.)一文针对海浪波动引起的建模误差和外界扰动提出了自抗扰高度控制律设计方法。采用扩张状态观测器对未建模动态和干扰作用进行估计与补偿，引入微分反馈解决超调与快速性的矛盾。该方法将海浪影响建模为广泛扰动，没有对海浪影响进行针对性补偿，无法验证实际应用中的有效性。The article "Design of Active Disturbance Rejection Height Control Law for Low-altitude Sea-skimming Flying Targets" (Fang Xiaoxing, Wang Yong, Wang Yingxun, Journal of Nanjing University of Science and Technology, 2012, 36(05): 835-839+845.) is aimed at modeling caused by wave fluctuations Errors and external disturbances, a design method of ADRC height control law is proposed. The extended state observer is used to estimate and compensate the unmodeled dynamics and disturbance effects, and the differential feedback is introduced to solve the contradiction between overshoot and rapidity. This method models the impact of waves as a wide range of disturbances, does not compensate for the impact of waves in a targeted manner, and cannot verify the effectiveness in practical applications.

发明内容SUMMARY OF THE INVENTION

要解决的技术问题technical problem to be solved

针对飞行器超低空掠海飞行中海浪对高度干扰影响以及气动参数不确定问题，本发明设计了一种参考轨迹动态修正的低空飞行气动辨识控制方法。Aiming at the influence of sea waves on altitude interference and the uncertainty of aerodynamic parameters in the ultra-low-altitude sea-skimming flight of the aircraft, the present invention designs a low-altitude flight aerodynamic identification control method with dynamic correction of reference trajectory.

技术方案Technical solutions

一种参考轨迹动态修正的低空飞行气动辨识控制方法，其特征在于步骤如下：A low-altitude flight aerodynamic identification control method with reference trajectory dynamic correction, characterized in that the steps are as follows:

步骤1：考虑飞行器纵向通道动力学模型：Step 1: Consider the aircraft longitudinal channel dynamics model:

Figure BDA0003020811040000011
Figure BDA0003020811040000011

Figure BDA0003020811040000021
Figure BDA0003020811040000021

Figure BDA0003020811040000022
Figure BDA0003020811040000022

Figure BDA0003020811040000023
Figure BDA0003020811040000023

Figure BDA0003020811040000024
Figure BDA0003020811040000024

所述的运动学模型由五个状态量X＝[V，h，γ，α，q]T和两个控制输入U＝[δe，T]T组成；V表示速度，h表示高度，γ表示航迹角，α表示攻角，q表示俯仰角速度，δe表示舵偏角，T表示推力；m、Iyy和g分别表示质量、俯仰轴的转动惯量和重力引起的加速度；The kinematic model described is composed of five state quantities X=[V, h, γ, α, q] T and two control inputs U=[δ e , T] T ; V represents speed, h represents height, and γ represents the track angle, α represents the angle of attack, q represents the pitch angular velocity, δ e represents the rudder deflection angle, and T represents the thrust; m, I yy and g represent the mass, the moment of inertia of the pitch axis and the acceleration caused by gravity, respectively;

力、力矩以及各系数的表达式为：The expressions of force, moment and each coefficient are:

Figure BDA0003020811040000025
Figure BDA0003020811040000025

其中，Q＝(1/2)ρV2表示动压，Sω表示气动参考面积，cA表示平均气动弦长，CL0、CLα、CD0、CDα、

Figure BDA0003020811040000026
Cm0、Cmα和
Figure BDA0003020811040000027
均表示气动参数；Among them, Q=(1/2)ρV 2 represents the dynamic pressure, S ω represents the aerodynamic reference area, c A represents the average aerodynamic chord length, C L0 , C Lα , C D0 , C Dα ,
Figure BDA0003020811040000026
C m0 , C mα and
Figure BDA0003020811040000027
Both represent aerodynamic parameters;

步骤2：自回归模型的最小二乘形式为Step 2: The least squares form of the autoregressive model is

ζ(k)＝ψT(k)θw (7)ζ(k)= ψT (k) θw (7)

ψ(k)＝[-ζ(k-1)，...，-ζ(k-nw)]T (8)ψ(k)=[-ζ(k-1),...,-ζ(kn w )] T (8)

θw＝[d1，...，dw]T (9)θ w = [d 1 , . . . , d w ] T (9)

其中，ζ(k)表示第k个时刻的海浪高度，θw表示未知参数，nw表示辨识阶次，由设计者给出；Among them, ζ(k) represents the wave height at the kth moment, θw represents the unknown parameter, and nw represents the identification order, which is given by the designer;

采用带遗忘因子的递推最小二乘法对θw进行估计Estimation of θ w using recursive least squares with forgetting factor

Figure BDA0003020811040000028
Figure BDA0003020811040000028

Figure BDA0003020811040000031
Figure BDA0003020811040000031

Figure BDA0003020811040000032
Figure BDA0003020811040000032

其中，μ表示遗忘因子，由设计者给出；Among them, μ represents the forgetting factor, which is given by the designer;

步骤3：设计高度参考指令hd为Step 3: Design the height reference instruction h d as

Figure BDA0003020811040000033
Figure BDA0003020811040000033

其中，xg表示飞行器重心沿地面坐标系x轴的投影，满足

Figure BDA0003020811040000034
Z2表示飞行器直线下滑时的初始高度，由设计者给出；
Figure BDA0003020811040000035
Z1表示飞行器末端拉平时的初始高度，由设计者给出；
Figure BDA0003020811040000036
Z0表示低空掠海段的初始高度，由设计者给出；Among them, x g represents the projection of the center of gravity of the aircraft along the x-axis of the ground coordinate system, satisfying
Figure BDA0003020811040000034
Z 2 represents the initial height when the aircraft slides straight down, which is given by the designer;
Figure BDA0003020811040000035
Z 1 represents the initial height when the end of the aircraft is pulled out, which is given by the designer;
Figure BDA0003020811040000036
Z 0 represents the initial height of the low-altitude sea-skimming section, which is given by the designer;

步骤4：将飞行器动力学模型解耦得到速度子系统(1)和高度子系统(2)-(5)；Step 4: Decouple the aircraft dynamics model to obtain the velocity subsystem (1) and the altitude subsystem (2)-(5);

速度子系统(1)写为The velocity subsystem (1) is written as

Figure BDA0003020811040000037
Figure BDA0003020811040000037

式中，

Figure BDA0003020811040000038
Figure BDA0003020811040000039
In the formula,
Figure BDA0003020811040000038
Figure BDA0003020811040000039

取x1＝h，x2＝γ，x3＝θ，x4＝q，其中θ＝α+γ表示俯仰角，高度子系统(2)-(5)写为Take x 1 =h, x 2 =γ, x 3 =θ, x 4 =q, where θ=α+γ represents the pitch angle, and the altitude subsystems (2)-(5) are written as

Figure BDA00030208110400000310
Figure BDA00030208110400000310

式中，

Figure BDA00030208110400000311
Figure BDA0003020811040000041
Figure BDA0003020811040000042
Figure BDA0003020811040000043
In the formula,
Figure BDA00030208110400000311
Figure BDA0003020811040000041
Figure BDA0003020811040000042
Figure BDA0003020811040000043

步骤5：对于速度子系统，定义速度跟踪误差

Figure BDA0003020811040000044
设计控制输入T为Step 5: For the velocity subsystem, define the velocity tracking error
Figure BDA0003020811040000044
The design control input T is

Figure BDA0003020811040000045
Figure BDA0003020811040000045

式中，Vd为速度参考指令，kv＞0为控制参数，

Figure BDA0003020811040000046
In the formula, V d is the speed reference command, k v > 0 is the control parameter,
Figure BDA0003020811040000046

设计参数自适应估计律

Figure BDA0003020811040000047
为Design Parameter Adaptive Estimation Law
Figure BDA0003020811040000047
for

Figure BDA0003020811040000048
Figure BDA0003020811040000048

式中，γv＞0为控制参数；In the formula, γ v > 0 is the control parameter;

步骤6：定义高度跟踪误差

Figure BDA0003020811040000049
设计虚拟控制器
Figure BDA00030208110400000410
为Step 6: Define Altitude Tracking Error
Figure BDA0003020811040000049
Design a virtual controller
Figure BDA00030208110400000410
for

Figure BDA00030208110400000411
Figure BDA00030208110400000411

式中，k1＞0为控制参数；In the formula, k 1 > 0 is the control parameter;

引入一阶滤波器Introduce a first-order filter

Figure BDA00030208110400000412
Figure BDA00030208110400000412

式中，α2＞0为控制参数；In the formula, α 2 > 0 is the control parameter;

定义航迹角跟踪误差

Figure BDA00030208110400000413
设计虚拟控制器
Figure BDA00030208110400000414
为Define track angle tracking error
Figure BDA00030208110400000413
Design a virtual controller
Figure BDA00030208110400000414
for

Figure BDA00030208110400000415
Figure BDA00030208110400000415

式中，k2＞0为控制参数，

Figure BDA00030208110400000416
In the formula, k 2 >0 is the control parameter,
Figure BDA00030208110400000416

引入一阶滤波器Introduce a first-order filter

Figure BDA00030208110400000417
Figure BDA00030208110400000417

式中，α3＞0为控制参数；In the formula, α 3 > 0 is the control parameter;

设计参数自适应估计律

Figure BDA0003020811040000051
和
Figure BDA0003020811040000052
为Design Parameter Adaptive Estimation Law
Figure BDA0003020811040000051
and
Figure BDA0003020811040000052
for

Figure BDA0003020811040000053
Figure BDA0003020811040000053

式中，

Figure BDA0003020811040000054
和
Figure BDA0003020811040000055
为控制参数；In the formula,
Figure BDA0003020811040000054
and
Figure BDA0003020811040000055
is the control parameter;

定义俯仰角跟踪误差

Figure BDA0003020811040000056
设计虚拟控制器
Figure BDA0003020811040000057
为Define pitch tracking error
Figure BDA0003020811040000056
Design a virtual controller
Figure BDA0003020811040000057
for

Figure BDA0003020811040000058
Figure BDA0003020811040000058

式中，k3＞0为控制参数；In the formula, k 3 > 0 is the control parameter;

引入一阶滤波器Introduce a first-order filter

Figure BDA0003020811040000059
Figure BDA0003020811040000059

式中，α4＞0为控制参数；In the formula, α 4 > 0 is the control parameter;

设计俯仰角速度跟踪误差

Figure BDA00030208110400000510
设计控制输入δe为Design pitch rate tracking error
Figure BDA00030208110400000510
The design control input δ e is

Figure BDA00030208110400000511
Figure BDA00030208110400000511

式中，k4＞0为控制参数，

Figure BDA00030208110400000512
In the formula, k 4 > 0 is the control parameter,
Figure BDA00030208110400000512

设计参数自适应估计律

Figure BDA00030208110400000513
和
Figure BDA00030208110400000514
为Design Parameter Adaptive Estimation Law
Figure BDA00030208110400000513
and
Figure BDA00030208110400000514
for

Figure BDA00030208110400000515
Figure BDA00030208110400000515

式中，

Figure BDA00030208110400000516
和
Figure BDA00030208110400000517
为控制参数；In the formula,
Figure BDA00030208110400000516
and
Figure BDA00030208110400000517
is the control parameter;

步骤7：根据得到的推力T和舵偏角δe，返回到飞行器动力学模型(1)-(5)，对速度和高度进行跟踪控制。Step 7: According to the obtained thrust T and rudder deflection angle δ e , return to the aircraft dynamics model (1)-(5), and perform tracking control on the speed and altitude.

一种计算机系统，其特征在于包括：一个或多个处理器，计算机可读存储介质，用于存储一个或多个程序，其中，当所述一个或多个程序被所述一个或多个处理器执行时，使得所述一个或多个处理器实现上述的方法。A computer system, characterized by comprising: one or more processors, and a computer-readable storage medium for storing one or more programs, wherein when the one or more programs are processed by the one or more programs When the processor is executed, the one or more processors are caused to implement the above method.

一种计算机可读存储介质，其特征在于存储有计算机可执行指令，所述指令在被执行时用于实现上述的方法。A computer-readable storage medium is characterized in that computer-executable instructions are stored, and the instructions, when executed, are used to implement the above-mentioned method.

一种计算机程序，其特征在于包括计算机可执行指令，所述指令在被执行时用于实现上述的方法。A computer program characterized by comprising computer-executable instructions which, when executed, are used to implement the above-mentioned method.

有益效果beneficial effect

本发明提出的一种参考轨迹动态修正的低空飞行气动辨识控制方法，该方法首先设计了自回归模型对海浪高度进行预测，给出了引入海浪高度动态预测信息的飞行高度参考指令。然后将飞行器纵向动力学模型转换为速度子系统和高度子系统，针对速度子系统设计动态逆控制；针对高度子系统设计反步法控制，其中考虑海浪引起的气动特性未知问题，将气动函数转换为已知状态向量和未知气动参数向量相乘的线性参数化形式，设计自适应估计律对未知气动参数进行估计。该发明通过考虑海浪影响设计参考轨迹动态修正的气动辨识控制方法，实现了高度精细控制，对超低空掠海突防具有重要意义。The invention proposes a low-altitude flight aerodynamic identification control method with dynamic correction of reference trajectory. The method first designs an autoregressive model to predict the sea wave height, and provides a flight height reference command that introduces the dynamic prediction information of the sea wave height. Then, the longitudinal dynamics model of the aircraft is converted into the velocity subsystem and the altitude subsystem, and the dynamic inverse control is designed for the velocity subsystem; the backstepping control is designed for the altitude subsystem, which considers the unknown aerodynamic characteristics caused by ocean waves, and converts the aerodynamic function into An adaptive estimation law is designed to estimate the unknown aerodynamic parameters as a linear parametric form of multiplying the known state vector and the unknown aerodynamic parameter vector. The invention realizes highly precise control by considering the influence of ocean waves to design an aerodynamic identification control method for dynamic correction of reference trajectories, which is of great significance for ultra-low-altitude sea skimming penetration.

本发明与现有技术相比有益效果为：Compared with the prior art, the present invention has the following beneficial effects:

(1)本发明通过自回归模型实现了海浪高度预测，将环境动态预测信息引入到参考轨迹设计中，提升了超低空掠海飞行的自适应性能和实时响应特性。(1) The present invention realizes the wave height prediction through the autoregressive model, introduces the environmental dynamic prediction information into the reference trajectory design, and improves the self-adaptive performance and real-time response characteristics of ultra-low altitude sea-skimming flight.

(2)本发明考虑超低空飞行过程中复杂海效应造成的气动参数不确定问题，将气动函数转化为线性参数化形式，设计自适应更新律对未知气动参数向量进行估计。(2) The present invention considers the uncertainty of aerodynamic parameters caused by the complex sea effect during ultra-low altitude flight, converts the aerodynamic function into a linear parameterized form, and designs an adaptive update law to estimate the unknown aerodynamic parameter vector.

(3)本发明充分考虑海浪波动对飞行性能的影响，给出了参考轨迹动态修正的低空飞行气动辨识控制方法，实现了飞行器相对海平面高度的精细控制。(3) The present invention fully considers the influence of wave fluctuations on flight performance, and provides a low-altitude flight aerodynamic identification control method with dynamic correction of reference trajectory, and realizes fine control of the aircraft relative to sea level.

附图说明Description of drawings

附图仅用于示出具体实施例的目的，而并不认为是对本发明的限制，在整个附图中，相同的参考符号表示相同的部件。The drawings are for the purpose of illustrating specific embodiments only and are not to be considered limiting of the invention, and like reference numerals refer to like parts throughout the drawings.

图1本发明方法流程图Figure 1 is a flow chart of the method of the present invention

具体实施方式Detailed ways

为了使本发明的目的、技术方案及优点更加清楚明白，以下结合附图和实施例，对本发明进行进一步详细说明。应当理解，此处所描述的具体实施例仅仅用以解释本发明，并不用于限定本发明。此外，下面描述的本发明各个实施方式中所涉及到的技术特征只要彼此之间未构成冲突就可以相互组合。In order to make the objectives, technical solutions and advantages of the present invention clearer, the present invention will be further described in detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present invention, but not to limit the present invention. In addition, the technical features involved in the various embodiments of the present invention described below can be combined with each other as long as they do not conflict with each other.

参照图1，本发明一种参考轨迹动态修正的低空飞行气动辨识控制方法。具体步骤如下：Referring to FIG. 1, the present invention is a low-altitude flight aerodynamic identification control method with dynamic correction of reference trajectory. Specific steps are as follows:

(a)考虑飞行器纵向通道动力学模型：(a) Consider the aircraft longitudinal channel dynamics model:

Figure BDA0003020811040000071
Figure BDA0003020811040000071

Figure BDA0003020811040000072
Figure BDA0003020811040000072

Figure BDA0003020811040000073
Figure BDA0003020811040000073

Figure BDA0003020811040000074
Figure BDA0003020811040000074

Figure BDA0003020811040000075
Figure BDA0003020811040000075

所述的运动学模型由五个状态量X＝[V,h,γ,α,q]T和两个控制输入U＝[δe,T]T组成；V表示速度，h表示高度，γ表示航迹角，α表示攻角，q表示俯仰角速度，δe表示舵偏角，T表示推力；m、Iyy和g分别表示质量、俯仰轴的转动惯量和重力引起的加速度。The kinematic model described is composed of five state quantities X=[V,h,γ,α,q] T and two control inputs U=[δ e ,T] T ; V represents speed, h represents height, and γ represents the track angle, α represents the angle of attack, q represents the pitch angular velocity, δ e represents the rudder deflection angle, and T represents the thrust; m, I yy and g represent the mass, the moment of inertia of the pitch axis, and the acceleration caused by gravity, respectively.

力、力矩以及各系数的表达式为：The expressions of force, moment and each coefficient are:

Figure BDA0003020811040000076
Figure BDA0003020811040000076

其中，Q＝(1/2)ρV2表示动压，Sω＝1.1712表示气动参考面积，cA＝0.4118表示平均气动弦长，CL0＝0.1651、CLα＝4.5111、CD0＝0.0230、CDα＝0.0765、

Figure BDA0003020811040000081
Cm0＝-0.1539、Cmα＝-5.2369和
Figure BDA0003020811040000082
均表示气动参数。Among them, Q=(1/2)ρV 2 represents the dynamic pressure, S ω =1.1712 represents the aerodynamic reference area, c A =0.4118 represents the average aerodynamic chord length, C L0 =0.1651, C Lα =4.5111, C D0 =0.0230, C Dα = 0.0765,
Figure BDA0003020811040000081
C m0 = -0.1539, C mα = -5.2369 and
Figure BDA0003020811040000082
Both represent aerodynamic parameters.

(b)自回归模型的最小二乘形式为(b) The least squares form of the autoregressive model is

ζ(k)＝ψT(k)θw (7)ζ(k)= ψT (k) θw (7)

ψ(k)＝[-ζ(k-1)，...，-ζ(k-nw)]T (8)ψ(k)=[-ζ(k-1),...,-ζ(kn w )] T (8)

θw＝[d1，...，dw]T (9)θ w = [d 1 , . . . , d w ] T (9)

其中，ζ(k)表示第k个时刻的海浪高度，θw表示未知参数，nw＝4。Among them, ζ(k) represents the wave height at the k-th moment, θw represents the unknown parameter, and n w = 4.

采用带遗忘因子的递推最小二乘法对θw进行估计Estimation of θ w using recursive least squares with forgetting factor

Figure BDA0003020811040000083
Figure BDA0003020811040000083

Figure BDA0003020811040000084
Figure BDA0003020811040000084

Figure BDA0003020811040000085
Figure BDA0003020811040000085

其中，μ＝0.95。where μ=0.95.

(c)设计高度参考指令hd为(c) The design height reference command h d is

Figure BDA0003020811040000086
Figure BDA0003020811040000086

其中，xg表示飞行器重心沿地面坐标系x轴的投影，满足

Figure BDA0003020811040000087
Z2＝200m；
Figure BDA0003020811040000088
Z1＝120m；
Figure BDA0003020811040000089
Z0＝105m。Among them, x g represents the projection of the center of gravity of the aircraft along the x-axis of the ground coordinate system, satisfying
Figure BDA0003020811040000087
Z 2 =200m;
Figure BDA0003020811040000088
Z 1 =120m;
Figure BDA0003020811040000089
Z 0 =105m.

(d)将飞行器动力学模型解耦得到速度子系统(1)和高度子系统(2)-(5)。(d) Decoupling the aircraft dynamics model to obtain the velocity subsystem (1) and the altitude subsystems (2)-(5).

速度子系统(1)写为The velocity subsystem (1) is written as

Figure BDA00030208110400000810
Figure BDA00030208110400000810

式中，

Figure BDA0003020811040000091
Figure BDA0003020811040000092
In the formula,
Figure BDA0003020811040000091
Figure BDA0003020811040000092

取x1＝h，x2＝γ，x3＝θ，x4＝q，其中θ＝α+γ表示俯仰角，高度子系统(2)-(5)写为Take x 1 =h, x 2 =γ, x 3 =θ, x 4 =q, where θ=α+γ represents the pitch angle, and the altitude subsystems (2)-(5) are written as

Figure BDA0003020811040000093
Figure BDA0003020811040000093

式中，

Figure BDA0003020811040000094
Figure BDA0003020811040000095
Figure BDA0003020811040000096
Figure BDA0003020811040000097
In the formula,
Figure BDA0003020811040000094
Figure BDA0003020811040000095
Figure BDA0003020811040000096
Figure BDA0003020811040000097

(e)对于速度子系统，定义速度跟踪误差

Figure BDA0003020811040000098
设计控制输入T为(e) For the velocity subsystem, define the velocity tracking error
Figure BDA0003020811040000098
The design control input T is

Figure BDA0003020811040000099
Figure BDA0003020811040000099

式中，Vd为速度参考指令，kv＝5，

Figure BDA00030208110400000910
In the formula, V d is the speed reference command, k v =5,
Figure BDA00030208110400000910

设计参数自适应估计律

Figure BDA00030208110400000911
为Design Parameter Adaptive Estimation Law
Figure BDA00030208110400000911
for

Figure BDA00030208110400000912
Figure BDA00030208110400000912

式中，γv＝3。In the formula, γ v =3.

(f)定义高度跟踪误差

Figure BDA00030208110400000913
设计虚拟控制器
Figure BDA00030208110400000914
为(f) Define altitude tracking error
Figure BDA00030208110400000913
Design a virtual controller
Figure BDA00030208110400000914
for

Figure BDA00030208110400000915
Figure BDA00030208110400000915

式中，k1＝7。In the formula, k 1 =7.

引入一阶滤波器Introduce a first-order filter

Figure BDA0003020811040000101
Figure BDA0003020811040000101

式中，α2＝0.05。In the formula, α 2 =0.05.

定义航迹角跟踪误差

Figure BDA0003020811040000102
设计虚拟控制器
Figure BDA0003020811040000103
为Define track angle tracking error
Figure BDA0003020811040000102
Design a virtual controller
Figure BDA0003020811040000103
for

Figure BDA0003020811040000104
Figure BDA0003020811040000104

式中，k2＝10，

Figure BDA0003020811040000105
In the formula, k 2 =10,
Figure BDA0003020811040000105

引入一阶滤波器Introduce a first-order filter

Figure BDA0003020811040000106
Figure BDA0003020811040000106

式中，α3＝0.05。In the formula, α 3 =0.05.

设计参数自适应估计律

Figure BDA0003020811040000107
和
Figure BDA0003020811040000108
为Design Parameter Adaptive Estimation Law
Figure BDA0003020811040000107
and
Figure BDA0003020811040000108
for

Figure BDA0003020811040000109
Figure BDA0003020811040000109

式中，

Figure BDA00030208110400001010
和
Figure BDA00030208110400001011
In the formula,
Figure BDA00030208110400001010
and
Figure BDA00030208110400001011

定义俯仰角跟踪误差

Figure BDA00030208110400001012
设计虚拟控制器
Figure BDA00030208110400001013
为Define pitch tracking error
Figure BDA00030208110400001012
Design a virtual controller
Figure BDA00030208110400001013
for

Figure BDA00030208110400001014
Figure BDA00030208110400001014

式中，k3＝2。In the formula, k 3 =2.

引入一阶滤波器Introduce a first-order filter

Figure BDA00030208110400001015
Figure BDA00030208110400001015

式中，α4＝0.05。In the formula, α 4 =0.05.

设计俯仰角速度跟踪误差

Figure BDA00030208110400001016
设计控制输入δe为Design pitch rate tracking error
Figure BDA00030208110400001016
The design control input δ e is

Figure BDA00030208110400001017
Figure BDA00030208110400001017

式中，k4＝1，

Figure BDA00030208110400001018
In the formula, k 4 =1,
Figure BDA00030208110400001018

设计参数自适应估计律

Figure BDA00030208110400001019
和
Figure BDA00030208110400001020
为Design Parameter Adaptive Estimation Law
Figure BDA00030208110400001019
and
Figure BDA00030208110400001020
for

Figure BDA0003020811040000111
Figure BDA0003020811040000111

式中，

Figure BDA0003020811040000112
和
Figure BDA0003020811040000113
In the formula,
Figure BDA0003020811040000112
and
Figure BDA0003020811040000113

(g)根据得到的推力T和舵偏角δe，返回到飞行器动力学模型(1)-(5)，对速度和高度进行跟踪控制。(g) According to the obtained thrust T and rudder deflection angle δ e , return to the aircraft dynamics model (1)-(5), and perform tracking control on the speed and altitude.

以上所述，仅为本发明的具体实施方式，但本发明的保护范围并不局限于此，任何熟悉本技术领域的技术人员在本发明公开的技术范围内，可轻易想到各种等效的修改或替换，这些修改或替换都应涵盖在本发明的保护范围之内。The above are only specific embodiments of the present invention, but the protection scope of the present invention is not limited to this. Any person skilled in the art can easily think of various equivalents within the technical scope disclosed by the present invention. Modifications or substitutions should be included within the protection scope of the present invention.

Claims (3)

1. A low-altitude flight pneumatic identification control method with reference to track dynamic correction is characterized by comprising the following steps:
step 1: considering the aircraft longitudinal passage dynamics model:
Figure FDA0003643856020000011
Figure FDA0003643856020000012
Figure FDA0003643856020000013
Figure FDA0003643856020000014
Figure FDA0003643856020000015
the dynamic model consists of five state variables X ═ V, h, gamma, alpha and q] T And two control inputs U ═ δ e ,T] T Composition is carried out; v represents velocity, h represents altitude, γ represents track angle, α represents angle of attack, q represents pitch angle velocity, δ e Denotes rudder deflection angle, T denotes thrust; m, I yy And g represents mass, moment of inertia of pitch axis, and acceleration due to gravity, respectively;
the expressions for the force, moment and coefficients are:
Figure FDA0003643856020000016
wherein Q ═ p V (1/2) 2 Denotes dynamic pressure, S ω Denotes the aerodynamic reference area, c A Representing mean aerodynamic chord length，C L0 、C Lα 、C D0 、C Dα 、
Figure FDA0003643856020000017
C m0 、C mα And
Figure FDA0003643856020000018
all represent pneumatic parameters;
step 2: the least squares form of the autoregressive model is
ζ(k)＝ψ T (k)θ w (7)
ψ(k)＝[-ζ(k-1),...,-ζ(k-n w )] T (8)
θ w ＝[d 1 ,...,d w ] T (9)
Where ζ (k) represents the wave height at the k-th instant, θ w Representing an unknown parameter, n w Representing the recognition order, given by the designer;
theta is measured by adopting recursive least square method with forgetting factor w Make an estimation
Figure FDA0003643856020000021
Figure FDA0003643856020000022
Figure FDA0003643856020000023
Wherein μ represents a forgetting factor, given by the designer;
and step 3: design height reference instruction h d Is composed of
Figure FDA0003643856020000024
Wherein x is g Represents the projection of the gravity center of the aircraft along the x axis of a ground coordinate system, and satisfies
Figure FDA0003643856020000025
Z 2 The initial height of the aircraft during straight gliding is shown and is given by a designer;
Figure FDA0003643856020000026
Z 1 representing the initial height of the aircraft at the tail end when the aircraft is flat, and is given by a designer;
Figure FDA0003643856020000027
Z 0 representing the initial height of the low altitude sea-swept segment, given by the designer;
and 4, step 4: decoupling the aircraft dynamics model to obtain a speed subsystem (1) and altitude subsystems (2) - (5);
the speed subsystem (1) is written as
Figure FDA0003643856020000028
In the formula (I), the compound is shown in the specification,
Figure FDA00036438560200000211
Figure FDA00036438560200000210
get x 1 ＝h，x 2 ＝γ，x 3 ＝θ，x 4 Q, where θ α + γ denotes the pitch angle, and the altitude subsystems (2) - (5) are written as
Figure FDA0003643856020000031
In the formula (I), the compound is shown in the specification,
Figure FDA0003643856020000032
Figure FDA0003643856020000033
Figure FDA0003643856020000034
Figure FDA0003643856020000035
and 5: for the velocity subsystem, a velocity tracking error is defined
Figure FDA0003643856020000036
Design control input T is
Figure FDA0003643856020000037
In the formula, V d For the speed reference command, k v The control parameter is more than 0, and the control parameter is more than 0,
Figure FDA0003643856020000038
design parameter adaptive estimation law
Figure FDA0003643856020000039
Is composed of
Figure FDA00036438560200000310
In the formula, gamma v The control parameter is more than 0;
step 6: defining height tracking error
Figure FDA00036438560200000311
Designing virtual controllers
Figure FDA00036438560200000312
Is composed of
Figure FDA00036438560200000313
In the formula, k 1 The control parameter is more than 0;
introducing a first order filter
Figure FDA00036438560200000314
In the formula, alpha 2 The control parameter is more than 0;
defining track angle tracking error
Figure FDA00036438560200000315
Designing virtual controllers
Figure FDA00036438560200000316
Is composed of
Figure FDA00036438560200000317
In the formula, k 2 The control parameter is more than 0, and the control parameter is more than 0,
Figure FDA00036438560200000318
introducing a first order filter
Figure FDA0003643856020000041
In the formula, alpha 3 The control parameter is more than 0;
design parameter adaptive estimation law
Figure FDA0003643856020000042
And
Figure FDA0003643856020000043
is composed of
Figure FDA0003643856020000044
In the formula (I), the compound is shown in the specification,
Figure FDA0003643856020000045
and
Figure FDA0003643856020000046
is a control parameter;
defining pitch tracking error
Figure FDA0003643856020000047
Designing virtual controllers
Figure FDA0003643856020000048
Is composed of
Figure FDA0003643856020000049
In the formula, k 3 The control parameter is more than 0;
introducing a first order filter
Figure FDA00036438560200000410
In the formula, alpha 4 The control parameter is more than 0;
designed pitch angle velocity tracking error
Figure FDA00036438560200000411
Design control input delta e Is composed of
Figure FDA00036438560200000412
In the formula, k 4 The control parameter is more than 0, and the control parameter is more than 0,
Figure FDA00036438560200000413
design parameter adaptive estimation law
Figure FDA00036438560200000414
And
Figure FDA00036438560200000415
is composed of
Figure FDA00036438560200000416
In the formula (I), the compound is shown in the specification,
Figure FDA00036438560200000417
and
Figure FDA00036438560200000418
is a control parameter;
and 7: according to the obtained thrust T and rudder deflection angle delta e Returning to the aircraft dynamics models (1) - (5), tracking control is performed on the speed and the altitude.
2. A computer system, comprising: one or more processors, a computer readable storage medium, for storing one or more programs, wherein the one or more programs, when executed by the one or more processors, cause the one or more processors to implement the method of claim 1.
3. A computer-readable storage medium having stored thereon computer-executable instructions for, when executed, implementing the method of claim 1.
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