CN104460681A

CN104460681A - Method for controlling flying of tilt-rotor unmanned helicopter in transition section

Info

Publication number: CN104460681A
Application number: CN201410495971.8A
Authority: CN
Inventors: 浦黄忠; 甄子洋; 郜晨
Original assignee: Nanjing University of Aeronautics and Astronautics
Current assignee: Nanjing University of Aeronautics and Astronautics
Priority date: 2014-09-24
Filing date: 2014-09-24
Publication date: 2015-03-25
Anticipated expiration: 2034-09-24
Also published as: CN104460681B

Abstract

A flight control method for the transition section of a tilt-rotor unmanned helicopter, characterized in that the method adopts optimal predictive control to control the tilt-rotor The flight of the unmanned helicopter transition section is controlled; the specific method is: firstly, a nonlinear model is established for the tiltrotor unmanned helicopter transition section, and then the linear equation described by the state space is obtained by trimming linearization at different nacelle inclination angles; secondly, based on The linear model is used to design the optimal predictive controller to control the flight speed, altitude and attitude angle during the transition process; finally, the inertial delay desalter is designed for the tilt rotor unmanned helicopter to form a smooth transition between different nacelle inclination angles. The purpose of the present invention is to provide a flight control method for the transition section of the tilt-rotor unmanned helicopter, which integrates information such as predictable expected speed, height and attitude angle, and completes the safe switching between the modes of the tilt-rotor unmanned helicopter.

Description

Flight control method for transition section of tilt-rotor unmanned helicopter

技术领域 technical field

本发明涉及倾转旋翼无人直升机过渡段的飞行控制方法，尤其涉及一种最优预见控制的飞行控制方法，属于飞行控制技术领域。 The invention relates to a flight control method for a transition section of a tilt-rotor unmanned helicopter, in particular to a flight control method for optimal predictive control, which belongs to the technical field of flight control.

背景技术 Background technique

倾转旋翼机是一种混合飞行器，既可以像固定翼飞机那样前飞，同时也可以像直升机那样实现垂直起降和悬停。倾转旋翼机的特殊性能，扩展了其在军用和民用中使用范围。美国“Eagle Eye”是目前正在研制的一款倾转旋翼无人直升机。 A tiltrotor is a hybrid aircraft that can fly forward like a fixed-wing aircraft, while also being able to take off and land vertically and hover like a helicopter. The special performance of the tilting rotor aircraft has expanded its range of use in military and civilian use. The American "Eagle Eye" is a tilting rotor unmanned helicopter currently under development.

倾转旋翼机带来了许多新的问题和技术难点：一方面，飞机控制方式和飞行模式比较复杂，增加了过渡段的模型描述难度，既有空气舵控制又有拉力矢量控制，不仅有直升机模式、固定翼模式，又存在直升机模式和固定翼模式之间的切换过程。另一方面，在模式切换过程中飞行器气动力特性极其复杂，而且发动机短舱倾转时有较强的俯仰耦合以及俯仰/沉浮运动耦合，气动力干扰严重，飞机稳定性差。因此，如何使飞机稳定地完成倾转旋翼机模式间的切换对于安全飞行尤其重要。而倾转旋翼无人直升机对飞行控制系统的要求相比有人倾转旋翼机更为苛刻。 Tilting rotorcraft has brought many new problems and technical difficulties: on the one hand, the control method and flight mode of the aircraft are relatively complicated, which increases the difficulty of model description in the transition section. There are both air rudder control and pull vector control. Not only helicopters mode, fixed-wing mode, and there is a switching process between helicopter mode and fixed-wing mode. On the other hand, during the mode switching process, the aerodynamic characteristics of the aircraft are extremely complex, and there is strong pitch coupling and pitch/heavy motion coupling when the engine nacelle is tilted, the aerodynamic interference is serious, and the stability of the aircraft is poor. Therefore, how to make the aircraft stably complete the switch between the modes of the tiltrotor is particularly important for safe flight. The tilt-rotor unmanned helicopter has more stringent requirements on the flight control system than the manned tilt-rotor.

针对倾转旋翼机过渡段的研究，特别是过渡段控制方案的研究，国内外学者提出了多种的控制方法。除了多数的经典控制方法外，Dr.Anthony J.Calise和Rolf T.Rysdyk提出了自适应非线性控制方法，该方法采用网络和模型结合。然而，这些控制方法往往只利用了系统过去和现在的信息。但对于飞机控制来讲，未来轨迹信息是实现规划和设计的。预见控制不仅考虑系统当前的信息，还根据系统的未来信息来决定当前的控制方案，因此可以改善系统的控制性能。 For the research on the transition section of the tilt rotor aircraft, especially the research on the control scheme of the transition section, scholars at home and abroad have proposed a variety of control methods. In addition to most of the classical control methods, Dr.Anthony J.Calise and Rolf T.Rysdyk proposed an adaptive nonlinear control method, which uses a combination of networks and models. However, these control methods often only utilize the past and present information of the system. But for aircraft control, future trajectory information is for planning and design. Foresight control not only considers the current information of the system, but also determines the current control scheme according to the future information of the system, so it can improve the control performance of the system.

目前，预见控制已成功应用在一些领域。如机械装置主动制振时，可应用预见控制在振动产生前或干扰信号出现前就加入控制作用，使控制能量减少很多；又如在机床、机器人等的路径跟踪控制中，由于路径大都事先知道,应用预见控制可大大减少轨迹跟踪误差。而在飞行器控制技术领域的应用很少。 At present, predictive control has been successfully applied in some fields. For example, when the mechanical device actively suppresses vibration, predictive control can be applied to add control before the vibration or interference signal appears, so that the control energy is greatly reduced; and in the path tracking control of machine tools, robots, etc., since the path is mostly known in advance , the application of predictive control can greatly reduce the trajectory tracking error. However, there are few applications in the field of aircraft control technology.

发明内容 Contents of the invention

所要解决的技术问题： Technical problems to be solved:

本发明的目的是提供一种倾转旋翼无人直升机过渡段的飞行控制方法，融合可预见的期望轨迹和期望速度等信息，完成倾转旋翼无人直升机模态之间的安全切换。 The purpose of the present invention is to provide a flight control method for the transition section of the tilt-rotor unmanned helicopter, which integrates information such as predictable expected trajectory and expected speed, and completes the safe switching between the modes of the tilt-rotor unmanned helicopter.

技术方案： Technical solutions:

为了实现以上功能，本发明提供了一种倾转旋翼无人直升机过渡段的飞行控制方法，其特征在于，该方法采用最优预见控制通过融合系统运动方程约束、可预见的理想状态以及最优性能指标约束信息，对倾转旋翼无人直升机过渡段的飞行进行控制；具体方法为： In order to achieve the above functions, the present invention provides a flight control method for the transition section of a tilt-rotor unmanned helicopter. The performance index constraint information is used to control the flight of the tilt-rotor unmanned helicopter in the transition section; the specific method is:

首先对倾转旋翼无人直升机过渡段建立非线性模型，然后在不同短舱倾角下配平线性化得到状态空间描述的线性方程； Firstly, a nonlinear model is established for the transition section of the tilt-rotor unmanned helicopter, and then the linear equation described by the state space is obtained by trimming and linearizing under different nacelle inclination angles;

其次，基于线性模型，设计最优预见控制器，控制过渡过程中的飞行速度、高度和姿态角； Secondly, based on the linear model, an optimal predictive controller is designed to control the flight speed, altitude and attitude angle during the transition process;

最后为倾转旋翼无人直升机在不同短舱倾角之间形成平滑过渡设计惯性延迟式淡化器。 Finally, an inertial delay desalter is designed for the tilt-rotor unmanned helicopter to form a smooth transition between different nacelle inclination angles.

具体的讲，所述不同短舱倾角状态为β_M的0°:15°，15°:30°，30°:45°，45°:60°，60°:75°，75°:90°；在β_M为15°、30°、45°、60°、75°、90°，建立对应于上述六个阶段的倾转旋翼无人直升机过渡段的线性模型，分别表示为{{A1,B1,C1,D1},{A2,B2,C2,D2},…,{A6,B6,C6,D6}}；其中，上述Ai、Bi、Ci、Di分别表示线性模型的系统矩阵，所述每一阶段的{Ai,Bi,Ci,Di}包括纵向模型{Ai_lon,Bi_lon,Ci_lon,Di_lon}和横侧向模型{Ai_lat,Bi_lat,Ci_lat,Di_lat}。 Specifically, the different nacelle inclination states are 0°: 15°, 15°: 30°, 30°: 45°, 45°: 60°, 60°: 75°, 75°: 90° of β _M ; When β _M is 15°, 30°, 45°, 60°, 75°, and 90°, establish a linear model corresponding to the transition section of the tilt-rotor unmanned helicopter in the above six stages, expressed as {{A1, B1, C1, D1}, {A2, B2, C2, D2}, ..., {A6, B6, C6, D6}}; wherein, the above-mentioned Ai, Bi, Ci, and Di respectively represent the system matrix of the linear model, and the {Ai,Bi,Ci,Di} of each stage includes longitudinal model {Ai _lon ,Bi _lon ,Ci _lon ,Di _lon } and lateral model {Ai _lat ,Bi _lat ,Ci _lat ,Di _lat }.

作为一种优选，当所述短舱倾角不超过45°时，倾转旋翼无人直升机过渡段的纵向方式为总距δ_c和纵向周期变距δ_lon联合控制，横侧向为横向周期变距δ_lat和纵向差动周期变距δ_dlon联合控制；则前三个阶段的纵向控制律形式为 As a preference, when the inclination angle of the nacelle does not exceed 45°, the longitudinal mode of the transition section of the tilt-rotor unmanned helicopter is combined control of collective distance δ _c and longitudinal periodic pitch δ _lon , and the lateral direction is horizontal periodic variable δ _lat and longitudinal differential periodic variable pitch δ _dlon joint control; then the longitudinal control law form of the first three stages is

$ΔU Δ U ((k k)) = = [\begin{matrix} {δ δ}_{c c} \\ {δ δ}_{lon the lon} \end{matrix}] = = {F f}_{00} {X x}_{00} ((k k)) + + {Σ Σ}_{j j = = 00}^{{M m}_{R R}} {F f}_{R R} ((j j)) ΔR ΔR ((k k + + j j))$

X₀(k)＝[*u *w *q u w q θ H]^T，R(k)＝[u w q]^T X ₀ (k)＝[*u *w *q u w q θ H] ^T ，R(k)＝[u w q] ^T

式中，u为沿机体坐标x轴方向的飞行速度，w为沿机体坐标z轴方向的飞行速度，q为俯仰角速度，变量前加*表示该变量测量值与理想值的差；δ_c为总距，δ_lon为纵向周期变距，θ为俯仰角；H为飞行高度； In the formula, u is the flight speed along the x-axis direction of the body coordinate, w is the flight speed along the z-axis direction of the body coordinate, q is the pitch angular velocity, adding * before the variable indicates the difference between the measured value and the ideal value of the variable; δ _c is Collective distance, δ _lon is the longitudinal periodic pitch, θ is the pitch angle; H is the flight height;

前三个阶段的横侧向控制律形式为 The form of the lateral control law in the first three stages is

$ΔU Δ U ((k k)) = = [\begin{matrix} {δ δ}_{lat lat} \\ {δ δ}_{dlon dlon} \end{matrix}] = = {F f}_{00} {X x}_{00} ((k k)) + + {Σ Σ}_{j j = = 00}^{{M m}_{R R}} {F f}_{R R} ((j j)) ΔR ΔR ((k k + + j j))$

X₀(k)＝[*β *p *r β p r φ]^T，R(k)＝[β p r]^T。 X ₀ (k)=[*β *p *r β p r φ] ^T , R(k)=[β p r] ^T .

式中，β为侧滑角，p为滚转角速度，r为偏航角速度，φ为滚转角，变量前加*表示该变量测量值与理想值的差；δ_lat为横向周期变距，δ_dlon为纵向差动周期变距 In the formula, β is the sideslip angle, p is the roll angular velocity, r is the yaw angular velocity, φ is the roll angle, adding * in front of the variable indicates the difference between the measured value and the ideal value of the variable; δ _lat is the lateral periodic variable pitch, δ _dlon is longitudinal differential periodic pitch change

作为一种优选，当所述短舱倾角超过45°后，倾转旋翼无人直升机过渡段的纵向方式为总距δ_c和升降舵δ_e联合控制，横侧向为副翼δ_a和方向舵δ_r联合控制；则后三个阶段的纵向控制律形式为 As a preference, when the inclination angle of the nacelle exceeds 45°, the longitudinal mode of the transition section of the tilt-rotor unmanned helicopter is joint control of collective pitch δ _c and elevator δ _e , and the lateral direction is joint control of aileron δ _a and rudder δ _r joint control; then the form of the longitudinal control law in the last three stages is

$ΔU Δ U ((k k)) = = [\begin{matrix} {δ δ}_{c c} \\ {δ δ}_{e e} \end{matrix}] = = {F f}_{00} {X x}_{00} ((k k)) + + {Σ Σ}_{j j = = 00}^{{M m}_{R R}} {F f}_{R R} ((j j)) ΔR ΔR ((k k + + j j))$

式中，δ_e为升降舵偏角； where, δ _e is the elevator deflection angle;

后三个阶段的横侧向控制律形式为 The form of the lateral control law in the last three stages is

$ΔU Δ U ((k k)) = = [\begin{matrix} {δ δ}_{a a} \\ {δ δ}_{r r} \end{matrix}] = = {F f}_{00} {X x}_{00} ((k k)) + + {Σ Σ}_{j j = = 00}^{{M m}_{R R}} {F f}_{R R} ((j j)) ΔR ΔR ((k k + + j j))$

式中，δ_a为副翼偏角，δ_r为方向舵偏角。 In the formula, δ _a is the aileron deflection angle, and δ _r is the rudder deflection angle.

具体的讲，惯性延迟式淡化器的设计方法具体为： Specifically, the design method of the inertial delay desalter is as follows:

设原控制律的输出为U_A(t)，重构控制律的输出为U_B(t)，切换控制律的输出为U(t)；假设在t₀时刻进行控制律切换，到t₁时刻，经过指数衰减作用，而完成控制律切换。该型淡化器的具体算法为 Suppose the output of the original control law is U _A (t), the output of the reconfigured control law is U _B (t), and the output of the switching control law is U(t); suppose the control law is switched at time t ₀ , and at t ₁ time, after exponential decay, and Complete control law switching. The specific algorithm of this type of desalter is

$U u ((t t)) = = \{\begin{matrix} {U u}_{A A} ((t t)) & t t \leq \leq {t t}_{00} \\ {U u}_{A A} ((t t)) {e e}^{- - a a ((t t - - {t t}_{00}))} + + {U u}_{B B} ((t t)) ((11 - - {e e}^{- - a a (({t t - - t t}_{00}))})) & {t t}_{00} < < t t \leq \leq {t t}_{11} \\ {U u}_{B B} ((t t)) & t t > > {t t}_{11} \end{matrix}$

参数a是影响淡化性能的主要因素，其值越大，淡化时间越短，反之，其值越小，淡化时间越长。 The parameter a is the main factor affecting the fade performance, the larger the value, the shorter the fade time, on the contrary, the smaller the value, the longer the fade time.

有益效果： Beneficial effect:

本发明的目的是提供一种倾转旋翼无人直升机过渡段的飞行控制方法，融合可预见的期望速度、高度和姿态角信息，完成倾转旋翼无人直升机模态之间的安全切换。 The purpose of the present invention is to provide a flight control method for the transition section of the tilt-rotor unmanned helicopter, which fuses predictable expected speed, height and attitude angle information to complete the safe switching between the modes of the tilt-rotor unmanned helicopter.

本发明所述的一种倾转旋翼无人直升机过渡段的飞行控制方法，采用最优预见控制通过融合系统运动方程硬约束、可预见的理想状态软约束以及最优性能指标软约束等信息，对倾转旋翼无人直升机过渡段的飞行进行控制。 The flight control method of the transition section of a tilt-rotor unmanned helicopter according to the present invention adopts the optimal predictive control by fusing information such as the hard constraints of the system motion equation, the predictable soft constraints of the ideal state, and the soft constraints of the optimal performance index. Controlling the flight of a tilt-rotor unmanned helicopter in transition.

附图说明 Description of drawings

下面结合附图和实施例对本发明进一步说明： Below in conjunction with accompanying drawing and embodiment the present invention is further described:

图1为倾转旋翼无人直升机过渡段速度安全范围； Figure 1 is the speed safety range of the transition section of the tilt-rotor unmanned helicopter;

图2为本发明所述的倾转旋翼无人直升机过渡段前三个阶段纵向最优预见控制器结构； Fig. 2 is the structure of the longitudinal optimal forecast controller in the first three stages of the transition section of the tilt rotor unmanned helicopter according to the present invention;

图3为本发明所述控制律切换淡化器仿真曲线图； Fig. 3 is a control law switching desalter simulation curve diagram of the present invention;

图4为采用本发明所述的飞行控制方法得到的短舱倾角15度的纵向速度响应曲线； Fig. 4 is the longitudinal velocity response curve of the nacelle inclination angle of 15 degrees obtained by adopting the flight control method of the present invention;

图5为采用本发明所述的飞行控制方法得到的短舱倾角45度的纵向速度响应曲线； Fig. 5 is the longitudinal speed response curve of the nacelle inclination angle of 45 degrees obtained by adopting the flight control method of the present invention;

图6为采用本发明所述的飞行控制方法得到的短舱倾角60度的纵向速度响应曲线； Fig. 6 is the longitudinal velocity response curve obtained by adopting the flight control method of the present invention at a nacelle inclination angle of 60 degrees;

图7为采用本发明所述的飞行控制方法得到的一维轨迹仿真曲线； Fig. 7 is the one-dimensional trajectory simulation curve obtained by adopting the flight control method of the present invention;

图8为采用本发明所述的飞行控制方法得到的二维轨迹仿真曲线。 Fig. 8 is a two-dimensional trajectory simulation curve obtained by using the flight control method of the present invention.

具体实施方式 Detailed ways

本发明提供一种倾转旋翼无人直升机过渡段的最优预见飞行控制方法，为使本发明的目的，技术方案及效果更加清楚，明确，以及参照附图并举实例对本发明进一步详细说明。应当理解，此处所描述的具体实施仅用以解释本发明，并不用于限定本发明。 The present invention provides an optimal predictive flight control method for the transition section of a tilt-rotor unmanned helicopter. In order to make the purpose of the present invention, technical solutions and effects clearer and clearer, the present invention is further described in detail with reference to the accompanying drawings . It should be understood that the specific implementations described here are only used to explain the present invention, not to limit the present invention.

采用本发明的倾转旋翼无人直升机过渡段的最优预见飞行控制方法，具体按照以下步骤。 The optimal predictive flight control method for the transition section of the tilt-rotor unmanned helicopter of the present invention is specifically in accordance with the following steps.

步骤1、首先建立倾转旋翼无人直升机过渡段的非线性模型。在机体坐标系下，飞行器动力学及运动学六自由度方程组可采用如下形式。 Step 1. First establish a nonlinear model of the transition section of the tilt-rotor unmanned helicopter. In the body coordinate system, the six-degree-of-freedom equations of aircraft dynamics and kinematics can take the following form.

$X x = = m m ((\overset{\cdot \cdot}{u u} + + qw qw - - rv rv)) + + mg mg sin sin θ θ$

$Y Y = = m m ((\overset{\cdot &Center Dot;}{v v} + + ru the - - pw pw)) - - mg mg sin sin φ φ cos cos θ θ$

$Z Z = = m m ((\overset{\cdot &Center Dot;}{w w} + + pv PV - - qu qu)) - - mg mg cos cos φ φ cos cos θ θ$

$L L = = {I I}_{xx xxx} \overset{\cdot \cdot}{p p} - - (({I I}_{yy yy} - - {I I}_{zz zz})) qr qr + + {I I}_{yz yz} (({r r}^{22} - - {q q}^{22})) - - {I I}_{xz xz} ((pq pq + + \overset{\cdot &Center Dot;}{r r})) + + {I I}_{xy xy} ((pr pr - - \overset{\cdot \cdot}{q q}))$

$M m = = {I I}_{yy yy} \overset{\cdot &Center Dot;}{q q} - - (({I I}_{zz zz} - - {I I}_{xx xxx})) pr pr + + {I I}_{xz xz} (({p p}^{22} - - {r r}^{22})) - - {I I}_{xy xy} ((qr qr + + \overset{\cdot &Center Dot;}{p p})) {+ + I I}_{yz yz} ((pq pq - - {r r}^{22}))$

$N N = = {I I}_{zz zz} \overset{\cdot &Center Dot;}{r r} - - (({I I}_{xx xxx} - - {I I}_{yy yy})) pq pq + + {I I}_{xy xy} (({q q}^{22} - - {p p}^{22})) - - {I I}_{yz yz} ((pr pr + + \overset{\cdot \cdot}{q q})) + + {I I}_{xz xz} ((qr qr - - \overset{\cdot \cdot}{p p}))$

$p p = = \overset{\cdot &Center Dot;}{φ φ} - - \overset{\cdot &Center Dot;}{ψ ψ} cos cos θ θ$

$q q = = \overset{\cdot &Center Dot;}{θ θ} cos cos φ φ + + \overset{\cdot &Center Dot;}{ψ ψ} sin sin φ φ cos cos θ θ$

$r r = = \overset{\cdot &Center Dot;}{ψ ψ} cos cos φ φ cos cos θ θ - - \overset{\cdot &Center Dot;}{θ θ} sin sin φ φ$

在计算力和力矩时，要考虑机身、旋翼、机翼、平尾和垂尾等每一部分的作用。如机体轴y轴方向的侧力的构成如下 When calculating forces and moments, the role of each part of the fuselage, rotor, wing, horizontal tail, and vertical tail is considered. For example, the composition of the lateral force in the y-axis direction of the body axis is as follows

Y＝Y_fuse+Y_wing+Y_ROT+Y_HT+Y_VT Y＝Y _fuse +Y _wing +Y _ROT +Y _HT +Y _VT

式中，Y_fuse是机身产生的侧力，Y_wing是机翼产生的侧力，Y_ROT是旋翼产生的侧力，Y_HT是平尾产生的侧力，Y_VT是垂尾产生的侧力。 In the formula, Y _fuse is the lateral force generated by the fuselage, Y _wing is the lateral force generated by the wing, Y _ROT is the lateral force generated by the rotor, Y _HT is the lateral force generated by the horizontal tail, Y _VT is the lateral force generated by the vertical tail .

分别对机身、旋翼、机翼、平尾和垂尾进行数学建模后，在MATLAB/Simulink中搭建倾转旋翼无人直升机的非线性数学模型。 After mathematical modeling of the fuselage, rotor, wing, horizontal tail and vertical tail, a nonlinear mathematical model of the tilt-rotor unmanned helicopter is built in MATLAB/Simulink.

然后使用MATLAB/Simulink中的trim函数来实现不同短舱角下对系统的配平，再利用linmod函数，将配平输入量和配平状态量作为linmod函数的参数，在配平点附近将非线性模型线性化。由于倾转旋翼无人直升机过渡段的运动具有很强的时变特性，将过渡段划分六个阶段，分别对应短舱倾角β_M的0°:15°，15°:30°，30°:45°，45°:60°，60°:75°，75°:90°。在β_M为15°，30°，45°，60°，75°，90°建立对应于上述六个阶段的倾转旋翼无人直升机过渡段的线性模型，分别表示为{{A1,B1,C1,D1},{A2,B2,C2,D2},…,{A6,B6,C6,D6}}。其中每一阶段的{Ai,Bi,Ci,Di}包括以x＝[u w q θ H]^T为状态变量、以U＝[δ_c δ_lon δ_e]为控制量的纵向模型{Ai_lon,Bi_lon,Ci_lon,Di_lon}和以x＝[β p r φ]^T为状态变量、以U＝[δ_lat δ_dlon δ_a δ_r]为控制量的横侧向模型{Ai_lat,Bi_lat,Ci_lat,Di_lat}。 Then use the trim function in MATLAB/Simulink to realize the trimming of the system under different nacelle angles, and then use the linmod function to take the trim input and trim state as parameters of the linmod function, and linearize the nonlinear model near the trim point . Since the motion of the transition section of the tilt-rotor unmanned helicopter has strong time-varying characteristics, the transition section is divided into six stages, corresponding to the nacelle inclination angle _βM of 0°: 15°, 15°: 30°, 30°: 45°, 45°:60°, 60°:75°, 75°:90°. When β _M is 15°, 30°, 45°, 60°, 75°, and 90°, a linear model corresponding to the transition section of the tilt-rotor unmanned helicopter in the above six stages is established, which are respectively expressed as {{A1, B1, C1,D1},{A2,B2,C2,D2},...,{A6,B6,C6,D6}}. The { _Ai , Bi, Ci, Di} _of _each stage includes a ^longitudinal model {Ai _lon , Bi _lon , _Ci _lon , _Di _lon _} _and the ^lateral model {Ai _lat , Bi _lat , Ci _lat , Di _lat }.

为了便于公众了解本发明的技术方案，用下表1将本发明涉及的物理量进行说明： In order to facilitate the public's understanding of the technical solution of the present invention, the physical quantities involved in the present invention are described with the following table 1 :

变量 variable 含义 the meaning u u 沿机体坐标x轴方向的飞行速度 Flight speed along the x-axis direction of the body coordinates w w 沿机体坐标z轴方向的飞行速度 Flight speed along the z-axis direction of the body coordinates β_M β _M 短舱倾角 nacelle inclination θ θ 俯仰角 Pitch angle φ φ 滚转角 roll angle ψ ψ 偏航角 Yaw angle p p 滚转角速度 roll rate q q 俯仰角速度 Pitch rate r r 偏航角速度 Yaw rate δ_c δ _c 总距 collective distance δ_lon δ _lon 纵向周期变距 Longitudinal cycle pitch

δ_dlon δ _dlon 纵向差动周期变距 Longitudinal differential periodic pitch change δ_lat δ _lat 横向周期变距 Horizontal periodic variable pitch δ_e δ _e 升降舵偏角 Elevator Declination δ_a δ _a 副翼偏角 Aileron deflection δ_r _δr 方向舵偏角 Rudder deflection m m 飞行器质量 aircraft mass β beta 侧滑角 side slip angle φ φ 滚转角 roll angle p p 滚转角速度 roll rate r r 偏航角速度 Yaw rate δ_a δ _a 副翼偏角 Aileron deflection δ_r _δr 方向舵偏角 Rudder deflection I_xx I _xx 飞机器绕机体坐标系x轴的转动惯量 The moment of inertia of the aircraft around the x-axis of the body coordinate system I_yy I _yy 飞机器绕机体坐标系y轴的转动惯量 The moment of inertia of the aircraft around the y-axis of the body coordinate system I_zz I _zz 飞机器绕机体坐标系z轴的转动惯量 The moment of inertia of the aircraft around the z-axis of the body coordinate system I_xy I _x 飞机器对机体坐标系x、y轴的惯性积 The product of inertia of the aircraft about the x and y axes of the body coordinate system I_xz I _x 飞机器对机体坐标系x、z轴的惯性积 The product of inertia of the aircraft about the x and z axes of the body coordinate system I_yz I _yz 飞机器对机体坐标系y、z轴的惯性积 The product of inertia of the aircraft about the y and z axes of the body coordinate system X x 飞机器受到的沿机体坐标系x轴的合力 The resultant force on the aircraft along the x-axis of the body coordinate system Y Y 飞机器受到的沿机体坐标系y轴的侧力 The lateral force on the aircraft along the y-axis of the body coordinate system Z Z 飞机器受到的沿机体坐标系z轴的合力 The resultant force on the aircraft along the z-axis of the body coordinate system L L 滚转力矩 rolling moment M m 俯仰力矩 pitching moment N N 偏航力矩 Yaw moment 变量前面加Δ Add Δ in front of the variable 该变量的增量 Increment of the variable 变量前面加* Add * in front of the variable 该变量测量值与理想值的差 The difference between the measured value of the variable and the ideal value

表1Table 1

步骤2、基于上述六个线性模型，设计最优预见控制器。 Step 2. Design an optimal predictive controller based on the above six linear models.

首先确定控制方案，如下： First determine the control scheme, as follows:

(1)在短舱倾角不超过45°时，倾转旋翼无人直升机过渡段的纵向方式为总距δ_c和纵向周期变距δ_lon联合控制，横侧向为横向周期变距δ_lat和纵向差动周期变距δ_dlon联合控制。 (1) When the inclination angle of the nacelle does not exceed 45°, the longitudinal mode of the transition section of the tilt-rotor unmanned helicopter is joint control of the collective pitch δ _c and the longitudinal periodic pitch δ _lon , and the lateral direction is the lateral periodic pitch variable δ _lat and Longitudinal differential periodic pitch variable δ _dlon joint control.

(2)短舱倾角超过45°后，倾转旋翼无人直升机过渡段的纵向方式为总距δ_c和升降舵δ_e联合控制，横侧向为副翼δ_a和方向舵δ_r联合控制。 (2) After the inclination angle of the nacelle exceeds 45°, the longitudinal mode of the transition section of the tilt-rotor unmanned helicopter is the joint control of the collective distance δ _c and the elevator δ _e , and the lateral direction is the joint control of the aileron δ _a and the rudder δ _r .

根据上述控制方案，对步骤1中的六个阶段的状态方程进行划分，得到前三个阶段的纵向运动状态方程为 According to the above control scheme, the state equations of the six stages in step 1 are divided, and the longitudinal motion state equations of the first three stages are obtained as

$\{\begin{matrix} {[\begin{matrix} \overset{\cdot &Center Dot;}{u u} & \overset{\cdot \cdot}{w w} & \overset{\cdot &Center Dot;}{q q} & \overset{\cdot &Center Dot;}{θ θ} & \overset{\cdot &Center Dot;}{H h} \end{matrix}]}^{T T} = = {A A}_{lon the lon}^{' '} {[\begin{matrix} u u & w w & q q & θ θ & H h \end{matrix}]}^{T T} + + {B B}_{lon the lon}^{' '} {[\begin{matrix} {δ δ}_{c c} & {δ δ}_{lon the lon} \end{matrix}]}^{T T} \\ {[\begin{matrix} u u & w w & q q & θ θ & H h \end{matrix}]}^{T T} {= = C C}_{lon the lon}^{' '} {[\begin{matrix} u u & w w & q q & θ θ & H h \end{matrix}]}^{T T} \end{matrix}$

前三个阶段的横侧向运动状态方程为 The state equations of lateral motion in the first three stages are

$\{\begin{matrix} {[\begin{matrix} \overset{\cdot \cdot}{β β} & \overset{\cdot \cdot}{p p} & \overset{\cdot \cdot}{r r} & \overset{\cdot \cdot}{φ φ} \end{matrix}]}^{T T} = = {A A}_{lat lat}^{' '} {[\begin{matrix} β β & p p & r r & φ φ \end{matrix}]}^{T T} {+ + B B}_{lat lat}^{' '} {[\begin{matrix} {δ δ}_{lat lat} & {δ δ}_{dlon dlon} \end{matrix}]}^{T T} \\ {[\begin{matrix} β β & p p & r r & φ φ \end{matrix}]}^{T T} = = {C C}_{lat lat}^{' '} {[\begin{matrix} β β & p p & r r & φ φ \end{matrix}]}^{T T} \end{matrix}$

后三个阶段的纵向运动状态方程为 The state equations of longitudinal motion in the last three stages are

$\{\begin{matrix} {[\begin{matrix} \overset{\cdot &Center Dot;}{u u} & \overset{\cdot \cdot}{w w} & \overset{\cdot \cdot}{q q} & \overset{\cdot \cdot}{θ θ} & \overset{\cdot \cdot}{H h} \end{matrix}]}^{T T} = = {A A}_{lon the lon}^{' '' '} {[\begin{matrix} u u & w w & q q & θ θ & H h \end{matrix}]}^{T T} + + {B B}_{lon the lon}^{' '' '} {[\begin{matrix} {δ δ}_{c c} & {δ δ}_{lon the lon} \end{matrix}]}^{T T} \\ {[\begin{matrix} u u & w w & q q & θ θ & H h \end{matrix}]}^{T T} {= = C C}_{lon the lon}^{' '' '} {[\begin{matrix} u u & w w & q q & θ θ & H h \end{matrix}]}^{T T} \end{matrix}$

后三个阶段的横侧向运动状态方程为 The state equations of lateral motion in the last three stages are

$\{\begin{matrix} {[\begin{matrix} \overset{\cdot &Center Dot;}{β β} & \overset{\cdot &Center Dot;}{p p} & \overset{\cdot &Center Dot;}{r r} & \overset{\cdot &Center Dot;}{φ φ} \end{matrix}]}^{T T} = = {A A}_{lat lat}^{' '' '} {[\begin{matrix} β β & p p & r r & φ φ \end{matrix}]}^{T T} {+ + B B}_{lat lat}^{' '' '} {[\begin{matrix} {δ δ}_{lat lat} & {δ δ}_{dlon dlon} \end{matrix}]}^{T T} \\ {[\begin{matrix} β β & p p & r r & φ φ \end{matrix}]}^{T T} = = {C C}_{lat lat}^{' '' '} {[\begin{matrix} β β & p p & r r & φ φ \end{matrix}]}^{T T} \end{matrix}$

以一定的采样时间把上述的线性模型进行离散化。得到下式表示的线性离散时间系 Discretize the above linear model with a certain sampling time. The linear discrete time system represented by the following formula is obtained

x(k+1)＝Ax(k)+BU(k) x(k+1)=Ax(k)+BU(k)

y(k)＝Cx(k) y(k)=Cx(k)

设目标值信号为R(k)，定义误差信号为 Let the target value signal be R(k), define the error signal as

e(k)＝R(k)-y(k) e(k)=R(k)-y(k)

然后导出下面的误差信号系统 Then derive the following error signal system

$[\begin{matrix} e e ((k k + + 11)) \\ Δx Δx ((k k + + 11)) \end{matrix}] = = [\begin{matrix} {I I}_{m m} & - - CA CA \\ 00 & A A \end{matrix}] [\begin{matrix} e e ((k k)) \\ Δx Δx ((k k)) \end{matrix}] + + [\begin{matrix} - - CB CB \\ B B \end{matrix}] ΔU Δ U ((k k)) + + [\begin{matrix} {I I}_{m m} \\ 00 \end{matrix}] ΔR ΔR ((k k + + 11))$

或者 or

X₀(k+1)＝ΦX₀(k)+GΔU(k)+G_RΔR(k+1) X ₀ (k+1)＝ΦX ₀ (k)+GΔU(k)+G _R ΔR(k+1)

对于误差系统，假设已知其从现在时刻开始直到未来M_R步的目标值，定义评价函数为如下包含误差项和输入项的二次型： For the error system, assuming that its target value from the current moment to the future _MR step is known, the evaluation function is defined as the following quadratic form including the error term and the input term:

$\begin{matrix} J J = = {Σ Σ}_{k k = = - - {M m}_{R R} + + 11}^{\infty \infty} [[{e e}^{T T} ((k k)) {Q Q}_{e e} e e ((k k)) {+ + ΔU Δ U}^{T T} ((k k)) HΔU HΔU ((k k))]] \\ = = {Σ Σ}_{k k = = - - {M m}_{R R} + + 11}^{\infty \infty} [[{X x}_{00}^{T T} ((k k)) Q Q {X x}_{00} ((k k)) + + {ΔU Δ U}^{T T} ((k k)) HΔU HΔU ((k k))]] \end{matrix}$

其中：Q是半正定矩阵；Q_e,H为正定矩阵。Q和H是由设计者决定的权重矩阵。 Among them: Q is a positive semi-definite matrix; Q _e , H are positive definite matrices. Q and H are weight matrices decided by the designer.

对于求解误差系统的最优控制问题有几种解法，常用的一种是依据偏微分最优化法，该方法是先假定控制输入的形式，代入评价函数中，依据偏微分要求此函数取最小值，从而求出预见前馈系数。 There are several solutions to the optimal control problem of the error system. The commonly used one is based on the partial differential optimization method. This method first assumes the form of the control input, substitutes it into the evaluation function, and requires this function to take the minimum value according to the partial differential. , so as to obtain the predictive feed-forward coefficient.

设最优预见控制的控制输入形式为 Let the control input form of optimal predictive control be

$ΔU Δ U ((k k)) = = {F f}_{00} {X x}_{00} ((k k)) + + {Σ Σ}_{j j = = 00}^{{M m}_{R R}} {F f}_{R R} ((j j)) ΔR ΔR ((k k + + j j))$

这里，F_R(j)为待定的系数。利用上述方法求解结果为 Here, _FR (j) is a coefficient to be determined. Using the above method to solve the result is

F₀＝-[H+G^TPG]^-1G^TPΦ F ₀ ＝-[H+G ^T PG] ^-1 G ^T PΦ

F_R(j)＝-[H+G^TPG]^-1G^T(ξ^T)^j-1PG_R (j≥1) F _R (j)＝-[H+G ^T PG] ^-1 G ^T (ξ ^T ) ^j-1 PG _R (j≥1)

(ξ＝Φ+GF₀) (ξ=Φ+GF ₀ )

P＝Q+Φ^TPΦ-Φ^TPG[H+G^TPG]^-1G^TPΦ P＝Q+Φ ^T PΦ-Φ ^T PG[H+G ^T PG] ^-1 G ^T PΦ

如果令F₀＝[F_e,F_x]，系统初始值为零，得到 If F ₀ ＝[F _e , F _x ], the initial value of the system is zero, we get

$U u ((k k)) = = {F f}_{e e} \frac{z z}{z z - - 11} e e ((k k)) + + {F f}_{x x} x x ((k k)) + + U u ((k k - - 11)) + + {Σ Σ}_{j j = = 11}^{{M m}_{R R}} {F f}_{R R} ((j j)) R R ((k k + + j j)) + + {Σ Σ}_{j j = = 00}^{{M m}_{R R}} {F f}_{d d} ((j j)) d d ((k k + + j j))$

在每个阶段控制器的目的在于使飞行器在该阶段内前飞速度达到安全过渡的范围，即处于如图1所示的安全范围内。结合控制方案设计的控制器结构为 The purpose of the controller in each stage is to make the forward flight speed of the aircraft reach the range of safe transition in this stage, that is, within the safe range as shown in FIG. 1 . The controller structure designed in combination with the control scheme is

前三个阶段的纵向控制律形式为 The form of the longitudinal control law in the first three stages is

前三个阶段的纵向控制器结构图如图2所示。 The structure diagram of the vertical controller in the first three stages is shown in Fig. 2 .

X₀(k)＝[*β *p *r β p r φ]^T，R(k)＝[β p r]^T X ₀ (k)=[*β *p *r β p r φ] ^T , R(k)=[β p r] ^T

后三个阶段的纵向控制律形式为 The form of the longitudinal control law in the last three stages is

步骤3、六个阶段控制器之间实现平滑过渡，为此设计了惯性延迟式淡化器，设原控制律的输出为U_A(t)，重构控制律的输出为U_B(t)，切换控制律的输出为U(t)。假设在t₀时刻进行控制律切换，到t₁时刻，经过指数衰减作用，而完成控制律切换。该型淡化器的具体算法为 Step 3. To realize smooth transition among the six-stage controllers, an inertial delay desalter is designed for this purpose. The output of the original control law is U _A (t), and the output of the reconstructed control law is U _B (t). The output of the switching control law is U(t). Assuming that the control law is switched at time t ₀ , at time t ₁ , after exponential decay, and Complete control law switching. The specific algorithm of this type of desalter is

假设切换时刻为5s，对于输入的控制量，切换前为-5，切换后为8，这里a＝1.5，切换时间为2s，则淡化器的切换过程如图3所示。此时淡化器的算法如下 Assuming that the switching time is 5s, the input control amount is -5 before the switching and 8 after the switching, here a=1.5, and the switching time is 2s, the switching process of the desalter is shown in Figure 3 . At this point, the algorithm of the fader is as follows

$u u ((t t)) = = \{\begin{matrix} - - 55 & t t \leq \leq 55 \\ {- - 55 e e}^{- - 1.5 1.5 ((t t - - 55))} + + 88 (({11 - - e e}^{- - 1.5 1.5 ((t t - - 55))})) & 55 < < t t \leq \leq 77 \\ 88 & t t > > 77 \end{matrix}$

为了验证本发明在倾转旋翼过渡段飞行控制上的有效性，进行如下仿真实验。仿真工具采用MATLAB软件，对象采用XV-15倾转旋翼飞行器进行分析。仿真参数设置如表2所示。 In order to verify the effectiveness of the present invention on the flight control of the tilt rotor transition section, the following simulation experiments are carried out. The simulation tool adopts MATLAB software, and the object adopts XV-15 tilt-rotor aircraft for analysis. The simulation parameter settings are shown in Table 2 .

时间t(s) time t(s) 0 0 0～5 0～5 5～15 5～15 15～55 15～55 β_M β _M 0° 0° 15° 15° 30° 30° 60° 60°

u(m/s) u(m/s) 20 20 30 30 50 50 60 60 w(m/s) w(m/s) 2.85 2.85 3.45 3.45 4.65 4.65 7.6 7.6

表2Table 2

在0～5s内以β_M＝15°时的配平线性模型进行控制，在5～15s内以β_M＝30°时的配平线性模型进行控制，在15～55s内以β_M＝60°时的配平线性模型进行控制，表2中前飞速度u的设定安全落在图1所示的安全范围内。设置速度的跟踪方式为前15s内，u以加速度a_u＝2m/s²，w以加速度a_w＝0.12m/s²的爬升；在15s时刻u输入幅值为10的阶跃信号，w输入幅值为2.95的阶跃信号。设初始高度为100m，按照上述步骤设计控制最优预见控制器，控制效果如图4到图8所示，其中图7、8所示的整个阶段轨迹根据图4到图6的速度曲线计算得到。在图4到图6的纵坐标表示当前速度减去初始速度的差值，在图4中，虽然u的控制结果出现稳态误差，但是在安全过渡范围内，不会影响倾转旋翼的过渡。在图6中w的控制初始时有个下降的过程，但是由于w初值大于0，所以会出现短时间内的高度下降现象，之后会恢复爬升状态，从图7的高度曲线变化中也可以看出。本仿真中控制的目的主要在于速度控制，不是轨迹跟踪，因此轨迹会跟理想轨迹有偏差。上述仿真结果充分表明，基于最有预见控制的倾转旋翼无人直升机系统具有良好的响应性能。 Control with the trim linear model when β _M =15° within 0~5s, control with the trim linear model when β _M =30° within 5~15s, and control with β _M =60° within 15~55s The trim linear model is controlled, and the setting of the forward flight speed u in Table 2 is safely within the safe range shown in Figure 1 . Set the tracking mode of the speed as within the first 15s, u climbs with acceleration a _u =2m/s ² , w with acceleration a _w =0.12m/s ² ; Enter a step signal with an amplitude of 2.95. Assuming the initial height is 100m, design and control the optimal predictive controller according to the above steps, the control effect is shown in Figure 4 to Figure 8 , and the trajectory of the whole stage shown in Figure 7 and 8 is calculated according to the speed curves in Figure 4 to Figure 6 . The ordinate in Figure 4 to Figure 6 represents the difference between the current speed minus the initial speed. In Figure 4 , although the control result of u has a steady-state error, it will not affect the transition of the tilt rotor within the safe transition range . In Figure 6 , the initial control of w has a descending process, but because the initial value of w is greater than 0, there will be a short-term height drop phenomenon, and then it will resume the climbing state, which can also be seen from the height curve changes in Figure 7 see. The purpose of control in this simulation is mainly speed control, not trajectory tracking, so the trajectory will deviate from the ideal trajectory. The above simulation results fully show that the tilt-rotor unmanned helicopter system based on the most predictable control has good response performance.

可以理解的是，对本领域普通技术人员来说，可以根据本发明的技术方案及其发明构思加以等同替换或改变，而所有这些改变或替换都应属于本发明所附的权利要求的保护范围。 It can be understood that those skilled in the art can make equivalent replacements or changes according to the technical solutions and inventive concepts of the present invention, and all these changes or replacements should belong to the protection scope of the appended claims of the present invention.

Claims

1. A flight control method for the transition section of a tilting rotor unmanned helicopter , characterized in that the method adopts optimal predictive control to control the inclination by integrating system motion equation constraints, foreseeable ideal state and optimal performance index constraint information. Control the flight of the transition section of the rotary-wing unmanned helicopter; the specific method is:

Firstly, a nonlinear model is established for the transition section of the tilt-rotor unmanned helicopter, and then the linear equation described by the state space is obtained by trimming and linearizing under different nacelle inclination angles;

Secondly, based on the linear model, an optimal predictive controller is designed to control the flight speed, altitude and attitude angle during the transition process;

Finally, an inertial delay desalter is designed for the tilt-rotor unmanned helicopter to form a smooth transition between different nacelle inclination angles.

2. the flight control method of a kind of tilt rotor unmanned helicopter transition section according to claim 1, is characterized in that: described different nacelle inclination angle states are 0 °: 15 °, 15 °: 30 ° of β _M , 30°:45°, 45°:60°, 60°:75°, 75°:90°; when β _M is 15°, 30°, 45°, 60°, 75°, 90°, the corresponding The linear models of the transition section of the tiltrotor unmanned helicopter in the above six stages are expressed as {{A1,B1,C1,D1},{A2,B2,C2,D2},…,{A6,B6,C6, D6}}; wherein, the above-mentioned Ai, Bi, Ci, Di respectively represent the system matrix of the linear model, and the {Ai, Bi, Ci, Di} of each stage includes the longitudinal model {Ai _lon , Bi _lon , Ci _lon , Di _lon } and the lateral model {Ai _lat , Bi _lat , Ci _lat , Di _lat }.

3. The flight control method of a kind of tilt rotor unmanned helicopter transition section according to claim 2, characterized in that: when the nacelle inclination angle does not exceed 45 °, the longitudinal direction of the tilt rotor unmanned helicopter transition section The method is joint control of collective pitch δ _c and longitudinal periodic variable pitch δ _lon , and lateral direction is joint control of lateral periodic variable pitch δ _lat and longitudinal differential periodic variable pitch δ _dlon ; then the form of the longitudinal control law in the first three stages is

X ₀ (k)＝[*u *w *q u w q θ H] ^T ，R(k)＝[u w q] ^T

In the formula, u is the flight speed along the x-axis direction of the body coordinate, w is the flight speed along the z-axis direction of the body coordinate, q is the pitch angular velocity, adding * before the variable indicates the difference between the measured value and the ideal value of the variable; δ _c is Collective distance, δ _lon is the longitudinal periodic pitch, θ is the pitch angle; H is the flight height;

The form of the lateral control law in the first three stages is

X ₀ (k)=[*β *p *r β p r φ] ^T , R(k)=[β p r] ^T .

In the formula, β is the sideslip angle, p is the roll angular velocity, r is the yaw angular velocity, φ is the roll angle, adding * in front of the variable indicates the difference between the measured value and the ideal value of the variable; δ _lat is the lateral periodic variable pitch, δ _dlon is the longitudinal differential periodic pitch.

4. The flight control method of a transition section of a tilt-rotor unmanned helicopter according to claim 3, characterized in that: after the nacelle inclination exceeds 45°, the longitudinal mode of the transition section of the tilt-rotor unmanned helicopter is the combined control of the collective pitch δ _c and the elevator δ _e , and the lateral direction is the joint control of the aileron δ _a and the rudder δ _r ; then the form of the longitudinal control law in the last three stages is

X ₀ (k)＝[*u *w *q u w q θ H] ^T ，R(k)＝[u w q] ^T

where, δ _e is the elevator deflection angle;

The form of the lateral control law in the last three stages is

X ₀ (k)=[*β *p *r β p r φ] ^T , R(k)=[β p r] ^T .

In the formula, δ _a is the aileron deflection angle, and δ _r is the rudder deflection angle.

5. according to the flight control method of any kind of tilt-rotor unmanned helicopter transition section described in claim 1 to 4, it is characterized in that: the design method of inertia delay type desalter is specifically:

Suppose the output of the original control law is U _A (t), the output of the reconfigured control law is U _B (t), and the output of the switching control law is U(t); suppose the control law is switched at time t ₀ , and at t ₁ time, after exponential decay, and Complete control law switching. The specific algorithm of this type of desalter is

The parameter a is the main factor affecting the fade performance, the larger the value, the shorter the fade time, on the contrary, the smaller the value, the longer the fade time.