An angle-constrained guidance and control method for aircraft

Abstract

This invention discloses an angle-constrained guidance and control method for an aircraft, comprising the following steps: establishing a guidance system based on the relative motion equations between the aircraft and the target; setting up a timing disturbance observer for sub-estimation and feedforward compensation of the target's flight state; obtaining a timing sliding mode guidance law based on the timing disturbance observer and the guidance system; and controlling the aircraft's flight attitude through the timing guidance law. The angle-constrained guidance and control method disclosed in this invention can estimate and compensate for the target even without prior information, accurately acquiring target maneuvers and improving target strike accuracy.

Description

Technical Field

[0001] This invention relates to an angle-constrained guidance and control method for aircraft, belonging to the field of guidance and control. Background Technology

[0002] For guided aircraft, flight control methods generally aim to intercept targets with minimal miss distance. However, when the target is a high-speed moving target, traditional flight control methods struggle to accurately identify the target's maneuvers and are unable to eliminate the influence of unknown target maneuvers during the control process.

[0003] Existing technologies include methods that use observers to observe target maneuvers and then compensate for guidance laws. However, these methods suffer from long observation times, difficulty in completing compensation within a fixed time, and inability to take into account the aircraft's impact angle.

[0004] Therefore, it is necessary to conduct more in-depth research on guidance methods to solve the above problems. Summary of the Invention

[0005] To overcome the above problems, the inventors conducted in-depth research and proposed an angle-constrained guidance and control method for aircraft, comprising the following steps:

[0006] A guidance system is established based on the relative motion equations between the aircraft and the target;

[0007] A timed disturbance observer is set up to estimate the target's flight state and provide feedforward compensation.

[0008] Based on the guidance system, a timing sliding mode guidance law is obtained using a timing disturbance observer;

[0009] The flight attitude of the aircraft is controlled by a timed sliding mode guidance law.

[0010] In a preferred embodiment, the equation of relative motion is expressed as:

[0011]

[0012]

[0013]

[0014] Where r represents the relative distance between the aircraft and the target, θ L Indicates the aircraft's line-of-sight tilt angle. Indicates the aircraft's line-of-sight deflection angle; This represents the acceleration of the aircraft in the line-of-sight coordinate system. Let be the acceleration of the target in the line-of-sight coordinate system.

[0015] In a preferred embodiment, the guidance system is configured as follows:

[0016]

[0017]

[0018]

[0019] Where, x1=[x 11 ,x 12 ] T x2=[x 21 ,x 22 ] T For system state variables, F, B, a T a M b 1θ , It is an intermediate variable.

[0020] In a preferred embodiment, the timing disturbance observer is configured as follows:

[0021]

[0022]

[0023]

[0024]

[0025]

[0026] Among them, V L =[V L1 V L2 ] T G represents the components of relative velocity in the vertical direction and the line-of-sight direction. 1θ For a Tθ The first derivative, for The first derivative.

[0027] In a preferred embodiment, the estimation error of the timing disturbance observer Set to:

[0028]

[0029] in, For V L The estimated value satisfies:

[0030]

[0031] for The estimated value satisfies:

[0032]

[0033] Among them, c 1d ,c 2d ,α 1d ,α 2d c is a constant. 3d ,c 4d ,α 3d ,α 4d η is a constant; d Let σ be a sequence of numbers with two positive constants. d For sliding mode variables, the symbol is... This represents the Hadamard product operation, where sig(·) is a user-defined function, and the user-defined function sig(·) is represented as sig a (x)=|x| a sign(x).

[0034] In a preferred embodiment, the sliding mode variable is set as follows:

[0035]

[0036] In a preferred embodiment, a timing sliding mode guidance law is obtained through sliding mode control, and the sliding surface s1 of the system is set as follows:

[0037]

[0038]

[0039] Where ρ(x1) is an intermediate parameter, β is a constant, preferably 2, and λ1 and μ1 are positive constants, and satisfy the following conditions:

[0040] The derivative of the sliding surface is expressed as:

[0041]

[0042] in,

[0043] In a preferred embodiment, the obtained timing sliding mode guidance law 'a' is expressed as:

[0044]

[0045] Where k1 > 0, λ2 and μ2 are positive constants, and satisfy the following conditions:

[0046] The beneficial effects of this invention include:

[0047] (1) It can estimate and feedforward compensation of the target without prior information, and can accurately obtain the target maneuver, thus improving the accuracy of the target strike;

[0048] (2) It can achieve precise interception and strike of maneuvering targets under different angle constraints, thereby enhancing the damage capability of aircraft. Attached Figure Description

[0049] Figure 1 A schematic flowchart of an aircraft angle constraint guidance and control method according to a preferred embodiment of the present invention is shown.

[0050] Figure 2 A diagram showing the three-dimensional spatial relative motion relationship between the projectile and the target in an angle-constrained guidance and control method for an aircraft according to a preferred embodiment of the present invention is provided.

[0051] Figure 3 The simulation curves of the flight trajectory under different conditions in Example 1 are shown;

[0052] Figure 4 The simulation results of the aircraft's line-of-sight angle under different conditions in Example 1 are shown. Detailed Implementation

[0053] The present invention will be further described in detail below with reference to the accompanying drawings and embodiments. Through these descriptions, the features and advantages of the present invention will become clearer and more apparent.

[0054] The term “exemplary” as used herein means “serving as an example, embodiment, or illustration.” Any embodiment illustrated herein as “exemplary” is not necessarily to be construed as superior to or better than other embodiments. Although various aspects of embodiments are shown in the accompanying drawings, the drawings are not necessarily drawn to scale unless specifically indicated otherwise.

[0055] According to the present invention, an angle-constrained guidance and control method for an aircraft is provided, such as... Figure 1 As shown, it includes the following steps:

[0056] A guidance system is established based on the relative motion equations between the aircraft and the target;

[0057] A timed disturbance observer is set up to estimate the target's flight state and provide feedforward compensation.

[0058] Based on the guidance system, a timing sliding mode guidance law is obtained using a timing disturbance observer;

[0059] The flight attitude of the aircraft is controlled by a timed sliding mode guidance law.

[0060] According to the present invention, the equation of relative motion is expressed as:

[0061]

[0062]

[0063]

[0064] Where r represents the relative distance between the aircraft and the target, θ L Indicates the aircraft's line-of-sight tilt angle. Indicates the aircraft's line-of-sight deflection angle; This represents the acceleration of the aircraft in the line-of-sight coordinate system. Let be the acceleration of the target in the line-of-sight coordinate system.

[0065] Furthermore, the relationship between the line-of-sight coordinate system and the inertial coordinate system is as follows: Figure 2 As shown, where OX E Y E Z E Represents an inertial coordinate system, OX L Y L Z L This represents the line-of-sight coordinate system.

[0066] In a preferred embodiment, the guidance system is configured as follows:

[0067]

[0068]

[0069]

[0070] Where, x1=[x 11 ,x 12 ] T x2=[x 21 ,x 22 ] T For system state variables, F, B, a T a M b 1θ , θ is an intermediate variable. Ld Indicates the aircraft's desired line-of-sight tilt angle. This indicates the aircraft's desired line-of-sight deflection angle.

[0071] Traditional sliding mode control guidance methods typically only include line-of-sight tilt angle and line-of-sight deflection angle in the system state. The system state variables set in this invention include not only the aircraft's line-of-sight tilt angle and line-of-sight deflection angle, but also the aircraft's desired line-of-sight tilt angle and desired line-of-sight deflection angle, thereby constraining the aircraft's flight angle.

[0072] The disturbance observer is used to observe the target maneuver and eliminate the impact of the target maneuver on the system. It converts the difference between the actual model and the ideal model output caused by external disturbances and changes in model parameters into control inputs, that is, it observes the equivalent disturbance and introduces an equal amount of compensation in the control to achieve complete suppression of the disturbance.

[0073] Traditional perturbation observers typically require the acquisition of the target's maneuver or the upper bound of the target's maneuver derivative in order to reduce the impact of the target's accuracy on the system. In this invention, a perturbation observer is proposed that does not require the target's maneuver or the upper bound of its derivative, and achieves its performance without overestimation. This perturbation observer can not only eliminate the influence of unknown maneuvers, but also estimate and feedforward the maneuvering target within a fixed time.

[0074] Specifically, the timed disturbance observer is configured as follows:

[0075]

[0076]

[0077]

[0078]

[0079]

[0080] Among them, V L =[V L1 V L2 ] T G represents the components of relative velocity in the vertical direction and the line-of-sight direction. 1θ For a Tθ The first derivative, for The first derivative.

[0081] Preferably,

[0082] The timing disturbance observer set in this invention can estimate and feedforward the target's flight state within a fixed time without prior information, thereby accurately acquiring the target's maneuver and improving the target strike accuracy.

[0083] Furthermore, the estimation error of the timing disturbance observer Set to:

[0084]

[0085] in, For V L The estimated value satisfies:

[0086]

[0087] for The estimated value satisfies:

[0088]

[0089] Among them, c 1d ,c 2d ,α 1d ,α 2d c is a constant. 3d ,c 4d ,α 3d ,α 4d η is a constant; d Let σ be a sequence of numbers with two positive constants. d For sliding mode variables, the symbol is... This represents the Hadamard product operation, where sig(·) is a user-defined function, and the user-defined function sig(·) is represented as sig a (x)=|x| a sign(x).

[0090] In a preferred embodiment, the sliding mode variable is set as follows:

[0091]

[0092] The sliding mode variable in this invention is set in an exponential form using a custom function, achieving fixed-time convergence.

[0093] Preferably, c 1d ,c 2d ,α 1d <1, α 1d <1, α 2d >1;

[0094] More preferably, c 1d =c 2d =0.5,α 1d =0.5,α 2d =1.5;

[0095] Preferably, c 3d ,c 4d <1,α 3d <1,α 4d >1,;

[0096] More preferably, c 3d =0.5,c 4d =0,α 3d =0.5,α 4d =1.5;

[0097] Preferably,

[0098] In a preferred embodiment, a timing sliding mode guidance law is obtained through sliding mode control, and the sliding surface s1 of the system is set as follows:

[0099]

[0100]

[0101] Where ρ(x1) is an intermediate parameter, β is a constant, preferably 2, and λ1 and μ1 are positive constants, and satisfy the following conditions:

[0102] The derivative of the sliding surface is expressed as:

[0103]

[0104] in,

[0105] Unlike traditional sliding surfaces, the exponential coefficients in the sliding surface of this invention include state variables x1, which can ensure that x1 and x2 achieve stability over a fixed time.

[0106] In this invention, the sliding surface enables the system variables x1 and x2 to achieve fixed-time stability.

[0107] According to the present invention, the obtained timing sliding mode guidance law 'a' is expressed as:

[0108]

[0109] Where k1 > 0, λ2 and μ2 are positive constants, and satisfy the following conditions:

[0110] Example

[0111] Example 1

[0112] To conduct simulation experiments and obtain the guidance law, the following steps are included:

[0113] A guidance system is established based on the relative motion equations between the aircraft and the target;

[0114] A timed disturbance observer is set up to estimate the target's flight state and provide feedforward compensation.

[0115] Based on the guidance system, a timing sliding mode guidance law is obtained using a timing disturbance observer;

[0116] The equation of relative motion is expressed as:

[0117]

[0118]

[0119]

[0120] The guidance system is configured as follows:

[0121]

[0122]

[0123]

[0124] The timing disturbance observer is configured as follows:

[0125]

[0126]

[0127]

[0128]

[0129]

[0130] in,

[0131] The estimation error of the timing disturbance observer Set to:

[0132]

[0133] in, For V L The estimated value satisfies:

[0134]

[0135] for The estimated value satisfies:

[0136]

[0137] Among them, c 1d =c 2d =0.5,α 1d =0.5,α 2d =1.5.

[0138] The sliding mode variable is set as follows:

[0139]

[0140] in,

[0141] A timing sliding mode guidance law is obtained through sliding mode control, and the sliding surface s1 of the system is set as follows:

[0142]

[0143]

[0144] The obtained timing sliding mode guidance law is expressed as:

[0145]

[0146] Where k1 takes the value of 2, and λ1, λ2, μ1, μ2 take the values ​​of 2.5, 2, 1, and 0.5, respectively.

[0147] In the simulation, the initial position of the target is set to (X). T ,Y T Z T )=(7424,5735.8,3461), the relative distance between the missile and the target is r0=10000m, and the speed of the aircraft is V. m =1000m / s, initial trajectory inclination angle is θ M =65°, initial ballistic deviation angle ψ M =0°; target speed is V T =800m / s, initial trajectory inclination angle is θ T =20°, initial ballistic deviation angle ψ T =120°.

[0148] The aircraft is configured to engage targets based on different terminal missile line-of-sight angles, with the desired target line-of-sight angles being:

[0149] Simulation curves of the flight trajectories of the aircraft and target under different conditions are as follows: Figure 3 As shown, Figure 4 The simulation results of the aircraft's line-of-sight angle under different conditions are shown. From Figure 3 As can be seen, the aircraft can accurately hit the target under different angular constraints; from Figure 4 As can be seen, under different conditions, the aircraft's line-of-sight angle can gradually converge to the desired value.

[0150] In the description of this invention, it should be noted that the terms "upper," "lower," "inner," "outer," "front," and "rear," etc., indicate the orientation or positional relationship based on the orientation or positional relationship in the working state of this invention, and are only for the convenience of describing this invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation, and therefore should not be construed as a limitation of this invention. Furthermore, the terms "first," "second," "third," and "fourth" are used for descriptive purposes only and should not be construed as indicating or implying relative importance.

[0151] In the description of this invention, it should be noted that, unless otherwise explicitly specified and limited, the terms "installation," "connection," and "linking" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; and they can refer to the internal communication between two components. Those skilled in the art can understand the specific meaning of the above terms in this invention based on the specific circumstances.

[0152] The present invention has been described above with reference to preferred embodiments; however, these embodiments are merely exemplary and illustrative. Various substitutions and modifications can be made to the present invention based on these embodiments, all of which fall within the scope of protection of the present invention.

Claims

1. An aircraft angle-constrained guidance control method, characterized in that, The method comprises the following steps: a guidance system is established according to a relative motion equation of the aircraft and the target; a timed disturbance observer is set up for estimating and feedforward compensating the flight state of the target; a timed sliding mode guidance law is obtained based on the timed disturbance observer on the basis of the guidance system; the flight attitude of the aircraft is controlled through the timed sliding mode guidance law; the relative motion equation is expressed as: where r represents the relative distance between the aircraft and the target, θ L represents the line-of-sight angle of the aircraft, represents the line-of-sight angle of the aircraft, is the acceleration of the aircraft in the line-of-sight coordinate system; is the acceleration of the target in the line-of-sight coordinate system; the guidance system is set as: where x1 = [x 11 ,x 12 ] T x2 = [x 21 ,x 22 ] T are system state variables, F, B, a T , a M , b 1θ , are intermediate variables, θ Ld represents the desired line-of-sight angle of inclination of the aircraft, represents the desired line-of-sight angle of declination of the aircraft; the timed disturbance observer is set as: where V L = [V L1 ,V L2 ] T denotes the components of the relative velocity in the vertical and line-of-sight directions, g 1θ is the first derivative of a Tθ , is the first derivative of ; Estimation error of the timing perturbation observer is set to: wherein V is L an estimate of the value of V, satisfying: For an estimate of the value of, satisfying: where c 1d ,c 2d ,α 1d ,α 2d are constants, c 3d ,c 4d ,α 3d ,α 4d are constants; η d is a sequence of two normal numbers, σ d is a sliding mode variable, and the sign denotes the Hadamard product operation, and sig(·) is a self-defined function, which is denoted as sig a (x) = |x| a sign(x). a sliding mode surface s1 of the system is set as: wherein p(x1) is an intermediate variable, β is a constant greater than 1, λ1, μ1 are normal numbers, and satisfy a derivative of the sliding mode surface is expressed as: wherein the obtained timed sliding mode guidance law a is expressed as: where k1>0, λ2, μ2are normal numbers, and satisfy 2. The aircraft angle constraint guidance control method according to claim 1, characterized in that, The sliding mode variable is set as:

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