An integrated method of guidance and control for exoatmospheric interceptor based on parameter identification

Abstract

The invention relates to the field of spacecraft guidance and control, discloses an integrated guidance and control method for an exoatmospheric interceptor based on parameter identification, constructs an integrated guidance and control model for an axially uncontrollable interceptor, proposes a non-singular fixed-time sliding surface based on a tanh function in order to improve the robustness of the control system in the presence of measurement noise and external interference, designs an adaptive attitude-orbit joint control system in combination with parameter identification theory under the condition that the seeker field of view constraint and the interceptor inertial parameters are unknown, applies fixed-time control theory to design a guidance law and an attitude tracking controller for a terminal guidance stage of a space interceptor, comprehensively considers the coupling effect between attitude and orbit in the interception process, and the uncertainty of the interceptor's own model, and finally realizes the precise strike of the interceptor spacecraft on a target in the terminal guidance stage under the unfavorable factors such as measurement noise, model uncertainty and attitude-orbit coupling.

Description

Technical Field

[0001] The present invention relates to the field of spacecraft guidance and control, and in particular to an integrated method for guidance and control of an exoatmospheric interceptor based on parameter identification. Background Art

[0002] As the status of aerospace systems in the military continues to rise, countries around the world have developed their own anti-satellite technologies. Among them, space interception is a typical anti-satellite method, and the accuracy of terminal interception and the robustness of the control system are the key to the success of the mission. During the mission execution, due to the limitations of sensors, actuators and satellite structures, the system needs to meet different constraints and convergence time requirements.

[0003] Accurate terminal guidance is the key to the successful attack of kinetic interceptors. Considering the size of the target satellite and the measurement error of the sensor, higher requirements are placed on the accuracy and robustness of the space interception control system. As the earliest guidance law, the proportional guidance law has been proven to be able to successfully intercept targets in both ground-to-air and air-to-air scenarios. Unlike the interception situation within the atmosphere, the interception of targets outside the atmosphere often uses direct force control, and the guidance law adopts a true proportional guidance law with acceleration perpendicular to the line of sight of the missile and target. In engineering applications, proportional guidance has certain advantages because it is easy to design and requires less information. However, proportional guidance has the problem of large misses when intercepting highly maneuverable targets, and it is also difficult to cope with system interference. The sliding mode variable structure control method has attracted much attention due to its strong robustness and insensitivity to model parameters. The guidance law design based on sliding mode control theory has the following advantages: (1) strong robustness to external interference; (2) only requires line of sight angular rate and relative velocity information; (3) simple structure and can be combined with methods such as adaptive control. Compared with proportional guidance, the sliding mode guidance law can better meet the requirements of convergence time and accuracy for exoatmospheric interception, and is robust to measurement errors and inertial parameter changes. Although the existing sliding mode guidance law can complete the interception of maneuvering targets and consider the attack angle constraint, its convergence time is still affected by the initial state value, and the inertial parameters of the interceptor cannot be unknown. Therefore, it is particularly important to design a terminal interception control system that considers field of view constraints, attitude and orbit coupling, and does not require inertial parameters, and has fast convergence characteristics. Based on this, the present invention proposes an integrated guidance and control method for exoatmospheric interceptors based on parameter identification. Summary of the invention

[0004] The object of the present invention is to provide an integrated method for guidance and control of an exoatmospheric interceptor based on parameter identification to solve the problems raised in the above-mentioned background technology.

[0005] To achieve the above object, the present invention provides the following technical solution: an integrated method for guidance and control of an exoatmospheric interceptor based on parameter identification, comprising the following steps:

[0006] Step 1: Establish the relative orbital dynamics model of the intercepting spacecraft and the attacking spacecraft;

[0007] Step 2: Establish the relative attitude dynamics model of the spacecraft;

[0008] Step 3: Based on the possible singularity problem of the terminal sliding mode variable structure control method, a non-singular fixed-time sliding mode surface based on the tanh function is designed;

[0009] Step 4: Based on the inertial parameters of the interceptor spacecraft itself and combined with parameter identification theory, an adaptive attitude and orbit joint system is designed;

[0010] Step 5: Based on steps 3 and 4, design a fixed-time attitude tracking controller and terminal guidance law considering field of view constraints.

[0011] Preferably: the relative motion relationship between the interceptor and the target spacecraft in step 1 can be expressed as

[0012]

[0013] Where r = [r x r y r z ] T is the position vector of the interceptor and the target spacecraft in the geocentric inertial coordinate system; are the positions of the interceptor and target spacecraft in the geocentric inertial coordinate system; μ = 3.986 × 10 14 is the gravitational constant; is the mass of the interceptor spacecraft; It is the control force of the interceptor in the inertial frame; The unknown disturbance force acting on the interceptor.

[0014] Define the interceptor's line of sight angle and angular rate relative to the target in the inertial frame:

[0015]

[0016]

[0017] Among them, q1 and q2 are the line of sight inclination and declination expressed in the inertial system respectively; is the relative distance between the interceptor spacecraft and the target spacecraft.

[0018] Define state quantity q = [q1 q2] T , combined with the above formula, we have

[0019]

[0020] in, The complete expression is as follows

[0021]

[0022]

[0023]

[0024]

[0025] Let the control force on the interceptor spacecraft body coordinate system be f c =[f c1 ,f c2 ] T , converted to the inertial system, we have

[0026] f I =R1f c

[0027] in, R1 is the transformation matrix from the spacecraft body coordinates to the inertial coordinate system. The relative orbital dynamics of the interceptor in the inertial system can be further written as

[0028]

[0029] Preferably, in step 2, the modified Rodrigues parameter method is used to describe the attitude operation equation of the interception spacecraft:

[0030]

[0031] in, denote the Rodrigues parameter and angular velocity of the interceptor spacecraft in the body coordinate system relative to the inertial coordinate system, is the control torque for the intercepting spacecraft, is the disturbance torque, is the moment of inertia of the intercepting spacecraft, the complete expression is as follows

[0032]

[0033] For the intercepting spacecraft, its moment of inertia has the following properties

[0034] J 12 =J 21 ,J 13 =J 31 ,J 23 =J 32

[0035] According to the line of sight inclination and line of sight deflection during the interception process, the conversion matrix from the inertial system to the line of sight represented by the line of sight angle is compared with the conversion matrix from the inertial system to the line of sight represented by the Euler angle, and the expected attitude angle of the interceptor represented by the Euler angle can be obtained, and then the expected Rodrigues parameter can be obtained, which is expressed as σ d The specific expression is as follows

[0036]

[0037] For the desired angular velocity, we have

[0038]

[0039] According to the definition of Rodrigues parameter and angular velocity, the relative Rodrigues parameter and relative angular velocity can be calculated.

[0040]

[0041] ω e =ω c -R cs ω d

[0042] Among them, R cs is the coordinate transformation matrix from the sight system to this system.

[0043] Further, the kinematic and dynamic equations of the relative attitude of the intercepting spacecraft in the inertial system are obtained as follows:

[0044]

[0045] in,

[0046] Preferably, the non-singular fixed-time sliding surface based on the tanh function designed in step 3 is as follows:

[0047]

[0048] Among them, x is the state quantity, is a positive constant, a s ∈(1,2), b s ∈(0,1), a s b s >1.

[0049] The derivative of S with respect to time is

[0050]

[0051] When x approaches zero, due to The above formula can be simplified as

[0052]

[0053] Considering a s >1 and a s b s >1, There will be no singular phenomenon, so the designed sliding surface is non-singular. When the sliding surface S converges, the system state x also converges. The Lyapunov equation is selected as

[0054]

[0055] The following relationship holds

[0056]

[0057] The convergence of the system state x is proved.

[0058] Preferably: In step 4, the adaptive attitude and trajectory joint system is designed in combination with the parameter identification theory, specifically, the variable θ is defined as [J 11 ,J 12 ,J 13 ,J 22 ,J 23 ,J 33 ] T and linear operators At this time, the moment of inertia can be rewritten as follows

[0059]

[0060] Then the relative motion model of the interceptor spacecraft can be written as

[0061]

[0062]

[0063] Identify τ dc , Q 2 f and mQ3 are small, the relative motion model can be simplified as

[0064]

[0065]

[0066] definition and right and To estimate, use the following differential tracker

[0067]

[0068] Among them, k z,1 ,k z,2 ,k z,3 ,k z,4 ,k z,5 ,k z,6 are positive constants, z2 and z5 are and The estimated value of have

[0069]

[0070]

[0071] To estimate the inertial parameters θ and m of the interceptor spacecraft, define the following variables:

[0072]

[0073]

[0074] Then there is

[0075] Φ 11 θ=τ c

[0076] mΦ 21 =Q2R1f c

[0077] Further transformation gives

[0078]

[0079]

[0080] definition and are the estimated values ​​of θ and m respectively, let The above formula can be further written as

[0081]

[0082]

[0083] Due to Φ 11 θ=τ c , and perform the inverse linear operator transformation to obtain

[0084] Φ 11 =L(J -1 τ c )

[0085] From the definition of linear operator, we know that when J -1 τ c ≠[0,0,0] T When 11 ≠0 3×6 ,but And x≠0 6×1 ,satisfy Similarly, when Q2R1f c ≠[0,0] T When 21 ≠[0,0] T ,but And x≠0 satisfies So when J -1 τ c ≠[0,0,0] T And Q2R1f c ≠[0,0] T When , the system meets the identification conditions.

[0086] Preferably: In step 5, a fixed-time attitude tracking controller and a terminal guidance law are designed with field of view constraints. Specifically, in terms of attitude tracking control, considering the limited field of view of the interceptor seeker, its maximum field of view angle is θ max , define L1 = [1,0,0] T is the unit vector of the line of sight, then the angle θ between the line of sight and the centerline of the intercepting spacecraft is expressed as

[0087]

[0088] The following variables are defined

[0089]

[0090] When the conversion variable X s When bounded, Θ is always less than the maximum field of view angle θ max .

[0091] Then the following glide surface is designed to achieve the desired attitude σ of the interceptor spacecraft d Tracking

[0092]

[0093] Among them, k1>0, a1∈(1,2), b1∈(0,1) and a1b1>1,

[0094] In the identification of the moment of inertia of the intercepting spacecraft, the following design is made Update rate

[0095]

[0096] in,

[0097]

[0098] Among them, k θ is a positive constant, and is defined for The upper bound d′ m1 The estimated value of is, and the estimated error is The design is as follows The update law

[0099]

[0100] Among them, k d1 ,k d2 All are positive numbers, adaptive fixed-time sliding mode controllers, as follows

[0101]

[0102] Among them, k w1 ,k w2 and k w3 is a positive constant; r w1 >1, r w2 ∈(0,1),

[0103] In terms of interception spacecraft guidance, the following non-singular fixed-time sliding mode surface is designed

[0104]

[0105] Among them, q * is the desired approach angle, k2>0, a2∈(1,2), b2∈(0,1) and a2b2>1. In addition to considering the zeroing of the line of sight angular velocity, the sliding surface S2 also ensures that the line of sight angle can track the desired approach angle q * ,

[0106] In terms of the identification of the interceptor spacecraft mass, the following design is made Update rate

[0107]

[0108] in,

[0109]

[0110] Among them, k m is a positive constant, and is defined As ||Q2R1d f +m Q3|| upper bound d′ m2 The estimation error is Design the following update rate

[0111]

[0112] The adaptive fixed-time sliding mode guidance law is as follows:

[0113]

[0114] Among them, k q1 ,k q2 and k q3 is a positive constant; r q1 >1, r q2 ∈(0,1).

[0115] Compared with the prior art, the present invention has the following beneficial effects:

[0116] The present invention fully considers the attitude-orbit coupling characteristics of the interceptor spacecraft, and on this basis establishes a joint system of attitude and orbit;

[0117] The present invention proposes a non-singular fixed-time sliding mode surface based on a tanh function, which realizes the field of view constraint of the seeker and ensures the fixed-time convergence of the system error.

[0118] The present invention combines parameter identification theory to complete the design of an adaptive attitude and orbit joint control system, ensuring that the interceptor spacecraft can complete the target strike mission when the inertial parameters are unknown, which has strong engineering significance. BRIEF DESCRIPTION OF THE DRAWINGS

[0119] Figure 1 is a flow chart of the present invention;

[0120] Figure 2 It is the coordinate system and vector definition in the process of spacecraft modeling;

[0121] Figure 3 It is a schematic diagram of the attitude control and orbit control engine layout;

[0122] Figure 4 This is the PWPF regulator structure diagram;

[0123] Figure 5 are the simulation results of the guidance law and attitude tracking controller. DETAILED DESCRIPTION

[0124] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0125] Embodiment 1

[0126] See also Figure 1 , the flowchart of the present invention shown in the figure, an integrated method for guidance and control of an exoatmospheric interceptor based on parameter identification, comprises the following steps:

[0127] Step 1: Establish the relative orbital dynamics model of the intercepting spacecraft and the attacking spacecraft;

[0128] Step 2: Establish the relative attitude dynamics model of the spacecraft;

[0129] Step 3: Based on the possible singularity problem of the terminal sliding mode variable structure control method, a non-singular fixed-time sliding mode surface based on the tanh function is designed;

[0130] Step 4: Based on the inertial parameters of the interceptor spacecraft itself and combined with parameter identification theory, an adaptive attitude and orbit joint system is designed;

[0131] Step 5: Based on steps 3 and 4, design a fixed-time attitude tracking controller and terminal guidance law considering field of view constraints.

[0132] In this embodiment, Figure 2 As shown, only when considering the two-body problem of the interceptor and the target spacecraft, the relative motion relationship between the two is established:

[0133]

[0134] Where r = [r x r y r z ] T is the position vector of the interceptor and the target spacecraft in the geocentric inertial coordinate system; are the positions of the interceptor and target spacecraft in the geocentric inertial coordinate system; μ = 3.986 × 10 14 is the gravitational constant; is the mass of the interceptor spacecraft; It is the control force of the interceptor in the inertial frame; The unknown disturbance force acting on the interceptor.

[0135] Define the interceptor's line of sight angle and angular rate relative to the target in the inertial frame:

[0136]

[0137]

[0138] Among them, q1 and q2 are the line of sight inclination and declination expressed in the inertial system respectively; is the relative distance between the interceptor spacecraft and the target spacecraft,

[0139] Define state quantity q = [q1 q2] T , combined with the above formula, we have

[0140]

[0141] in, The complete expression is as follows

[0142]

[0143]

[0144]

[0145]

[0146] Since the interceptor spacecraft has uncontrollable axial characteristics, such as Figure 3 As shown, it is possible to consider equipping the interceptor spacecraft with engines in the lateral and longitudinal directions to achieve control of the sight angle and angular rate.

[0147] Assume that the control force in the interceptor spacecraft coordinate system is f c =[f c1 ,f c2 ] T , converted to the inertial system, we have

[0148] f I =R1f c

[0149] in, R1 is the transformation matrix from the spacecraft body coordinates to the inertial coordinate system. The relative orbital dynamics of the interceptor in the inertial system can be further written as

[0150]

[0151] The modified Rodrigues parameter method is used to describe the interceptor spacecraft's attitude equations:

[0152]

[0153] in, denote the Rodrigues parameter and angular velocity of the interceptor spacecraft in the body coordinate system relative to the inertial coordinate system, is the control torque for the intercepting spacecraft, is the disturbance torque, is the moment of inertia of the intercepting spacecraft, the complete expression is as follows

[0154]

[0155] For the intercepting spacecraft, its moment of inertia has the following properties

[0156] J 12 =J 21 ,J 13 =J 31 ,J 23 =J 32

[0157] According to the line of sight inclination and line of sight deflection during the interception process, the conversion matrix from the inertial system to the line of sight represented by the line of sight angle is compared with the conversion matrix from the inertial system to the line of sight represented by the Euler angle, and the expected attitude angle of the interceptor represented by the Euler angle can be obtained, and then the expected Rodrigues parameter can be obtained, which is expressed as σ d The specific expression is as follows

[0158]

[0159] For the desired angular velocity, we have

[0160]

[0161] According to the definition of Rodrigues parameter and angular velocity, the relative Rodrigues parameter and relative angular velocity can be calculated.

[0162]

[0163] ω e =ω c -R cs ω d

[0164] Among them, R cs is the coordinate transformation matrix from the sight system to the local system,

[0165] Further, the kinematic and dynamic equations of the relative attitude of the intercepting spacecraft in the inertial system are obtained as follows:

[0166]

[0167] in,

[0168] Design a non-singular fixed-time sliding surface based on the tanh function. During the terminal interception process, the intercepting spacecraft will be affected by external interference and model uncertainty. In view of the advantages of strong robustness and insensitivity to system parameters of the sliding mode variable structure control method, a non-singular fixed-time sliding surface based on the tanh function is given here. The specific form is as follows

[0169]

[0170] Among them, x is the state quantity, is a positive constant, a s ∈(1,2), b s ∈(0,1), a s b s >1,

[0171] The derivative of S with respect to time is

[0172]

[0173] When x approaches zero, due to The above formula can be simplified as

[0174]

[0175] Considering a s >1 and a s b s >1, There will be no singular phenomenon, so the designed sliding surface is non-singular.

[0176] When the sliding surface S converges, the system state x also converges. The Lyapunov equation is selected as

[0177]

[0178] The following relationship holds

[0179]

[0180] The convergence of the system state x is proved.

[0181] Combining the parameter identification theory to design the adaptive attitude and orbit joint system, in order to facilitate the subsequent parameter identification and controller design, we first define the variable θ = [J 11 ,J 12 ,J 13 ,J 22 ,J 23 ,J 33 ] T and linear operators At this time, the moment of inertia can be rewritten as follows

[0182]

[0183] Then the relative motion model of the interceptor spacecraft can be written as

[0184]

[0185]

[0186] In order to facilitate system analysis and controller design, it is assumed here that τdc , Q 2 f and mQ3 are small, the relative motion model can be simplified as

[0187]

[0188] definition and right and To estimate, use the following differential tracker

[0189]

[0190]

[0191] Among them, k z,1 ,k z,2 ,k z,3 ,k z,4 ,k z,5 ,k z,6 are positive constants, z2 and z5 are and The estimated value of have

[0192]

[0193]

[0194] In order to estimate the inertial parameters θ and m of the interceptor spacecraft, the following variables are defined:

[0195]

[0196] Then there is

[0197] Φ 11 θ=τ c

[0198] mΦ 21 =Q2R1f c

[0199] Further transformation can be obtained

[0200]

[0201] definition and are the estimated values ​​of θ and m respectively, let The above formula can be further written as

[0202]

[0203] Due to Φ 11 θ=τ c , and perform the inverse linear operator transformation to obtain

[0204] Φ 11 =L(J -1 τ c )

[0205] From the definition of linear operator, we know that when J -1 τ c ≠[0,0,0] T When 11 ≠0 3×6 ,but And x≠0 6×1 ,satisfy Similarly, when Q2R1f c ≠[0,0] T When 21 ≠[0,0] T ,but And x≠0 satisfies So when J -1 τ c ≠[0,0,0] T And Q2R1f c ≠[0,0] T When , the system meets the identification conditions.

[0206] Design a fixed-time attitude tracking controller and terminal guidance law considering the field of view constraint. In terms of attitude tracking control, considering the limited field of view of the interceptor seeker, its maximum field of view angle is θ max , define L1 = [1,0,0] T is the unit vector of the line of sight, then the angle θ between the line of sight and the centerline of the intercepting spacecraft can be expressed as

[0207]

[0208] In order to meet the field of view constraints, based on the idea of ​​state transformation, the following variables are defined

[0209]

[0210] When the conversion variable X s When bounded, Θ is always less than the maximum field of view angle θ max ,

[0211] Then the following glide surface is designed to achieve the desired attitude σ of the interceptor spacecraft d Tracking

[0212]

[0213] Among them, k1>0, a1∈(1,2), b1∈(0,1) and a1b1>1,

[0214] In the identification of the moment of inertia of the intercepting spacecraft, the following design is made Update rate

[0215]

[0216] in,

[0217]

[0218] Among them, k θ is a positive constant. In order to further improve the robustness of the system, we define for The upper bound d′ m1 The estimated value of is, and the estimated error is The design is as follows The update law

[0219]

[0220] Among them, k d1 ,k d2 are all positive numbers. In order to achieve posture tracking control when the moment of inertia is unknown, an adaptive fixed-time sliding mode controller can be designed, as follows:

[0221]

[0222] Among them, k w1 ,k w2 and k w3 is a positive constant; r w1 >1,

[0223] In terms of stability proof, the Lyapunov equation is selected as

[0224]

[0225]

[0226] Through the above formula, we can finally explain that the system state σ e ,ω e It can converge within a fixed time, and the field of view angle always meets the constraints during posture tracking.

[0227] In terms of interception spacecraft guidance, the following non-singular fixed-time sliding mode surface is designed

[0228]

[0229] Among them, q * is the desired approach angle, k2>0, a2∈(1,2), b2∈(0,1) and a2b2>1. In addition to considering the zeroing of the line of sight angular velocity, the sliding surface S2 also ensures that the line of sight angle can track the desired approach angle q * ,

[0230] In terms of the identification of the interceptor spacecraft mass, the following design is made Update rate

[0231]

[0232] in,

[0233]

[0234] Among them, k m is a positive constant. In order to enhance the robustness of the system, we define As ||Q2R1d f +mQ3|| upper bound d′ m2 The estimation error is Design the following update rate

[0235]

[0236] Furthermore, in order to achieve accurate strike of the target by the interceptor spacecraft when the mass is unknown, an adaptive fixed-time sliding mode guidance law is designed as follows:

[0237]

[0238] Among them, k q1 ,k q2 and k q3 is a positive constant; r q1 >1,

[0239] In terms of stability proof, the Lyapunov equation is selected as

[0240]

[0241] Through the above formula, we can finally explain the system state It is stable at a fixed time.

[0242] In order to be closer to the working principle of the interceptor attitude control and orbit control engine, the PWPF regulator is used here to convert the continuous control force and torque into pulse thrust, such as Figure 4As shown in the figure, the principle of the PWPF pulse width modulator is to set different duty cycles to convert the continuous quantity into a pulse thrust that can be directly recognized by the attitude control and orbit control engines, which is closer to the actual engineering application scenario.

[0243] Finally, the effectiveness and superiority of the algorithm are verified through Matlab simulation. Some parameters used in the simulation are as follows:

[0244] The initial mass of the interceptor spacecraft is 25 kg, the thrust of the attitude control engine is 10 N, the specific impulse is 250 s, the thrust of the orbit control engine is 350 N, the specific impulse is 300 s, and the main control parameters are selected as follows: a1 = 1.4, b1 = 0.9, k1 = 1, k q1 =k q2 =2, r q1 =1.4, r q2 =0.7, k d1 =50, k d2 =0.01, k θ =15, a2=1.2, b2=0.9, k2=0.5, k w1 =k w2 =4, r w1 =1.1, r w2 =0.8, k d3 =0.1, k d4 =0.001, k m =10; in the attitude control engine, the PWPF parameter is set to: K m,τ1 =10, K m,τ2 =K m,τ3 =1,T m,τ1 =T m,τ2 =T m,τ3 =0.05, U off,τ1 =0.005, U on,τ1 =0.008, U off,τ2 =U off,τ3 =0.01, U on,τ2 =U on,τ3 =0.04, K m,f1 =K m,f2 =1,T m,f1 =T m,f2 =1,U off,f1 =U off,f2 =2-0.02t, U on,τ2 =U on,τ3 =20-0.4t.

[0245] According to the above parameters, we get Figure 5From the simulation results in , it is not difficult to see that even when the system parameters are unknown, the interceptor spacecraft can intercept the target with a small miss, which can meet the actual mission requirements. In terms of attitude tracking, the algorithm in the present invention can also ensure a faster convergence speed and higher convergence accuracy, and the field of view constraint of the seeker is always satisfied. Figure 5 In (g)-(h), after the attitude control torque and orbit control force are discretized by PWPF, the three attitude control engines are turned on twice, and the two orbit control engines are turned on 30 and 37 times respectively, consuming a total of 1.84 kg of fuel, which is Figure 5 From (i) to (j), it can be seen that the inertial parameter identification strategy in the present invention can achieve a good estimation of the mass and moment of inertia of the interception spacecraft. In summary, the effectiveness of the control algorithm in the present invention is verified.

[0246] It should be noted that, in this article, relational terms such as first and second, etc. are only used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Moreover, the terms "include", "comprise" or any other variants thereof are intended to cover non-exclusive inclusion, so that a process, method, article or device including a series of elements includes not only those elements, but also other elements not explicitly listed, or also includes elements inherent to such process, method, article or device.

[0247] Although embodiments of the present invention have been shown and described, it will be appreciated by those skilled in the art that various changes, modifications, substitutions and variations may be made to the embodiments without departing from the principles and spirit of the present invention, and that the scope of the present invention is defined by the appended claims and their equivalents.

Claims

1. An integrated method for guidance and control of an exoatmospheric interceptor based on parameter identification, characterized in that: The steps include: Step 1: Establish the relative orbital dynamics model of the intercepting spacecraft and the attacking spacecraft; Step 2: Establish the relative attitude dynamics model of the spacecraft; Step 3: Based on the possible singularity problem of the terminal sliding mode variable structure control method, a non-singular fixed-time sliding mode surface based on the tanh function is designed; Step 4: Based on the inertial parameters of the interceptor spacecraft itself and combined with parameter identification theory, an adaptive attitude and orbit joint system is designed; Step 5: Based on steps 3 and 4, design a fixed-time attitude tracking controller and terminal guidance law considering field of view constraints.

2. The method for integrating guidance and control of an exoatmospheric interceptor based on parameter identification according to claim 1, characterized in that: The relative motion relationship between the interceptor and the target spacecraft in step 1 can be expressed as: Where r = [r x r y r z ] T is the position vector of the interceptor and the target spacecraft in the geocentric inertial coordinate system; are the positions of the interceptor and target spacecraft in the geocentric inertial coordinate system; μ = 3.986 × 10 14 is the gravitational constant; is the mass of the interceptor spacecraft; It is the control force of the interceptor in the inertial frame; The unknown disturbance force on the interceptor; Define the interceptor's line of sight angle and angular rate relative to the target in the inertial frame: Among them, q1 and q2 are the line of sight inclination and declination expressed in the inertial system respectively; is the relative distance between the interceptor spacecraft and the target spacecraft; Define state quantity q = [q1 q2] T , combined with the above formula, we have in, The complete expression is as follows Let the control force on the interceptor spacecraft body coordinate system be f c =[f c1 ,f c2 ] T , converted to the inertial system, we have f I =R1f c in, R1 is the transformation matrix from the spacecraft body coordinates to the inertial coordinate system. The relative orbital dynamics of the interceptor in the inertial system can be further written as 3. The integrated method for guidance and control of an exoatmospheric interceptor based on parameter identification according to claim 2 is characterized in that: In step 2, the modified Rodrigues parameter method is used to describe the attitude operation equation of the interceptor spacecraft: in, denote the Rodrigues parameter and angular velocity of the interceptor spacecraft in the body coordinate system relative to the inertial coordinate system, is the control torque of the intercepting spacecraft, is the disturbance torque, is the moment of inertia of the intercepting spacecraft, the complete expression is as follows For the intercepting spacecraft, its moment of inertia has the following properties I 12 =J 21 ,J 13 =J 31 ,J 23 =J 32 According to the line of sight inclination and line of sight deflection during the interception process, the conversion matrix from the inertial system to the line of sight represented by the line of sight angle is compared with the conversion matrix from the inertial system to the line of sight represented by the Euler angle, and the expected attitude angle of the interceptor represented by the Euler angle can be obtained, and then the expected Rodrigues parameter can be obtained, which is expressed as σ d ; The specific expression is as follows For the desired angular velocity, we have According to the definition of Rodrigues parameter and angular velocity, the relative Rodrigues parameter and relative angular velocity can be calculated. oh e =ω c -R cs oh d Among them, R cs is the coordinate transformation matrix from the sight system to this system; Further, the kinematic and dynamic equations of the relative attitude of the intercepting spacecraft in the inertial system are obtained as follows: in, 4. The integrated method for guidance and control of an exoatmospheric interceptor based on parameter identification according to claim 3 is characterized in that: The non-singular fixed-time sliding surface based on the tanh function designed in step 3 is as follows Among them, x is the state quantity, is a positive constant, a s ∈(1,2), b s ∈(0,1), a s b s >1; The derivative of S with respect to time is When x approaches zero, due to The above formula can be simplified as Considering a s >1 and a s b s >1, There will be no singular phenomenon, so the designed sliding surface is non-singular. When the sliding surface S converges, the system state x also converges. The Lyapunov equation is selected as The following relationship holds The convergence of the system state x is proved.

5. The integrated method for guidance and control of an exoatmospheric interceptor based on parameter identification according to claim 4, characterized in that: In step 4, the adaptive attitude and trajectory joint system is designed in combination with the parameter identification theory, specifically, the variable θ is defined as [J 11 ,J 12 ,J 13 ,J 22 ,J 23 ,J 33 ] T and linear operators At this time, the moment of inertia can be rewritten as follows Then the relative motion model of the intercepting spacecraft can be written as Identify τ dc , Q 2 f and mQ3 are small, the relative motion model can be simplified as definition and right and To estimate, use the following differential tracker Among them, k z,1 ,k z,2 ,k z,3 ,k z,4 ,k z,5 ,k z,6 are positive constants, z2 and z5 are and The estimated value of have To estimate the inertial parameters θ and m of the interceptor spacecraft, define the following variables: Then there is F 11 θ=τ c mΦ 21 =Q2R1f c Further transformation gives definition and are the estimated values ​​of θ and m respectively, let The above formula can be further written as Due to Φ 11 θ=τ c , and perform the inverse linear operator transformation to obtain F 11 =L(J -1 t c ) From the definition of linear operator, we know that when J -1 τ c ≠[0,0,0] T When 11 ≠0 3×6 ,but And x≠0 6×1 ,satisfy Similarly, when Q2R1f c ≠[0,0] T When 21 ≠[0,0] T ,but And x≠0 satisfies So when J -1 τ c ≠[0,0,0] T And Q2R1f c ≠[0,0] T When , the system meets the identification conditions.

6. The integrated method for guidance and control of an exoatmospheric interceptor based on parameter identification according to claim 5, characterized in that: In step 5, a fixed-time attitude tracking controller and a terminal guidance law are designed with field of view constraints. Specifically, in terms of attitude tracking control, considering the limited field of view of the interceptor seeker, its maximum field of view angle is θ max , define L1 = [1,0,0] T is the unit vector of the line of sight, then the angle θ between the line of sight and the centerline of the intercepting spacecraft is expressed as The following variables are defined When the conversion variable X s When bounded, θ is always less than the maximum field of view angle θ max ; Then the following glide surface is designed to achieve the desired attitude σ of the interceptor spacecraft d Tracking Among them, k1>0, a1∈(1,2), b1∈(0,1) and a1b1>1, In the identification of the moment of inertia of the intercepting spacecraft, the following design is made Update rate in, Among them, k θ is a positive constant, and is defined for The upper bound d′ m1 The estimated value of is, and the estimated error is The design is as follows The update law Among them, k d1 ,k d2 All are positive numbers, adaptive fixed-time sliding mode controllers, as follows Among them, k w1 ,k w2 and k w3 is a positive constant; r w1 >1, In terms of interception spacecraft guidance, the following non-singular fixed-time sliding mode surface is designed Among them, q * is the desired approach angle, k2>0, a2∈(1,2), b2∈(0,1) and a2b2>1. In addition to considering the zeroing of the line of sight angular velocity, the sliding surface S2 also ensures that the line of sight angle can track the desired approach angle q * , In terms of the identification of the interceptor spacecraft mass, the following design is made Update rate in, Among them, k m is a positive constant, and is defined As ||Q2R1d f +mQ3|| upper bound d′ m2 The estimation error is Design the following update rate The adaptive fixed-time sliding mode guidance law is as follows: Among them, k q1 ,k q2 and k q3 is a positive constant; r q1 >1,

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