Cooperative guidance method for low-cost strapdown seeker

Abstract

The invention relates to a collaborative guidance method for a low-cost strapdown seeker, and belongs to the technical field of collaborative guidance. Comprising the following steps: acquiring a sight angle under an inertial system based on a body sight angle under a missile system measured by a strapdown seeker; based on the line-of-sight angle under the inertial system, estimating the line-of-sight angle and the line-of-sight angle rate by using an alpha-beta filtering algorithm; considering the terminal falling angle constraint, and establishing a space cooperative guidance law based on the estimated line-of-sight angular rate; time constraint is considered, and time cooperative guidance is established based on the estimated line-of-sight angular rate; and selecting a space cooperative guidance law or time cooperative guidance according to different scenes to perform space cooperative guidance or time cooperative guidance on the strapdown aircraft. Compared with the prior art, the filtering algorithm is used for estimating the sight angle and the sight angle rate, the structure is simple, complex matrix operation is not involved, the engineering applicability is high, the collaborative guidance algorithm is designed, collaborative strike on the target is achieved, and the aircraft cluster efficiency is improved.

Description

Technical Field

[0001] The invention belongs to the technical field of cooperative guidance, and in particular relates to a cooperative guidance method for a low-cost strapdown seeker. Background Art

[0002] As an important subsystem of a self-seeking missile, the seeker plays a key role in the entire guidance and control system. In the terminal guidance stage of the missile trajectory, the seeker can directly search, calibrate and automatically track the predetermined target. The execution process is fast and accurate. The guidance information output by the seeker is output to the guidance and control system, and finally guides the missile to hit the target. The performance of the seeker directly affects the hit accuracy and combat effectiveness of the self-seeking missile. With the continuous development and progress of precision-guided weapon related technologies, the study of seekers has become an important topic for researchers in various countries around the world.

[0003] Due to the constraints of the mechanical frame, the traditional platform-type seeker has the disadvantages of complex structure, high cost, and large measurement error. With the continuous development of low manufacturing cost and miniaturized seekers, the research of strapdown seekers has received more attention. The strapdown seeker has a simple structure design, small size, light weight, high reliability, and is insensitive to missile overload, which improves the accuracy of the output signal. However, due to the special structure of the strapdown seeker, its engineering application is subject to certain restrictions. The line of sight angular rate required for guidance cannot be directly obtained, and a relatively large measurement noise is introduced into the guidance signal, affecting the accuracy of the guidance signal, and the coupling of the line of sight angle signal and the missile attitude angle information. As a new type of tactical missile, the low-cost strapdown missile has a curved and changeable flight trajectory, a low flight altitude, and is easily affected by dense atmospheric interference. In order to ensure a fast and accurate response to the guidance command, the control system needs to have a high anti-interference ability.

[0004] In order to better apply the strapdown seeker to engineering applications, give full play to its advantages, improve its performance utilization and guidance accuracy, it is particularly important to study and design more efficient guidance control methods to eliminate the above-mentioned constraints. Summary of the invention

[0005] The technical problems to be solved by the present invention are:

[0006] The strapdown seeker is fixedly connected to the missile body, the line of sight angular rate required for guidance cannot be directly obtained, and a relatively large measurement noise is introduced into the guidance signal. The line of sight angle signal measured by the seeker is coupled with the missile body attitude angle signal, which brings errors to the collaborative guidance. In order to avoid the shortcomings of the prior art, the present invention provides a collaborative guidance method for a low-cost strapdown seeker.

[0007] In order to solve the above technical problems, the technical solution adopted by the present invention is:

[0008] A collaborative guidance method for a low-cost strapdown seeker, characterized by comprising:

[0009] The line of sight angle in the inertial system is obtained based on the stereoscopic line of sight angle in the missile system measured by the strapdown seeker;

[0010] Based on the sight angle in the inertial system, the sight angle and sight angle rate are estimated using the α-β filtering algorithm;

[0011] Considering the terminal landing angle constraint, a space cooperative guidance law is established based on the estimated line-of-sight angular velocity.

[0012] Considering the time constraint, the temporal cooperative guidance is established based on the estimated line-of-sight angular rate;

[0013] According to different scenarios, the space cooperative guidance law or the time cooperative guidance is selected to respectively perform space cooperative guidance or time cooperative guidance on the strapdown aircraft.

[0014] A further technical solution of the present invention is as follows: the stereoscopic line of sight angle under the missile system measured by the strapdown seeker is used to obtain the line of sight angle under the inertial system, specifically:

[0015] Based on the volumetric line of sight angle measured by the strapdown seeker in the missile system, the two are decoupled by line of sight angle reconstruction, and the relationship between the line of sight azimuth and elevation angle in the inertial system is obtained:

[0016]

[0017] Where R 11 , R 12 , R 13 , R 21 , R 22 , R 23 , R 31 , R 32 , R 33 is the coordinate system transformation coefficient, q a ,q b is the sight elevation angle and sight azimuth angle under the missile system, q λ ,q γ are the line of sight azimuth and line of sight altitude angle in the inertial system.

[0018] A further technical solution of the present invention is as follows: the sight angle based on the inertial system uses a filtering algorithm to estimate the sight angle and the sight angle rate, and the formula used is:

[0019]

[0020] Among them, α and β are filter gains, T is the filtering algorithm time interval, is the estimated value of the system state, is the velocity estimate, Is the system at t k+1 The state value at the moment, Yes k+1 The speed of the system at a given moment.

[0021] A further technical solution of the present invention is: considering the terminal landing angle constraint, the space cooperative guidance law is established based on the estimated line of sight angular velocity, specifically:

[0022]

[0023] Where N is the proportional coefficient, v is the aircraft speed, is the line of sight angular velocity, a B is the falling angle constraint bias term. A further technical solution of the present invention is: the falling angle constraint bias term a B Calculated according to the following formula:

[0024] a B =-(σ-Nq+(N-1)σ d )v / t go

[0025] Among them, σ is the ballistic angle of the aircraft, σ d It is the expected landing angle.

[0026] A further technical solution of the present invention is: considering the time constraint, the time coordinated guidance is established based on the estimated line of sight angular velocity, specifically:

[0027]

[0028] Among them, t go is the remaining attack time, R go is the remaining flight distance, ε t is the attack time error.

[0029] A computer system, characterized in that it comprises: one or more processors, and a computer-readable storage medium for storing one or more programs, wherein when the one or more programs are executed by the one or more processors, the one or more processors implement the above-mentioned method.

[0030] A computer-readable storage medium, characterized in that it stores computer-executable instructions, which are used to implement the above method when executed.

[0031] A computer program product, characterized by comprising computer executable instructions, wherein the instructions are used to implement the above method when executed.

[0032] The beneficial effects of the present invention are:

[0033] The invention provides a collaborative guidance method for a low-cost strapdown seeker. Based on the volume line-of-sight angle and the angular velocity of a missile body measured by the strapdown seeker, the two are decoupled by reconstructing the line-of-sight angle, and then the line-of-sight angle and the line-of-sight angular velocity are estimated by filtering, and the line-of-sight angular velocity information required by the seeker is output. On the basis of the traditional method, the invention introduces a time collaborative bias term and a terminal angle-of-fall constraint bias term, and designs a multi-missile collaborative guidance law considering angle-of-fall constraints and time constraints, so that each low-cost small strapdown aircraft can accurately hit the target at the same time and meet different expected angle-of-fall constraints, thereby achieving a coordinated attack on the target and improving the efficiency of the aircraft cluster.

[0034] Compared with the prior art, the present invention uses a filtering algorithm to estimate the line of sight angle and line of sight angular rate. It has a simple structure, does not involve complex matrix operations, has strong engineering applicability, and designs a collaborative guidance algorithm to achieve coordinated strikes on targets and improve the effectiveness of aircraft clusters. BRIEF DESCRIPTION OF THE DRAWINGS

[0035] The drawings are only for the purpose of illustrating particular embodiments and are not to be considered limiting of the present invention. Like reference symbols denote like components throughout the drawings.

[0036] Figure 1 Flowchart of collaborative guidance algorithm design for low-cost strapdown seeker;

[0037] Figure 2 Schematic diagram of strapdown seeker line of sight angular rate extraction;

[0038] Figure 3 A diagram showing the comparison of the azimuth sight angles in the example of the present invention;

[0039] Figure 4 A diagram showing the comparison of azimuth line of sight angular velocity in an example of the present invention;

[0040] Figure 5 A comparison diagram of high and low sight angles in an example of the present invention;

[0041] Figure 6 A comparison of high and low sight line angular rates in an example of the present invention;

[0042] Figure 7 Schematic diagram of the relative motion relationship between the missile and the target;

[0043] Figure 8 The three-dimensional trajectory of the missile in the case of a stationary target in the example of the present invention;

[0044] Fig. 9 The XZ plane trajectory of the missile in the case of a stationary target in the example of the present invention;

[0045] Fig.10The remaining flight time of the missile in the case of a stationary target in the example of the present invention;

[0046] Fig.11 In the present invention, the missile is longitudinally overloaded in the case of a stationary target;

[0047] Fig.12 In the example of the present invention, the missile is lateral overloaded in the case of a stationary target;

[0048] Fig.13 The missile trajectory deviation angle in the case of a stationary target in the example of the present invention;

[0049] Fig.14 The relative distance between the projectile and the target in the case of a stationary target in the example of the present invention;

[0050] Fig.15 The three-dimensional trajectory of the target in the case of cosine maneuver in the example of the present invention;

[0051] Fig.16 XZ plane trajectory of the target in the case of cosine maneuver in the example of the present invention;

[0052] Fig.17 The remaining flight time of the target in the cosine maneuver case in the example of the present invention;

[0053] Fig.18 The lateral overload in the case of the target cosine maneuver in the example of the present invention;

[0054] Fig.19 Longitudinal overload in the case of target cosine maneuver in the example of the present invention;

[0055] Fig. 20 The trajectory deviation angle of the target in the cosine maneuvering case in the example of the present invention;

[0056] Fig.21 The relative distance between the projectile and the target when the target cosine maneuvers in the example of the present invention. DETAILED DESCRIPTION

[0057] In order to make the purpose, technical solutions and advantages of the present invention more clearly understood, the present invention is further described in detail below in conjunction with the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present invention and are not intended to limit the present invention. In addition, the technical features involved in the various embodiments of the present invention described below can be combined with each other as long as they do not conflict with each other.

[0058] The present invention provides a collaborative guidance method for a low-cost strapdown seeker, comprising the following contents:

[0059] (1) Based on the stereoscopic line-of-sight angle measured by the strapdown seeker, the two are decoupled by line-of-sight angle reconstruction, and then the line-of-sight angle and line-of-sight angular rate are estimated by filtering;

[0060] (2) Considering the terminal angle constraint, adding the angle constraint bias term, and designing the biased proportional guidance law with the hit angle constraint;

[0061] (3) Considering the terminal time constraints, adding time coordination terms, and designing the guidance law with time coordination terms;

[0062] Specific research contents include:

[0063] (1) During the guidance process, the strapdown seeker used in the present invention will be subjected to different degrees of signal amplification, delay and dead zone systems, and the guidance signal finally calculated and outputted may be distorted to varying degrees. Therefore, during the final flight of the aircraft, the terminal needs to perform signal filtering and amplitude limiting processing on the guidance information output by the strapdown seeker of the present invention to obtain a stable guidance signal. In order to meet the stability of the closed-loop guidance loop required by the present invention and the conditions for guidance accuracy, and then output the line of sight angular velocity information required by the strapdown seeker, an α-β filtering algorithm is used as the filtering algorithm of the strapdown seeker of the present invention.

[0064] Consider Figure 1 The strapdown seeker line of sight angle rate extraction schematic diagram shown in FIG. According to the above-mentioned decoupling method of missile body motion, based on the volume line of sight angle q measured by the strapdown seeker a ,q b , the inertial sight angle q is obtained through the sight angle reconstruction process λ ,q γ , after differentiation and noise filtering, the line of sight angular rate for proportional guidance can be obtained

[0065] Since the line of sight angle q output by the strapdown seeker a ,q b The missile body’s attitude motion information and a large amount of background noise and thermal noise are coupled in the observation equation. In order to establish the strapdown seeker’s inertial line of sight angular rate extraction model, the noisy seeker output q a ,q b As the observed quantity, the azimuth sight line angular rate and the high and low sight line angular rate as well as the azimuth sight line angle and the high and low sight line angle are selected as the state variables in the state equation.

[0066] In the sight coordinate system, the missile-target vector r has the following relationship:

[0067]

[0068] Among them, the rotation angular velocity of the sight coordinate system relative to the inertial coordinate system is

[0069]

[0070] The approach speed between the aircraft and the target can be expressed as

[0071]

[0072] Differentiating the approach velocity, the relative acceleration between the aircraft and the target is:

[0073]

[0074] Assuming that the target does not maneuver in the line of sight coordinate system, the following differential equations are obtained for the missile-target distance and the missile-target line of sight angular rate:

[0075]

[0076] make Then the following state equation of strapdown seeker line-of-sight angular rate decoupling is obtained:

[0077]

[0078] It can be seen from the above formula that the state equation of the system has strong nonlinearity.

[0079] The target coordinates measured by the strapdown seeker are the same in the line of sight and the stereoscopic line of sight. Assume it is [r 0 0] T , then the representations of the target vector in the two coordinate systems are:

[0080]

[0081] According to the following conversion relationship between the projectile system and the inertial system:

[0082]

[0083] Among them, γ, θ, and ψ are the roll angle, pitch angle, and yaw angle.

[0084] The relationship between the stereoscopic sight angle, inertial sight angle and missile attitude is obtained:

[0085] M(q a )L(q b )=M(q γ )L(q λ )[K(γ)M(θ)L(ψ)] -1 (10)

[0086] Right now:

[0087]

[0088] Arrange the above formula to get the relationship between the target sight angle in the stereoscopic sight system:

[0089]

[0090] In the above formula:

[0091]

[0092] Let y = [q a q b ] T , then the following measurement equation for the strapdown seeker line-of-sight angular rate decoupling is obtained:

[0093]

[0094] As a measurement component of the self-seeking guidance system, the main function of the seeker is to autonomously search, identify and track the target. However, the optical system and stabilization mechanism of the seeker do not work under ideal conditions, and the missile body will also be disturbed by the external airflow, resulting in the inevitable introduction of interference noise into the measurement information of the seeker. The traditional platform-type seeker has a small measurement field of view, and the observation noise can usually be assumed to be a Gaussian white noise sequence with an expected value of zero; for the strapdown seeker fixedly installed with the missile body, its observation field of view is larger, and the optical axis of the seeker moves with the posture movement of the missile body, and more complex colored noise will be introduced into its measurement information. In the present invention, Gaussian white noise is set to reflect the noise characteristics of the strapdown seeker, and subsequent line of sight angular rate extraction and simulation are performed.

[0095] Let f n (x) is the probability density function of Gaussian white noise with normal distribution, expected value E(x) = μ, variance value

[0096]

[0097] The present invention uses α-β filtering to estimate the sight angle and sight angular velocity. The algorithm principle is as follows:

[0098] 1) Input data: The α-β filtering algorithm uses two sets of input data, namely, measurement values ​​and prior estimates.

[0099] 2) Prior estimation: The prior estimation is the initial estimation or the last estimation result of the system state, which is usually based on the system model or historical data.

[0100] 3) State estimation: The α-β filtering algorithm combines the measured value and the prior estimate to obtain the state estimate. The state estimate is an estimate or prediction of the actual system state.

[0101] The algorithm steps are as follows:

[0102] 1) Assume that the estimated values ​​of the line-of-sight angular velocity in the altitude and azimuth directions are q y , q z , Establishing the system state equation

[0103]

[0104] 2) Select the time interval of the filtering algorithm as T s , discretize the above system state equation into:

[0105]

[0106] In the α-β filtering process, according to the estimated value q at the previous moment y,k-1，k-1 , q z,k-1，k-1 , Predict the current value q y,k，k-1 , q z,k，k-1 , Then according to the measured value q at the current moment y,k ,q z.k For the measured value q y,k，k-1 , q z,k，k-1 , Make corrections to get the estimated value q at the current moment y,k,k , q z,k,k ,

[0107] 3) The α-β filter estimate finally obtained by the filtering algorithm is:

[0108]

[0109] Where: α and β are filter gains, which can be adjusted according to different situations.

[0110] By considering the trade-off between prior estimates and measured values, the α-β filtering algorithm can provide relatively smooth state estimates and reduce jitter caused by measurement noise. Moreover, due to the existence of β gain, the α-β filtering algorithm has a faster response speed and can adjust the changes in system states more quickly. The α-β filtering algorithm can provide smooth and fast-responding state estimation results by combining measured values ​​and prior estimates and adjusting α gain and β gain.

[0111] The decoupling algorithm established by the present invention is used to model and calculate the line of sight angle information of the aircraft in the inertial system, and α-β filtering is performed. At the same time, the results are introduced into the six-degree-of-freedom simulation model of the aircraft. The simulation results Figures 3 to 6As shown. The red line is the value of the true angular rate, which is the result of directly differentiating the angular rate; the blue line represents the result after α-β filtering. From the results, we can see that after α-β filtering, the line of sight angle and angular rate are relatively smooth. The maximum error of the extracted line of sight angular rate is 0.0286° / s, and the average error is 0.0061° / s. The maximum error of the high and low angular rates is 0.1834° / s, and the average error is 0.0308° / s. The estimation accuracy of the α-β filtering algorithm can meet the line of sight angular rate required for proportional guidance.

[0112] (2) The biased proportional guidance law with impact angle constraint is designed as follows:

[0113] For point targets, multiple strapdown aircraft are needed to conduct penetration attacks and quickly destroy the target from multiple angles. Figure 7 The relative motion relationship between the aircraft and the target is shown in the figure. T is the target. Assuming the target is stationary, (x T ,z T ) is the position of the target; M is the aircraft, (x, z) is the instantaneous position of the aircraft, v is the tangential velocity of the aircraft, assuming v is a constant, a n is the normal acceleration of the aircraft, σ is the ballistic angle of the aircraft, and it is stipulated that σ is positive when the velocity vector points above the AX axis, otherwise it is negative; R is the relative distance between the missile and the target; q is the sight angle of the missile and the target, and it is stipulated that q is positive when it rotates counterclockwise from the horizontal reference line to the sight of the missile and the target, and otherwise it is negative; η is the lead angle of the aircraft velocity vector, and it is stipulated that η is positive when it rotates counterclockwise from the aircraft velocity vector to the sight of the missile and the target, and otherwise it is negative.

[0114] The relative motion equation between projectile and target is:

[0115]

[0116] The equation of motion of the aircraft is

[0117]

[0118] Since η=q-σ, then

[0119]

[0120] In order to enable the aircraft to hit the target at the desired angle, a method of adding a bias term to the proportional guidance is adopted. Its general form is:

[0121]

[0122] Where: N is the proportionality coefficient; a B is the bias term.

[0123] Application scenarios of space collaborative guidance: For point targets, multiple strapdown aircraft are required to conduct penetrating attacks and quickly destroy enemy targets through multi-angle encirclement.

[0124] Substituting (22) into (21), we get

[0125]

[0126] According to the definition of derivative, we know that within a short period of time Δt,

[0127] (σ t+Δt -σ t ) / Δt=N(q t+Δt -q t ) / Δt+a B / v (24)

[0128] Where: subscripts t and t+Δt represent the initial and final time of the Δt time period, respectively. B If it is a constant in each Δt time period, then Δt is expanded to the total flight time t f Available

[0129] σ f -σ 0 =N(q f -q 0 )+a B t f / v (25)

[0130] Where: subscripts 0 and f represent the initial and final moments of the entire flight process, respectively. When the aircraft hits the target, R(t f )=0, according to (19) we can get

[0131] σ f =q f (26)

[0132] Let σ d is the expected falling angle, let

[0133] σ f =σ d (27)

[0134] That is, we hope that the trajectory angle of the aircraft at the terminal moment can reach the expected landing angle. Then, by rearranging formula (25), we can get

[0135] a B =-(σ 0 -Nq 0 +(N-1)σ d )v / t f (28)

[0136] Due to tf It can be regarded as the remaining attack time of the aircraft at the initial moment of flight, so the bias term a B The current time parameters can be expressed as

[0137] a B =-(σ-Nq+(N-1)σ d )v / t go (29)

[0138] Where: t go Indicates the remaining attack time of the aircraft at the current moment.

[0139] In this guidance law, the estimation of the remaining attack time does not take into account the influence of the normal acceleration command on the flight trajectory curvature and the remaining flight time. If this guidance law is only applied to the guidance system of a single aircraft, and the system has no requirements for the attack time, then this guidance law has no significant effect on the aircraft's miss distance and the accuracy of the hit angle control. The present invention intends to extend this biased proportional guidance law to the collaborative guidance of networked aircraft. The collaborative guidance system requires that the attack time of each aircraft must be consistent, so the guidance system has higher requirements for the estimation accuracy of the remaining flight time. The following introduces the precise estimation of the remaining flight time.

[0140] Assume the position of the aircraft is (x,z) and the position of the target is (x T ,z T ), then the sight angle can be approximated as

[0141] q=(z T -z) / R (30)

[0142] From R = v c t go , and differentiate the above formula with respect to time to obtain

[0143]

[0144] Under linear conditions, q and η are both small angles, so we can approximate v c =v, Combining the above formula, we can get

[0145]

[0146] To facilitate the solution, the whole process of the attack is considered in reverse, and the attack is regarded as a low-cost small strapdown aircraft launched from the target position and flying to the launch origin. go =(R 0 -r) / v c ,r∈[0,R 0 ], where R 0 is the distance between the target and the launch origin, converting the above formula into a function with r as the independent variable:

[0147]

[0148] According to the initial conditions: r = 0, z(0) = z 0 =z t , Solving the above Euler-Cauchy equation yields

[0149]

[0150] Taking the derivative of the above formula with respect to r, we can get

[0151]

[0152] The length of the aircraft trajectory is Under linear conditions, we can approximate x(r)=r. is a small quantity, according to Taylor series we can get

[0153]

[0154] Since the sight angle q is small, s≈Vt go , Substituting the formula into

[0155]

[0156] (3) The guidance law with time coordination term is designed as follows:

[0157] For important targets, multiple strapdown aircraft are required to attack simultaneously and destroy the target quickly in a short time. The kinematic equation of a low-cost small strapdown aircraft is shown in Equation (20), and the initial and terminal conditions are written as

[0158]

[0159] Among them, T f The terminal time.

[0160] The steering problem with terminal time constraints can be transformed into the optimal control problem of the following linearized system:

[0161]

[0162] Where x is an independent parameter, u B ,u F is the control variable, and σ is assumed to be small, and each variable is dimensionless.

[0163]

[0164] Among them, T, X, Z, a B ,a F Is a dimensioned quantity.

[0165] The cost function is:

[0166]

[0167] The boundary conditions of each state variable are also dimensionless:

[0168]

[0169] Since v is a constant, the constraint on the attack time can be expressed as follows:

[0170]

[0171] This linear optimization problem can be solved analytically by applying the maximum principle. The Hamiltonian function of this problem can be written as

[0172]

[0173] Among them, λ z , σ is the undetermined Lagrange multiplier.

[0174] Optimal conditions The optimal acceleration instruction is given as follows

[0175] u B =-λ σ =-ν z (x f -x) (45)

[0176] Integrating equation (39) yields

[0177]

[0178] Therefore, the feedback guidance command at the initial moment can be expressed as u F Function:

[0179]

[0180] From this we can get

[0181]

[0182] in, x go ≡x f -x, and z go ≡z f -z.

[0183] The heading angle can be expressed as a polynomial of x, as shown below

[0184] θ(η)=αη2 +βη+γ (49)

[0185] in,

[0186] η∈[x,x f ]

[0187]

[0188] β=u F -ν z x f

[0189]

[0190] Taylor expansion of equation (43) and ignoring higher-order terms yields

[0191]

[0192] in, is the expected guidance time, c n is 2 The polynomial coefficients of (x) are as follows

[0193]

[0194] c 2 =βγ

[0195] c 3 =(β 2 +2αγ)

[0196] c 4 =αβ

[0197]

[0198] Use it F Expressed as

[0199]

[0200] Let u in formula (51) F The remaining flight time estimate of a low-cost small strapdown aircraft can be obtained by taking it as 0, as shown below

[0201]

[0202] The solution of formula (51) is

[0203]

[0204] where ε t is the attack time error, Select the one that satisfies ε t= 0 when u F =0 condition, the additional instruction is

[0205]

[0206] Combining equation (48) with equation (52) we get the acceleration command as follows:

[0207]

[0208] Rearranging the above formula, we can get the expression of the time coordinated guidance law:

[0209]

[0210] Application scenarios of time-coordinated guidance: For important targets, multiple strapdown aircraft are required to attack simultaneously and quickly destroy enemy targets in a short time.

[0211] At the same time, in order to achieve low-cost aircraft coordinated attack targets and improve the attack accuracy of low-cost aircraft, this project conducted simulation verification on the situation of multiple aircraft coordinated attack on ground moving or stationary targets, and verified the availability of angle-constrained and time-coordinated guidance laws.

[0212] The simulation verification of multiple aircraft coordinated attack on ground stationary targets is as follows: This simulation adopts three aircraft coordinated attack on a target, and the coordinated attack simulation verification is carried out when the target is stationary. The specific initial conditions of the simulation are as follows:

[0213] Table 1 Simulation initial conditions

[0214] X(m) Y(m) Z(m) Ballistic inclination (°) Ballistic angle (°) Speed ​​(m / s) Aircraft 1 0 1000 1000 0 20 225 Flying Machine 2 0 1000 600 0 -30 230 Aircraft 3 0 1000 200 0 0 245

[0215] When the target is stationary, the position of the target point is (2000, 0, 1340), and the trajectory deflection constraints of the three aircraft are 30°, -45°, and -30° respectively. The simulation step size is selected as 0.01s, and the simulation solver selects the Runge-Kutta method.

[0216] It can be seen from the simulation diagram that the three aircrafts hit the stationary target in 9.63 seconds, and the relative distance between the missile and the target can converge to 0, achieving a relatively good coordinated attack effect: Figure 8 and Fig. 9 It can be seen that the three aircraft attacked the target under different initial conditions and finally achieved coordination and hit the target at the same time; especially Fig. 9 It can be seen that in the XZ plane, the three aircraft hit the target in three directions according to the constraint angles; Fig.10 It can be seen that during the coordinated attack of the aircraft, the remaining attack time of the three aircraft tends to be consistent within about 6 seconds; Fig.11 and Fig.12It can be seen that the absolute values ​​of the longitudinal overload and lateral overload of the aircraft are kept within 15, which is in line with the aircraft's capabilities; Fig.13 It can be seen that the three aircraft converge to the constraint angles of 30°, -45°, and -30° respectively.

[0217] In the case of target cosine maneuver, the target starting point is (2000,0,1340), moving at a speed of 54km / h along the X-axis and an amplitude of 3m / s along the Z-axis. 2 , a cosine maneuver with a period of 2.5π is used to simulate ground vehicles. The simulation step size is selected as 0.01s, and the simulation solver selects the Runge-Kutta method.

[0218] From the above simulation diagram, we can see that the three aircrafts hit the stationary target in coordination within 11.04s. The minimum relative distance between the three missiles and the target was 2.0784m, which achieved a relatively good strike effect. Fig.15 and Fig.16 It can be seen that the three aircraft attacking the cosine maneuvering target under different initial conditions can achieve coordinated hitting of the cosine maneuvering target, but due to the influence of the target maneuvering, the coordinated effect is not as good as that of attacking stationary and linear maneuvering targets; Fig.16 It can be seen that all three vehicles successfully hit the target in the XZ plane; Fig.17 Compared with the attack on stationary targets, the remaining attack time tends to be consistent with the later ones; Fig.18 and Fig.19 It can be seen that the absolute values ​​of the longitudinal overload and lateral overload of the aircraft are kept within 15; Fig. 20 The ballistic deviation angle is within a reasonable range.

[0219] The above description is only a specific implementation mode of the present invention, but the protection scope of the present invention is not limited thereto. Any technician familiar with the technical field can easily think of various equivalent modifications or substitutions within the technical scope disclosed in the present invention, and these modifications or substitutions should be included in the protection scope of the present invention.

Claims

1. A collaborative guidance method for a low-cost strapdown seeker, characterized in that: include: The line of sight angle in the inertial system is obtained based on the stereoscopic line of sight angle in the missile system measured by the strapdown seeker; Based on the sight angle in the inertial system, the sight angle and sight angle rate are estimated using the α-β filtering algorithm; Considering the terminal landing angle constraint, a space cooperative guidance law is established based on the estimated line-of-sight angular velocity. Considering the time constraint, the temporal cooperative guidance is established based on the estimated line-of-sight angular rate; According to different scenarios, the space cooperative guidance law or the time cooperative guidance is selected to respectively perform space cooperative guidance or time cooperative guidance on the strapdown aircraft.

2. The collaborative guidance method for low-cost strapdown seeker according to claim 1, characterized in that: The stereoscopic line of sight angle in the missile system measured by the strapdown seeker is used to obtain the line of sight angle in the inertial system, specifically: Based on the volumetric line of sight angle measured by the strapdown seeker in the missile system, the two are decoupled by line of sight angle reconstruction, and the relationship between the line of sight azimuth and elevation angle in the inertial system is obtained: Where R 11 , R 12 , R 13 , R 21 , R 22 , R 23 , R 31 , R 32 , R 33 is the coordinate system transformation coefficient, q a ,q b is the sight elevation angle and sight azimuth angle under the missile system, q λ ,q γ are the line of sight azimuth and line of sight altitude angle in the inertial system.

3. The collaborative guidance method for low-cost strapdown seeker according to claim 1, characterized in that: The sight angle based on the inertial system uses a filtering algorithm to estimate the sight angle and the sight angle rate, and the formula used is: Among them, α and β are filter gains, T is the filtering algorithm time interval, is the estimated value of the system state, is the velocity estimate, Is the system at t k+1 The state value at the moment, Yes k+1 The speed of the system at a given moment.

4. The collaborative guidance method for low-cost strapdown seeker according to claim 1, characterized in that: Considering the terminal landing angle constraint, the space cooperative guidance law is established based on the estimated line of sight angular rate, specifically: Where N is the proportional coefficient, v is the aircraft speed, is the line of sight angular velocity, a B is the drop angle constraint offset term.

5. The collaborative guidance method for low-cost strapdown seeker according to claim 4, characterized in that: The angle constraint bias term a B Calculated according to the following formula: from B =-(σ-Nq+(N-1)σ d )v / t go Among them, σ is the ballistic angle of the aircraft, σ d It is the expected falling angle.

6. The collaborative guidance method for low-cost strapdown seeker according to claim 1, characterized in that: Considering the time constraint, the time coordinated guidance is established based on the estimated line of sight angular rate, specifically: Among them, t go is the remaining attack time, R go is the remaining flight distance, ε t is the attack time error.

7. A computer system, characterized in that include: One or more processors, and a computer-readable storage medium for storing one or more programs, wherein when the one or more programs are executed by the one or more processors, the one or more processors are enabled to implement the method of claim 1.

8. A computer-readable storage medium, characterized in that Computer executable instructions are stored, and when the instructions are executed, they are used to implement the method of claim 1.

9. A computer program product, characterized in that The invention comprises computer executable instructions, which are used to implement the method of claim 1 when being executed.

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