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Showing papers on "Missile published in 2014"



Journal ArticleDOI
TL;DR: In this article, an accurate time-to-go estimation of PN guidance law for a stationary target is proposed by using a simple interpolation of the estimated time-time of the PN trajectory.
Abstract: P ROPORTIONAL navigation (PN) is a popular guidance law because of its robustness, simplicity, and ease of implementation PN guidance law is capable of intercepting both nonmaneuvering and maneuvering targets [1] PN is also an optimal control effort guidance law in the linearized setting [2] Modern tactical guidance laws are designed not only to intercept a target but to satisfy other constraints [3] like impact time, impact angle, and other optimality constraints These guidance laws can often be represented as being comprising of a PN component and other components designed to satisfy other constraints As these guidance laws require an estimate of the time-to-go, it is important to develop methods to obtain an accurate estimate of the time-to-go for the PNguidance law An interesting application requiring the estimation of time-to-go is in salvo attack problems [4], in which several missiles, launched at different times, need to achieve interception of the target simultaneously Using PN guidance law to achieve this task requires an accurate estimation of time-to-go Impact time, impact angle, and other optimal guidance laws also need information of the time-to-go to compute the guidance command One of the simplest ways to estimate time-to-go is by computing range over closing velocity (or the negative of the range rate) The range over range rate estimate is accurate only when the closing velocity is constant and the initial heading error is small Improved methods for time-to-go estimation by solving a linear quadratic control problem is discussed in York and Pastrick [5] In Tahk et al [6], a recursive time-to-go method is discussed in which the timeto-go estimates are updated recursively The time-to-go estimate is updated in every guidance cycle time and a compensationmethod for the time-to-go error is obtained The time-to-go estimate improves as the engagement proceeds because the guidance law used reduces the heading error In Sang et al [7] the commanded guidance history is considered to obtain the time-to-go estimate The time-to-go estimate in Jeon et al [4] is obtained as a function of the initial range and heading error Our time-to-go estimation method also depends on the initial conditions and hence, our results are compared with Jeon et al [4] In this Note, an accurate method to obtain a time-to-go estimate of PN guidance law for a stationary target is proposed The time-to-go for a stationary target is obtained accurately by a simple interpolation It is shown that from the time-to-go estimate of a particular initial condition, the time-to-go estimate of any other initial condition can be obtained by a simple but elegant time scaling property of the PN guidance law The time-to-go estimate as a function of the heading error for the particular initial condition is used as the base solution for estimating the time-to-go for other initial conditions This method also gives estimates for initial conditions with large initial heading errors; even for cases when the missile is launched with a high heading error Themethod can be used to obtain time-to-go estimates for any initial condition and navigation gain The time-to-go estimation is extended to constant velocity targets by a simple iterative process The time-to-go obtained after the initial iteration is exact and need not be computed in every guidance cycle time

86 citations


Journal ArticleDOI
TL;DR: In this paper, a generalized model predictive static programming technique is presented for rapidly solving a class of finite-horizon nonlinear optimal control problems with hard terminal constraints, such as the double integrator problem.
Abstract: A new generalized model predictive static programming technique is presented for rapidly solving a class of finite-horizon nonlinear optimal control problems with hard terminal constraints. Two key features for its high computational efficiency include one-time backward integration of a small-dimensional weighting matrix dynamics, followed bya static optimization formulation that requires only a static Lagrange multiplier to update the control history. It turns out that under Euler integration and rectangular approximation of finite integrals it is equivalent to the existing model predictive static programming technique. In addition to the benchmark double integrator problem, usefulness of the proposed technique is demonstrated by solving a three-dimensional angle-constrained guidance problem for an air-to-ground missile, which demands that the missile must meet constraints on both azimuth and elevation angles at the impact point in addition to achieving near-zero miss distance, while minimizing the lateral acceleration demand throughout its flight path. Simulation studies include maneuvering ground targets along with a first-order autopilot lag. Comparison studies with classical augmented proportional navigation guidance and modern general explicit guidance lead to the conclusion that the proposed guidance is superior to both and has a larger capture region as well.

81 citations


Journal ArticleDOI
TL;DR: In this article, an integrated missile guidance and control law based on adaptive fuzzy sliding mode control is presented, where an adaptive fuzzy system is adopted to approximate the coupling nonlinear functions of the system, and for the uncertainties, an online-adaptive control law is used to estimate the unknown parameters.

60 citations


Proceedings ArticleDOI
01 Dec 2014
TL;DR: This paper addresses a three-body pursuit-evasion scenario where an Attacker missile using Pure Pursuit guidance pursues a Target aircraft and a Defender missile launched by a wingman aims at intercepting the Attacker before it reaches the aircraft.
Abstract: This paper addresses a three-body pursuit-evasion scenario where an Attacker missile using Pure Pursuit guidance pursues a Target aircraft and a Defender missile launched by a wingman aims at intercepting the Attacker before it reaches the aircraft. An optimal control problem is posed which captures the goal of the Target-Defender team, namely, to maximize the separation between Target and Attacker at the instant of capture of the Attacker by the Defender. The optimal control law provides the heading angles for the Target and the Defender team.

55 citations


Proceedings ArticleDOI
01 Sep 2014
TL;DR: The solution to this pursuit-evasion differential game provides optimal heading angles for the Target and the Defender team to maximize the terminal separation between Target and Attacker and it also provides the optimal heading angle for the Attacker to minimize the said distance.
Abstract: A pursuit-evasion differential game involving three agents is discussed. This scenario considers an Attacker missile pursuing a Target aircraft. The Target is however aided by a Defender missile launched by, say, its wingman, to intercept the Attacker before it reaches the Target aircraft. Thus, a team is formed by the Target and the Defender which cooperate to maximize the distance between the Target aircraft and the point where the Attacker missile is intercepted by the Defender missile, and the Attacker which tries to minimize said distance. The solution to this differential game provides optimal heading angles for the Target and the Defender team to maximize the terminal separation between Target and Attacker and it also provides the optimal heading angle for the Attacker to minimize the said distance.

54 citations


Journal ArticleDOI
TL;DR: In this paper, a linear quadratic optimal framework is proposed to solve the impact angle control problem with respect to a collision triangle, where the collision triangle is defined to be the triangle formed by the initial positions of target and missile, and the intercept point at which the missile hits the target when flown by a straight (with zero effort) line.
Abstract: C ONTROL of missile-target-relative geometry is one of the desired features of guidance in many modern applications. A typical example is to impact a ground target in a direction perpendicular to the tangent plane of the terrain with very high precision both in miss distance and impact angle [1]. Various needs for the maximum warhead effectiveness, and sometimes enhancement of survivability of the missile launch vehicle, in naval applications call for guidance laws that can achieve a specified final direction of approach to the target as well [2]. Also, ensuring a small angle of the missile body relative to the target during the whole engagement process is critical in the case of missiles with strapdown seekers [3]. This necessity of control of terminal engagement geometry has been amajor thrust for much of the researchwork in the area of guidance law design with impact angle constraints. In this Note, a novel method of optimal impact angle control guidance law development based on linear quadratic optimal framework [4] against an arbitrary maneuvering target is presented. Throughout the Note, the missile velocity profile is assumed to be arbitrary. The equation of motion of a missile is often written in terms of the angular variables associated with velocity vectors; in this case, the missile acceleration is computed as the angular rate of its velocity vector multiplied by the magnitude of the velocity, which is directly realizable for aerodynamically controlled missiles. The main problem here is the fact that the kinematics is now nonlinear, defying closed-form solutions of many optimal guidance problems of interest. Thus, it has been common practice to linearize the kinematics (e.g., [5]) or approximate by linear equations [6] to come up with a nice linear quadratic optimal guidance problem. A natural question to follow is then how we linearize the kinematics in a right manner. This question, however, has not been addressed adequately in the literature and linearization has been performed in many cases with the usual assumption of small values of angular variables involved. As a result, the guidance laws often yield poor performance when the associated angular variables get larger. Usually, the collision triangle is defined to be the triangle formed by the initial positions of target and missile, and the intercept point at which the missile hits the target when flown by a straight (with zero effort) line. When no specific requirement on final engagement geometry is posed, the linearization about the usual collision triangle works reasonably well (e.g., [7]; also see [8]). If a specific impact angle between the missile and target velocity vectors is required, however, the usual collision triangle no longer serves as zero-effort collision geometry because the missile trajectory may largely deviate from the collision triangle to satisfy the specific impact angle requirement. This is why some papers just assume before linearization that the end game is initiated with a collision triangle satisfying closely the impact angle requirement [9]. No attempt, however, has been made yet to address specific questions such as what collision trianglewe should be looking for and howwe compute and use it for the linear optimal guidance problem formulation. In this Note, we introduce and use, as the basis of linearization, the perfect (or zero effort) collision triangle for the impact angle control problem, which varies depending on the value of the prescribed impact angle, and solve a linear optimal guidance problem.

54 citations


Journal ArticleDOI
01 Feb 2014-Optik
TL;DR: Consensus of networks with time-delays under switching topologies is proved using common Lyapunov function method using feedback linearization to attain linear guidance law for each missile, which is the base law for cooperative.

53 citations


Journal ArticleDOI
TL;DR: In this paper, an approach of integrated guidance and control (IGC) design for interception of maneuvering targets (evaders) is presented, where an IGC model with uncertainties in the pitch plane is formulated, and, by adopting a backstepping scheme, an adaptive nonlinear IGC approach is developed.
Abstract: This paper presents an approach of integrated guidance and control (IGC) design for interception of maneuvering targets (evaders). An IGC model with uncertainties in the pitch plane is formulated, and, by adopting a backstepping scheme, an adaptive nonlinear IGC approach is developed. Theoretical analysis shows that the design approach makes the line-of-sight (LOS) rate be input-to-state stable (ISS) with respect to target maneuvers and missile model uncertainties, and the stability of the missile dynamics can be guaranteed as well. The numerical simulation confirms the effectiveness of the proposed design approach.

50 citations


Journal ArticleDOI
TL;DR: An omega-K algorithm for missile-borne synthetic aperture radar (SAR) diving with constant acceleration to derive a reference function multiplication from the equivalent range history with a form similar to that of conventional SAR.
Abstract: This letter proposes an omega-K algorithm for missile-borne synthetic aperture radar (SAR) diving with constant acceleration. The essence of the novel method is to derive a reference function multiplication from the equivalent range history. The true range history is divided into two parts: range independent and range dependent. After compensating the range-independent part with considering all the terms, we can regard the range-dependent part as an equivalent range history with a form similar to that of conventional SAR. Then, an analytical 2-D spectrum can be used for the omega-K algorithm. The proposed algorithm can handle general missile configurations with wide swaths and high squint angles. Simulation results are presented to validate the proposed method.

45 citations


Journal ArticleDOI
TL;DR: In this article, a robust nonlinear controller for a highly maneuverable air-to-air (ATA) missile is presented, where the reference signals in angle of attack, sideslip, and bank angle produced by the external guidance system are followed by robust sigmoid-like control functions.
Abstract: This paper deals with the design of a robust nonlinear controller for a highly maneuverable missile. Stabilization and tracking are achieved, exploiting a detailed nonlinear model of the six-degree-of-freedom, nonminimum phase, uncertain, and time-varying dynamics of a non-axial-symmetric air-to-air tail-controlled missile. A robust backstepping approach is applied to the multi-input/multi-output model to achieve both bank-to-turn and skid-to-turn maneuvers. Control objectives consist of following the reference signals in angle of attack, sideslip, and bank angle produced by the external guidance system, in order to pursue highly agile maneuvers. Uncertain terms, mostly due to aerodynamic coefficients and dynamic pressure, are suitably limited by bounding functions constructed using experience, a priori knowledge on system behavior, and a bit of conservatism. Robust sigmoidlike control functions are then used to dominate in size the uncertain terms. The whole control system is shown to be practically-robu...

Journal ArticleDOI
TL;DR: In this paper, an interactive multiple-model-based guidance law estimation filter was proposed to estimate an antiship missile's threat index, such as the remaining time of flight and impact angle.
Abstract: This paper presents an interactive multiple-model-based guidance law estimation filter. An antiship missile’s threat index, such as the remaining time of flight and impact angle, can be estimated by using the guidance law estimation filter. The filter bank of the filter was composed of models based on proportional navigation and impact angle control guidance laws. In the guidance law estimation filter, state variables, rather than fixed values of navigation constant or impact angle, are used for estimation. To that end, the modified interactive multiple model, which is a modified version of the conventional interactive multiple model method is presented in this paper. The modified interactive multiple model method minimizes the degradation of estimation performance by systematic approach for different state vectors. The effectiveness of the modified interactive multiple model method and the performance of the guidance law estimation filter are examined through the numerical simulation of various engagemen...

Journal ArticleDOI
TL;DR: In this paper, the authors presented a modified optical scanning seeker with programmed and tracking controls with the simultaneous influence of external interferences from the direction of a self-guiding rocket.

Journal ArticleDOI
01 Jun 2014
TL;DR: In this article, a global sliding mode controller (GSMC) is proposed for the missile electromechanical actuator (EA) servo system, where there exists high uncertainties, such as parameter variations and external disturba...
Abstract: A global sliding mode controller (GSMC) is proposed for the missile electromechanical actuator (EA) servo system, where exists high uncertainties, such as parameter variations and external disturba...

Journal ArticleDOI
TL;DR: In this article, the authors proposed a finite time convergent guidance law for a homing missile to intercept a maneuvering target. And they proved that the line-of-sight (LOS) angular rate converges to zero in finite time under the proposed guidance law.
Abstract: Taking into consideration both the autopilot dynamics and uncertainties, this paper proposes a finite time convergent guidance law for homing missile to intercept a maneuvering target Firstly, an exact observer (differentiator) is employed to estimate the target maneuvers in finite time Then, a finite time convergent guidance law is designed based on the existing finite time sliding-mode control theory It is proved that the line-of-sight (LOS) angular rate converges to zero in finite time under the proposed guidance law Compared with the existing finite time guidance laws, this guidance law can compensate for the effects of the autopilot dynamics and uncertainties, and the information used for feedback control is much easier to obtain Finally, simulation results show that our scheme, as a finite time convergent algorithm, has strong robustness to bounded disturbances

01 Apr 2014
TL;DR: In this paper, the authors present a survey of the People s Republic of China (PRC) ASCM and LACM programs and their implications for broader PLA capabilities, especially in a Taiwan scenario.
Abstract: : China s military modernization is focused on building modern ground, naval, air, and missile forces capable of fighting and winning local wars under informationized conditions. The principal planning scenario has been a military campaign against Taiwan, which would require the People s Liberation Army (PLA) to deter or defeat U.S. intervention. The PLA has sought to acquire asymmetric assassin s mace technologies and systems to overcome a superior adversary and couple them to the command, control, communications, computers, intelligence, surveillance, and reconnaissance (C4ISR) systems necessary for swift and precise execution of short-duration, high-intensity wars. A key element of the PLA s investment in antiaccess/area-denial (A2/AD) capabilities is the development and deployment of large numbers of highly accurate antiship cruise missiles (ASCMs) and land-attack cruise missiles (LACMs) on a range of ground, air, and naval platforms. China s growing arsenal of cruise missiles and the delivery platforms and C4ISR systems necessary to employ them pose new defense and nonproliferation challenges for the United States and its regional partners. This study surveys People s Republic of China (PRC) ASCM and LACM programs and their implications for broader PLA capabilities, especially in a Taiwan scenario. Key findings are presented below.

Journal ArticleDOI
TL;DR: While the numerical simulation is able to capture the experimental trend and results, a comparison of penetration depth and scabbing and perforation limits as per different empirical formulation shows substantial divergence.

Journal ArticleDOI
TL;DR: In this paper, the authors proposed a guidance law for guiding a missile against a maneuvering target while satisfying a circular no-fly zone (NFZ) constraint, which consists of two parts: virtual-target guidance (VTG) and boundary-constraint handling scheme (BCHS).
Abstract: The proposed guidance law is used for guiding a missile against a maneuvering target while satisfying a circular no-fly zone (NFZ) constraint. It consists of two parts: virtual-target guidance (VTG) and boundary-constraint handling scheme (BCHS). In order that the missile avoids the NFZ, VTG first maps the actual target to a virtual one, then obtains the relative motion between the virtual target and missile, and finally uses proportional navigation to steer the missile to the virtual target. The missile also hits the actual target when it hits the virtual target because the virtual and actual targets are coincident at this moment. In some cases, especially when the initial velocity vector of the missile points toward the center of the NFZ, if the evasive action taken by VTG is found to be insufficient, then BCHS will be enabled to keep the missile from entering the NFZ unless the target enters the NFZ.

Journal ArticleDOI
TL;DR: A new homing guidance law is introduced based on the energy cost weighted by a Gaussian function in order to shape the missile's trajectory and distribute the acceleration command during the engagement.
Abstract: In this paper, a new homing guidance law is introduced based on the energy cost weighted by a Gaussian function in order to shape the missile's trajectory and distribute the acceleration command during the engagement. This law can improve guidance performance by alleviating sensitivity with respect to initial heading error and by reducing the possibility of command saturation through distributing acceleration demand properly.

Journal ArticleDOI
TL;DR: The proposed autopilots have a simple structure and require no time-consuming gain scheduling for many flight conditions, while providing satisfactory tracking and robustness over the entire flight envelope.
Abstract: This paper considers two autopilot designs using H ∞ loop shaping for an agile missile that experiences high angle of attack, highly nonlinear and rapidly changing dynamics, and aerodynamic variation after launch. The main autopilot design is started with two H ∞ control designs intended to cover the low-speed and high-speed regions of the flight envelope. Then, the two control designs are combined (via Mach variation) to construct a global controller that covers the entire flight envelope. The proposed autopilots have a simple structure and require no time-consuming gain scheduling for many flight conditions, while providing satisfactory tracking and robustness over the entire flight envelope. The performance of the designed autopilots is checked via a comparison study and a challenging intercept scenario. These performance test results clearly demonstrate the merit of the proposed designs.

Journal ArticleDOI
TL;DR: In this article, an extended trajectory shaping guidance (ETSG) law is proposed under the assumption that the missile-target relative velocity is constant and the line of sight angle is small.

Journal ArticleDOI
TL;DR: An extended nonlinear chirp scaling algorithm for focusing missile borne synthetic aperture radar data is proposed, compensating the azimuth dependent characteristic of the Azimuth FM rate and adopting higher order approximation processing, which means easier implementation and higher efficiency.

Journal ArticleDOI
TL;DR: In this article, a Lyapunov-based pursuit guidance law against stationary targets is proposed to reduce the angle between the velocity vector and the distance vector between the missile and the target.

Proceedings ArticleDOI
01 Aug 2014
TL;DR: In this paper, a fault-tolerant integrated guidance and control law is proposed based on backstepping and input-to-state stability in the presence of actuator failures, target maneuvers and unknown bounded model uncertainties.
Abstract: A fault-tolerant integrated guidance and control law is proposed based on backstepping and input-to-state stability in the presence of actuator failures, target maneuvers and unknown bounded model uncertainties. An adaptive law is designed to estimate the unknown parameter with respect to the actuation effectiveness components. Theoretical analysis is shown that the line-of-sight (LOS) rate is input-to-state stable (ISS), and the stability of the missile dynamics can be guaranteed as well. The numerical simulation results confirm the effectiveness of the fault-tolerant IGC law.

Journal ArticleDOI
TL;DR: In this paper, a longitudinal autopilot for angle-of-attack tracking based on backstepping control is proposed to handle the nonlinear, rapidly changing dynamics and aerodynamic uncertainties.
Abstract: This paper deals with a nonlinear adaptive autopilot design for agile missile systems. In advance of the autopilot design, an investigation of the agile turn maneuver, based on the trajectory optimization, is performed to determine state behaviors during the agile turn phase. This investigation shows that there exist highly nonlinear, rapidly changing dynamics and aerodynamic uncertainties. To handle of these difficulties, we propose a longitudinal autopilot for angle-of-attack tracking based on backstepping control methodology in conjunction with the time-delay adaptation scheme.

Journal ArticleDOI
01 Mar 2014
TL;DR: The proposed DE enhanced guidance law is compared against the existing conventional laws in the literature, on the criteria of time and energy optimality, peak lateral acceleration demanded, terminal speed and robustness to unanticipated target maneuvers, to illustrate the superiority of the proposed law.
Abstract: This paper presents a novel, soft computing based solution to a complex optimal control or dynamic optimization problem that requires the solution to be available in real-time. The complexities in this problem of optimal guidance of interceptors launched with high initial heading errors include the more involved physics of a three dimensional missile-target engagement, and those posed by the assumption of a realistic dynamic model such as time-varying missile speed, thrust, drag and mass, besides gravity, and upper bound on the lateral acceleration. The classic, pure proportional navigation law is augmented with a polynomial function of the heading error, and the values of the coefficients of the polynomial are determined using differential evolution (DE). The performance of the proposed DE enhanced guidance law is compared against the existing conventional laws in the literature, on the criteria of time and energy optimality, peak lateral acceleration demanded, terminal speed and robustness to unanticipated target maneuvers, to illustrate the superiority of the proposed law.

Proceedings ArticleDOI
01 Aug 2014
TL;DR: Simulation results show the proposed architecture can realize coordinated attack effectively and has the characteristics of brief structure, short coordinate time and little information to hand on.
Abstract: According to the characteristics of multiple missiles cooperative guidance, a time-cooperative guidance architecture based on leader-follower strategy is proposed This architecture is composed of individual guidance for each missile and coordinating strategy of the whole system The central process unit situated at the leader coordinates the system, and the cooperative information is broadcasted from leader to followers The coordinate information includes expected impact time and relative motion information of leader and target The time-cooperative guidance architecture has the characteristics of brief structure, short coordinate time and little information to hand on For the features of structure and the follower have no seeker, a specific and feasible algorithm of line of sight rate is presented, the requirement of follower computation and detection equipment is decreased Simulation results show the proposed architecture can realize coordinated attack effectively

Proceedings ArticleDOI
16 Jun 2014
TL;DR: The Preliminary Aerothermal Structural Simulation (PASS) is a tool suite for MultiDisciplinary Analysis and Optimization of high speed vehicles at the conceptual or preliminary design level that includes geometry definition, preliminary packaging, aerodynamic estimation, trajectory modeling, structural sizing, and thermal analysis.
Abstract: The Preliminary Aerothermal Structural Simulation (PASS) is a tool suite for MultiDisciplinary Analysis and Optimization of high speed vehicles at the conceptual or preliminary design level. The PASS process includes geometry definition, preliminary packaging, aerodynamic estimation, trajectory modeling, structural sizing, and thermal analysis. PASS currently includes Missile Datcom (MDC) to provide aerodynamic data, Mass Estimator for Monocoque Aero Structures (MEMAS) to estimate the structural mass, Aerothermal Targets Analysis Program (ATAP) to provide thermal analysis, and MATLAB to perform the integration. MDC and ATAP are standard AFRL tools. MEMAS is a new in-house code developed in MATLAB for sizing high speed attack missile structures, including articulated fins. The Packaging algorithm is a new in-house code developed for placement of representative sub-systems and calculation of mass properties/moments of inertia.

Proceedings ArticleDOI
20 May 2014
TL;DR: In this article, the authors presented the model of two-axis gimbal seeker and improved its performance using proposed fuzzy controller The equations of gimbals motion are derived using Lagrange equation and the stabilization system is constructed considering the missile rates, torque disturbances, and cross coupling between elevation and azimuth channels The overall control system is simulated using MATLAB/Simulink This model is evaluated comparing with conventional PI control system
Abstract: The application of the guided missile seeker is to provide stability to the sensor by isolating it from the missile motion and vibration The aim of this paper is to present the model of two axes gimbal seeker and improve its performance using proposed fuzzy controller The equations of gimbals motion are derived using Lagrange equation and the stabilization system is constructed considering the missile rates, torque disturbances, and cross coupling between elevation and azimuth channels The overall control system is simulated using MATLAB/Simulink This model is evaluated comparing with conventional PI control system The comparative simulation results in different conditions have shown that the proposed fuzzy PID method offers a better performance than the classical one

Patent
30 Jul 2014
TL;DR: In this paper, a system for evaluating the disturbance rejection rate parasitical loop of a strap-down infrared seeker is presented, which consists of a target simulator, a three-axis rotary table, the seeker and an emulation computer.
Abstract: The invention discloses a system for evaluating the disturbance rejection rate parasitical loop of a strap down infrared seeker The system comprises a target simulator, a three-axis rotary table, the seeker and an emulation computer, wherein the target simulator simulates the movement relation of a target; the strap down infrared seeker detects the infrared signal source of the target simulator, and outputs a missile-target LOS (Line of Sight) rate obtained after decoupling filtering to the emulation computer, the emulation computer calculates the attitude and position information of a missile according to a practical kinetic model of the missile and transmits the attitude angle of a missile body to the three-axis rotary table to simulate the movement of the missile body; the emulation computer figures out a theoretical missile-target LOS rate and transmits the theoretical missile-target LOS rate to the target simulator to simulate the movement of the missile-target to form an entire closed-loop testing system In the whole process, the emulation computer stores test data for final disturbance rejection rate calculation The system eliminates the problem of singular point in the traditional disturbance rejection rate testing method, and enables that the disturbance rejection rate test is more accurate and a testing system is closer to the practical combat mode of the missile