Showing papers on "Missile published in 2010"

Homing Guidance Law for Cooperative Attack of Multiple Missiles

Abstract: OVER the past few years, there have been significant efforts devoted to the research and development of cooperative unmannedsystems [1–3].The formationflyingofmultipleunmanned aerial vehicles (UAVs) has been studied for radar deception, reconnaissance, surveillance, and surface-to-air-missile jamming in military operations. An example of a cooperative operational scenario of multiple vehicles is that of a small UAV flying over an urban area, dispensingmultiplemicro aerial vehicles to examinepointsof interest fromclosedistances [4].Agroupofwell-organized low-costmultiple vehicles can be far superior to a single high-technology and high-cost UAV in effectiveness. Tactical missile systems as well as UAVs provide more capabilities when they are organized as a coordinated group than when they are operated independently. Modern antiship missiles need to be able to penetrate the formidable defensive systems of battleships such as antiair defense missile systems and close-in weapon system (CIWS). CIWS is a naval shipboard weapon system for detecting and destroying incoming antiship missiles and enemy aircraft at short range. These defensive weapons with powerful fire capability and various strategies seriously intimidate the survivability of the conventional antiship missiles. Hence, antiship missile developers have made great efforts to develop a high-performance missile system with ultimate sea-skimming flight and terminal evasive maneuvering capabilities despite a huge cost. On the other hand, cooperative attack strategies have been studied to enhance survivability of the conventional ones. Here, a cooperative attack means that multiple missiles attack a single target or multiple targets cooperatively or, in a specific case, simultaneously [5,6]. Clearly, it is difficult to defend a group of attackers bursting into sight at the same time, even though each member is the conventional one in performance. So the simultaneous attack ofmultiple missiles is a cost-effective and efficient cooperative attack strategy. A simultaneous attack of a group of missiles against a single common target can be achieved by two ways. The first approach is individual homing, inwhich a common impact time is commanded to all members in advance, and thereafter each missile tries to home on the target on time independently. The second is cooperative homing, inwhich themissiles communicate among themselves to synchronize the arrival times. In other words, the missiles with larger times-to-go try to take shortcuts, whereas others with shorter times-to-go take detours to delay the arrival times. The first concept requires determination of a suitable common impact time before homing, but the second needs online data links throughout the engagement. Despite a number of studies on guidance problems related to timeto-go [7–10], studies on guidance laws to control impact time for a simultaneous attack are rare, except a few recent works by the authors. An impact-time-control guidance law (ITCG) for antiship missiles was developed in [5] and, as an extension of this study, a guidance law to control both impact time and angle (ITACG) was presented in [11]. These individual homing methods are based on optimal control theory, providing analytical closed-loop guidance laws. Herein, the desired impact time is assumed to be prescribed before the homing phase starts. Alternatively, this Note is concerned with a new guidance law based on the second approach, cooperative homing, for a simultaneous attack of multiple missiles. Proportional navigation (PN) is a well-known homing guidance method in which the rate of turn of the interceptor is made proportional with a navigation ratio N to the rate of turn of the line of sight (LOS) between the interceptor and the target. The navigation constant N is a unitless gain chosen in the range from 3 to 5 [12]. Although PNwithN 3 is known to be energy-optimal, an arbitrary N > 3 is also optimal if a time-varying weighting function is included into the cost function of the linear quadratic energy-optimal problem [13,14]. In general, the navigation ratio is held fixed. In some cases, however, it can be considered as a control parameter to achieve a desired terminal heading angle [15].Although PN results in successful intercepts under a wide range of engagement conditions, its control-efficiency is not optimal, in general, especially for the case of maneuvering targets [16]. Augmented proportional navigation, a variant of PN, is useful in cases in which target maneuvers are significant [12]. Biased proportional navigation is also commonly used to compensate for target accelerations and sensor noises or to achieve a desired attitude angle at impact [17]. Even if PN and its variants are alreadywell known andwidely used, they are not directly applicable to many-to-one engagements. This Note proposes a homing guidance law called cooperative proportional navigation (CPN) for many-to-one engagements: CPN has the same structure as conventional PN except that it has a time-varying navigation gain that is adjusted based on the onboard time-to-go and the times-to-go of the other missiles. CPN uses the time-varying navigation gain as a control parameter for reducing the variance of times-on-target of multiple missiles. This Note begins with the formulation of the homing problem of multiple missiles against a single target, subject to constraints on the impact time. Next, preliminary concepts such as the relative time-togo error and the variance of times-to-go of multiple missiles are introduced and a new guidance law is proposed. Then the major property of the law is investigated and the characteristics of the law for the case of twomissiles are examined in detail. Finally, numerical simulation results illustrate the performances of the proposed law.

Cooperative Multiple-Model Adaptive Guidance for an Aircraft Defending Missile

Abstract: A cooperative guidance law, for a defender missile protecting an aerial target from an incoming homing missile, is presented. The filter used is a nonlinear adaptation of a multiple model adaptive estimator, in which each model represents a possible guidance law and guidance parameters of the incoming homing missile. Fusion of measurements from both the defender missile and protected aircraft is performed. A matched defender’s missile guidance law is optimized to the identified homing missile guidance law. It utilizes cooperation between the aerial target and the defender missile. The cooperation stems from the fact that the defender knows the future evasive maneuvers to be performed by the protected target and thus can anticipate the maneuvers it will induce on the incoming homing missile. Moreover, the target performs a maneuver that minimizes the control effort requirements from the defender. The estimator and guidance law are combined in a multiple model adaptive control configuration. Simulation results show that combining the estimations with the proposed optimal guidance law, that utilizes cooperation between the defending missile and protected target, yields hit-to-kill closed loop performance with very low control effort. This facilitates the use of relatively small defending missiles to protect aircrafts from homing missiles.

Synthetic-Waypoint Guidance Algorithm for Following a Desired Flight Trajectory

Abstract: T HE development of aircraft guidance, navigation, and control systems has been a long-standing research area. Numerous methods relating to the enhancement of aircraft performance under various mission parameters have been developed in response to a need for more reliable and robust guidance systems. Current guidance systems applied to commercial, civilian, and unmanned aircraft rely on the knowledge of a flight path, specified bywaypoints located in inertial space. Most missions are considered successful when the vehicle reaches the designatedwaypoint at which new commands are issued to the vehicle to proceed to the next waypoint. Two common types of conventional aircraft guidance are the direct-to-waypoint (DTW) and track-to-waypoint (TTW) methods in relation to pathfollowing between designated waypoints. The DTWmethod simply issues heading commands to the vehicle based on the angular difference between thewaypoint and vehicle.When thevehicle reaches the waypoint, the control system issues a new command to guide the aircraft to the next waypoint. The TTW method aims to follow the track betweenwaypoints. In this guidancemethod the control system aims to minimize the lateral offset between the prescribed flight path and the aircraft’s position, issuing heading commands that return the vehicle to the nominal flight path. The track method therefore places the additional constraint on a flight path that the vehicle must follow in order to reach the waypoint, rather than simply reaching the waypoint. However, both methods are far from optimal. This is evident in how the aircraft transitions between flight paths after reaching awaypoint. During theseflight-path transitions, the aircraft will often overshoot the desiredflight path to correct its track, particularlywhen the flight-path transition angle is acute. Various control strategies have been investigated to alleviate or minimize flight-path deviations. Such strategies include applyingmodern control methods such as receding-horizon control [1,2] and model predictive control [3] to anticipate flight-path changes and take control action before reaching a goal while maintaining adequate vehicle flight performance. Missile guidance and control systems operate on similar principles to commercial, civilian, and unmanned aircraft guidance and control algorithms. The primary mission for missile systems is to intercept a moving target using information about the relative position and velocity between the pursuer and target. One of thefirstmethods used in missile guidance was pursuit guidance (PG) [4–7]. The method operates by forcing the angular displacement error between a pursuer and its target to zero. Control commands scaled by a proportional factor of the current error are then issued to direct the pursuer along the line of sight (LOS) between the pursuer and target. PG solutions, however, do not consider the path taken or the levels of system performance required by the pursuer in reaching the target, resulting in a far-from-optimal solution. To address this problem of suboptimality, additional parameters have been introduced to enhance missile performance. One method includes taking into account the motion of the commanded line of sight between the pursuer and target [4,7–10], issuing lateral acceleration commands based on tracking error and tracking error rate to the target. This approach has been shown to improve overall interceptor performance compared with conventional PG [4,7]. Another suchmethod aims tomodify the level of control the guidance algorithm possesses over the vehicle by adjusting the level of proportional gain. This is achieved by gain scheduling [11] to select gain values based on current interceptor states. In addition to modifying internal missile guidance and control parameters such as variable gains and LOS rate estimation, mission performance can be enhanced by manipulating the trajectory taken by the pursuer to the targets. A good example of such a method is discussed in [12], in which a missile aims to exploit the aerodynamic benefits of high-altitude flight by tracking a virtual target at some initially high altitude that is not necessarily along the trajectory to the true target. This Note discusses the development of a guidance law fusing the virtual-target concepts with those of pursuit guidance for implementation into an aircraft guidance system. This Note develops a path-following aircraft guidance algorithm that pursues synthetic waypoints using only a small set of guidance parameters, extending the virtual-target concept to complete aircraft guidance. The path is defined by the track between a minimal set of waypoints at specified locations, removing the need for a smooth path to be defined or the need for complicated path-switching logic or trajectory planning when awaypoint is reached. The synthetic waypoint travels along the path between waypoints, with the trailing aircraft traveling a smooth path generated through its own dynamics in following the synthetic waypoint. The guidance law is tested by varying guidance parameters, thus assessing vehicle sensitivity to and overall system performance of parameter variations. The following sections discuss the basic concepts of missile and aircraft guidance and provide a detailed description of the structure of the synthetic-waypoint guidance algorithm. A discussion on the implementation of the algorithm into the underlying aircraft control system will also be presented, followed by an analysis of the performance of the guidance algorithm in nonlinear simulation. Received 30 June 2009; revision received 3 November 2009; accepted for publication 9November 2009.Copyright©2009 by theAmerican Institute of Aeronautics and Astronautics, Inc. All rights reserved. Copies of this paper may be made for personal or internal use, on condition that the copier pay the $10.00 per-copy fee to the Copyright Clearance Center, Inc., 222 Rosewood Drive, Danvers, MA 01923; include the code 0731-5090/10 and $10.00 in correspondence with the CCC. ∗Graduate Research Student, School of Aerospace, Mechanical and Mechatronic Engineering; e.medagoda@aeromech.usyd.edu.au. Senior Lecturer, School of Aerospace, Mechanical and Mechatronic Engineering; pwg@aeromech.usyd.edu.au. JOURNAL OF GUIDANCE, CONTROL, AND DYNAMICS Vol. 33, No. 2, March–April 2010

Modern Homing Missile Guidance Theory and Techniques

Abstract: Classically derived homing guidance laws, such as proportional navigation, can be highly effective when the homing missile has significantly more maneuver capability than the threat. As threats become more capable, however, higher performance is required from the missile guidance law to achieve intercept. To address this challenge, most modern guidance laws are derived using linear-quadratic optimal control theory to obtain analytic feedback solutions. Generally, optimal control strategies use a cost function to explicitly optimize the missile performance criteria. In addition, it is typical for these guidance laws to employ more sophisticated models of the target and missile maneuver capability in an effort to improve overall performance. In this article, we will present a review of optimal control theory and derive a number of optimal guidance laws of increasing complexity. We also will explore the advantages that such guidance laws have over proportional navigation when engaging more stressing threats.

Basic Principles of Homing Guidance

Abstract: This article provides a conceptual foundation with respect to homing guidance upon which the next several articles are anchored. To this end, a basic geometric and notational framework is first established. Then, the well-known and often-used proportional navigation guidance concept is developed. The mechanization of proportional navigation in guided missiles depends on several factors, including the types of inertial and target sensors available on board the missile. Within this context, the line-of-sight reconstruction process (the collection and orchestration of the inertial and target sensor measurements necessary to support homing guidance) is discussed. Also, guided missiles typically have no direct control over longitudinal acceleration, and they maneuver in the direction specified by the guidance law by producing acceleration normal to the missile body. Therefore, we discuss a guidance command preservation technique that addresses this lack of control. The key challenges associated with designing effective homing guidance systems are discussed, followed by a cursory discussion of midcourse guidance for completeness' sake.

Air Force UAV's: The Secret History

Abstract: : In 1956, Air Force Maj. Gen. David Baker addressed a meeting of industrial leaders by stating, "We can readily see that except for certain types of missions, the manned combat aircraft will become technically obsolete in the future." He was referring not to UAVs, but to the possible replacement of manned penetrating bombers by intercontinental ballistic missiles and pilotless cruise missiles. The nuclear delivery mission was the keystone of Air Force organizational identity and budget share, however, so his statement speaks to the technological optimism of Air Force leaders concerning unproven, immature, and truly innovative aerospace technologies. His comments do not wholly concern missile developments, however, for at the same time the Air Force was also investing in fairly sophisticated, jet-powered target drones. Camera-carrying derivatives of those jet drones, operated by Air Force pilots, would soon evolve into the first significant combat UAV in history. This paper explores Air Force UAV systems using a comparative analytical framework. Systems are presented chronologically and the analysis focuses on external and internal variables contributing to weapon system innovation. As with the other services, independent externalities such as aviation technology, the military threat, and politics provide the context for Air Force UAV decision-making.

Triangle Intercept Guidance for Aerial Defense

Abstract: This paper focuses on missile guidance methods for a defense missile trying to protect a cooperative aircraft from guided munitions which have the same thrust and maneuvering capabilities as that of the defense missile. A novel guidance law for the missile-to-missile intercept is developed via a simple geometrical approach. The potential of the guidance law is demonstrated and discussed with some simulation results. Performance comparison with a conventional guidance law is also presented. Nomenclature C a = BM acceleration command normal to velocity vector, m/s 2 BA = blue aircraft (cooperative aircraft) BM = blue missile (defense missile) I, J = performance indices LOS ij = line-of-sight from the vehicle j to the vehicle i (Subscripts, A, M, and T denote BA, BM, and RM, respectively.) N , d N = navigation constant of the TG RM = red missile (attacking missile or target) ij r , ij r = LOS ij vector, and LOS ij range, m

Coning Motion of Spinning Missiles Induced by the Rate Loop

Abstract: Rate loops are commonly employed to compensate damping for missiles due to their great effectiveness. However, unexpected coning instability could be induced when the missile is subjected to spinning. The stability condition of coning motion as well as suitable design criteria for spinning missiles with rate loop feedback is established in this paper. Based on the theory of nutation movement, nonlinear equations of motion in terms of the nutation and precession angles are created. The mathematic model of the actuator acting on spinning missiles is also derived. After introducing the complex angle of attack into the system, the governing equations can be simplified to an analytically solvable second-order differential equation with respect to this variable. The proposed method is demonstrated by numerical simulations. It is also found that the upper limit of the critical gain in the damping loops' feedback for the stability decreases dramatically with a spinning missile. The dynamic inverse technique is employed as a remedy and it shows great effectiveness and applicability for the problem discussed in this paper.

Noncertainty-Equivalent Adaptive Missile Control via Immersion and Invariance

Abstract: [] This paper presents a noncertainty-equivalent new adaptive control (NCEA) system for the control of missile based on immersion and invariance approach. The mathematical model of the missile represents the nonlinear longitudinal dynamics, and it is assumed that all the aerodynamic parameters (except the sign of a single control input gain) are not known. Although, other choices of controlled output variables, which are nonlinear functions of state variables are possible, in this paper control of angle of attack is considered. A nonlinear adaptive autopilot for the trajectory control of the angle of attack is derived. The autopilot has a modular structure, which consists of a stabilizer and a parameter estimator. In the closed-loop system, it is seen that the trajectory is conflned to a manifold in the space of missile states and estimated parameters. A special feature of the designed autopilot (unlike the traditional certainty-equivalent (CEA) adaptive systems) is that, whenever the estimated parameters attain their true values, they remain frozen thereafter, and the autopilot recovers performance of deterministic control system. Simulation results are presented which show that the designed autopilot accomplishes trajectory control of the angle of attack despite uncertainties in the system parameters.

Overview of Missile Flight Control Systems

Abstract: The flight control system is a key element that allows the missile to meet its system performance requirements. The objective of the flight control system is to force the missile to achieve the steering commands developed by the guidance system. The types of steering commands vary depending on the phase of flight and the type of interceptor. For example, in the boost phase the flight control system may be designed to force the missile to track a desired flight-path angle or attitude. In the midcourse and terminal phases the system may be designed to track acceleration commands to effect an intercept of the target. This article explores several aspects of the missile flight control system, including its role in the overall missile system, its subsystems, types of flight control systems, design objectives, and design challenges. Also discussed are some of APL's contributions to the field, which have come primarily through our role as Technical Direction Agent on a variety of Navy missile programs.

Comparison of different Constitutive Models for Concrete in ABAQUS/Explicit for Missile Impact Analyses

Abstract: This EUR report describes numerical missile impact analyses on a reinforced concrete slab performed at JRC-IE using the Finite Element (FE) solver ABAQUS/Explicit. The FE model of the impacted reinforced concrete slab resembles a structure used in the missile impact tests by Hanchak et al. FE analyses with a hard (rigid) and a soft (deformable) missile are performed. Traditional Lagrangian formulation for both the missiles and reinforced concrete slabs are used. Two different build-in constitutive models for concrete in ABAQUS/Explicit, the Brittle Cracking Model and the Concrete Damaged Plasticity Model, are compared with each other and their suitability and limitations for missile impact analyses are explored. It turns out that only the Concrete Damaged Plasticity Model in combination with a soft missile leads to physically reasonable and sound results in terms of strains/stresses of the reinforced concrete slab, overall energy balances and overall deformation of the concrete slab when a traditional Lagrangian formulation is used for both missile and reinforced concrete slab. How to obtain EU publications Our priced publications are available from EU Bookshop (http://bookshop.europa.eu), where you can place an order with the sales agent of your choice. The Publications Office has a worldwide network of sales agents. You can obtain their contact details by sending a fax to (352) 29 29-42758. The mission of the JRC is to provide customer-driven scientific and technical support for the conception, development, implementation and monitoring of EU policies. As a service of the European Commission, the JRC functions as a reference centre of science and technology for the Union. Close to the policy-making process, it serves the common interest of the Member States, while being independent of special interests, whether private or national. LD -N A -2451-EN -C

Sliding Mode Integrated Missile Guidance and Control

Abstract: A novel method for the integrated guidance and control of a missile, called Sliding Mode Integrated Guidance and Control (SMIGC), is derived. In this formulation, the guidance and control systems are developed together and implemented in a single loop. One benefit of this “integrated” approach is that beneficial synergies between the guidance and control subsystems can be exploited. The design process can also be significantly reduced since only one control architecture must be designed. SMIGC makes use of the finite(time reaching phase of the sliding mode technique to ensure that a desired constraint, the Predicted Impact Point (PIP) heading error, is achieved in a finite time. Furthermore, the robustness of sliding mode to uncertainties is combined with a novel method for accounting for target acceleration to yield acceptable miss distances against maneuvering targets. An interesting and desirable aspect of the controller is that it does not require full information about the target acceleration. Further, to fully understand the effects of target acceleration on the problem, an in(depth target acceleration uncertainty analysis is performed. The effectiveness of the SMIGC approach is demonstrated through a series of simulations with a 6(DOF nonlinear missile model that includes all aerodynamic effects against maneuvering targets. Representative numerical results show that against agile targets, SMIGC obtains a high Hit(to(Kill accuracy consistently with reasonable fin deflections. �omenclature

The Effect of Tail Fin Parameters on the Induced Roll of a Canard-Controlled Missile

Abstract: Canard-controlled missiles often have adverse rolling moment characteristics that are difficult to predict. Several parameters that affect these rolling moment properties are investigated. An effort has been made to use a semi-empirical aerodynamic prediction code to determine the effects of varying tail geometry on the rolling moment. A generic canard controlled missile based on the NASA Blair configuration was examined at subsonic and supersonic Mach numbers with varying angle of attack and canard deflection angle. The span and area of the tail fins were systematically varied to determine the effects of each parameter alone on the rolling moment of the configuration.

Applying a novel extended Kalman filter to missile―target interception with APN guidance law: A benchmark case study

Abstract: This paper considers the estimation of the target acceleration with unknown dynamics along with other states of a benchmark example of a nonlinear 2D missile-target engagement system in presence of model uncertainties and measurement noises. The objective is to implement the augmented proportional navigation (APN) guidance law for the missile-target interception to minimize the distance between the missile and the target. The estimated target acceleration can be treated as an unknown input to the nonlinear 2D missile-target engagement system. A novel analytical recursive approach referred to as extended Kalman filter with unknown inputs without direct feedthrough (EKF- UI-WDF) is derived with the weighted least squares estimation for an extended state vector including states and unknown inputs which can be any type of signals without prior information. By applying the proposed EKF-UI-WDF approach to a 2D missile-target interception control system, simulation results demonstrate that this approach is capable of (i) estimating the states and unknown input (target acceleration) well, and (ii) achieving more reasonable interception performance comparing with the traditional extended Kalman filter (EKF) approach.

Missile guidance with constrained terminal body angle

Abstract: A novel missile guidance law that simultaneously minimizes the miss distance and imposes a predetermined terminal body angle is presented. The guidance law is derived by analyzing an interception scenario within the framework of a linear quadratic (LQ) terminal control problem. The derivation is for first order linear missile dynamics and a non-maneuvering target. The performance of the proposed guidance law is demonstrated in a planar interception scenario.

Noncertainty-Equivalent Adaptive Missile Control via Immersion and Invariance

Abstract: [] This paper presents a noncertainty-equivalent new adaptive control (NCEA) system for the control of missile based on immersion and invariance approach. The mathematical model of the missile represents the nonlinear longitudinal dynamics, and it is assumed that all the aerodynamic parameters (except the sign of a single control input gain) are not known. Although, other choices of controlled output variables, which are nonlinear functions of state variables are possible, in this paper control of angle of attack is considered. A nonlinear adaptive autopilot for the trajectory control of the angle of attack is derived. The autopilot has a modular structure, which consists of a stabilizer and a parameter estimator. In the closed-loop system, it is seen that the trajectory is conflned to a manifold in the space of missile states and estimated parameters. A special feature of the designed autopilot (unlike the traditional certainty-equivalent (CEA) adaptive systems) is that, whenever the estimated parameters attain their true values, they remain frozen thereafter, and the autopilot recovers performance of deterministic control system. Simulation results are presented which show that the designed autopilot accomplishes trajectory control of the angle of attack despite uncertainties in the system parameters.

Integrated missile flight control using quaternions and third-order sliding mode control

Abstract: An integrated autopilot-and-guidance algorithm that is robust to missile's model parametric and dynamic uncertainties is developed, using Higher Order Sliding Mode (HOSM) quasi-continuous control, for an endoatmospheric missile. Unlike in traditional guidance and control solutions that use a nested-loop design, a single loop integrated design of guidance/autopilot is proposed. 6-Degree of Freedom simulations of a simplified missile model with and without realistic actuator models, delays and parametric uncertainties confirms the efficacy of the proposed robust integrated HOSM guidance/control design.

Adaptive multiple sliding surface control for integrated missile guidance and autopilot with terminal angular constraint

Abstract: Integrated guidance and control for homing missiles with terminal angular constraint is considered. A time-varying integrated guidance and control model with unmatched uncertainties is formulated. In order to solve the multiple states regulation problem of such a system, an adaptive multiple sliding surface control algorithm is developed. It is proved that the proposed feedback controller can drive the missile to intercept the target accurately with a desired terminal attitude angle, and ensure the stability of the missile dynamics simultaneously. The 6DOF nonlinear missile simulation results demonstrate the feasibility of the proposed integrated guidance and control design scheme.

Design of guidance law for passive homing missile using sliding mode control

Abstract: This paper deals with a new passive homing guidance law for a stationary or a slowly moving target using sliding mode control technique. The proposed guidance law is designed in such a way of increasing the target observability. In addition, it can control a terminal impact angle to maximize warhead effect. The main idea is that we design the nonlinear sliding surface which makes an oscillatory missile motion along the desired impact angle frame so that the proposed guidance law introduces the LOS angle oscillation continuously. And then the enhanced target observability is achieved by comparing with the other conventional guidance laws such as the proportional navigation guidance. On the nonlinear sliding surface, lateral miss distance, velocity and acceleration command are converged to zero at terminal time. In addition, the proposed guidance law is robust with respect to autopilot lag and uncertainty. Finally, the performance of proposed guidance law is evaluated and demonstrated by number of simulations.

Dynamics and control of a gyroscope-stabilized platform in a self-propelled anti-aircraft system

Abstract: The paper presents a mathematical model of a triaxial gyroscopic platform on a moving platform base (motor vehicle). Control software platforms are designated with the inverse dynamics of the duties, while the control correction – using the LQR method. The considered platform can be used as an independent observation base for systems, cameras, parcels or gun machines. In the present study, it is shown in its application to stabilization of anti-aircraft missile launchers.

Missile guidance system

Abstract: In a CLOS missile guidance system, target and missile tracking data e.g. video image data are acquired on a UAV and transmitted to the missile where they are processed to provide guidance control data to the missile. Alternatively the video image data may be transmitted to a command station where the guidance control data is generated and transmitted to the missile, preferably via the UAV.

An analysis of the dynamics of a launcher-missile system on a moveable base

Abstract: The analysis focuses on the dynamics of a hypothetical anti-aircraft system during the missile launch. The results of the computer simulation of motion of the launcher-missile system are represented graphically. The diagrams provide information about the range of kinematic excitations acting on the opto-electro-mechanical coordinator of the target in a self-guided missile.

Remote air-to-air missile multi-platform cooperative guidance system and realization method thereof

Abstract: The invention discloses a remote air-to-air missile multi-platform cooperative guidance system and a realization method thereof. The multi-platform cooperative guidance system comprises an early warning plane module, a ground base radar detecting module, an aerial carrier module, a friend plane module, a remote air-to-air missile module, a target information space-time normalizing module and a target information fusion module. The realization method comprises the following steps: 1. an early warning plane, a ground base radar guided aerial carrier and a friend plane fly to a war area; 2. the radar of the friend plane operates to detect a target, and the aerial carrier launches an air-to-air missile; and 3. the targets detected by the radar of the early warning plane, the ground base radar and the radar of the friend plane are fused, and the friend plane carries out midcourse guide on the air-to-air missile. By adopting a multi-platform combat network formed by the early warning plane, the ground base radar, the aerial carrier and the friend plane to provide high-accuracy target information for the remote air-to-air missile, and adopting the friend plane to perform midcourse guide on the remote air-to-air missile, a remote target can be attacked stealthily and accurately.

Implementation of a reticle seeker missile simulator for jamming effect analysis

Abstract: In this paper, we implement a reticle seeker missile simulator on MATLAB-SIMULINK to analyze the jamming effect of the spin-scan and conscan reticle seeker. The DIRCM (Directed Infrared Countermeasures) system uses the pulsing flashes of infrared (IR) energy and its frequency and intensity have influence on the missile guidance system. Our simulation results show that jamming effect is indicated significantly when jammer frequency and reticle frequency are similar and present a 3D trajectory of missile motions by jamming.

Inertial Navigation for Guided Missile Systems

Abstract: An accurate inertial reference based on measurements of missile angular velocity and acceleration is needed for all of the major guidance and control functions of a guided missile. For example, intercept of a target would not be possible without a good inertial reference system to stabilize target line-of-sight measurements for the computation of missile guidance commands. This article provides an overview of inertial navigation for guided missiles. Missile navigation data (position, velocity, and attitude) are needed for missile guidance and control. Furthermore, this article describes in-flight alignment techniques that can be used to increase the accuracy of the missile navigation-system data by incorporating external non-inertial navigation-aiding data using a navigation Kalman filter. For guided missile systems, this aiding often is provided by an external radar track of the missile and/or Global Positioning System (GPS) receiver measurements. This article also discusses more recent advances in navigation for guided missiles. APL has been a major contributor to the development of many of the advanced guided missile navigation systems in use today.

Review on Empirical Studies of Local Impact Effects of Hard Missile on Concrete Structures

Abstract: Concrete is basic construction material used for any kind of structure However, in most vital and local structures such as nuclear plants, Power plants, Weapon Industries, weapons storage places, water retaining structures like dams, and also local industries, & etc, concrete structures have to be designed as defensive structures to provide protection against any accidents or knowingly generated incidents such as dynamic loading, dynamic local impact damage and global damage generated by kinetic missiles (steel rods, steel pipes, turbine blades, etc) The impacting missile (projectile) can be classified as ‘Hard’ and ‘Soft’ in nature, depending upon the implication of its deformation with respect to the deformation of target ‘Hard’ missile impact can generate both local impact damage and also overall dynamic global damage of concrete structure This paper only provides the review of previous empirical studies related to our study and can be used for making design recommendation and design procedures for determining the dynamic response of the target to prevent local and impact damage