Showing papers on "Missile published in 2012"

Sliding-Mode Guidance and Control for All-Aspect Interceptors with Terminal Angle Constraints

Abstract: In this paper, sliding-mode-control-based guidance laws to intercept stationary, constant-velocity, and maneuvering targets at a desired impact angle are proposed. The desired impact angle, which is defined in terms of a desired line-of-sight angle, is achieved in finite time by selecting the missile's lateral acceleration to enforce terminal sliding mode on a switching surface designed using nonlinear engagement dynamics. The conditions for capturability are also presented. In addition, by considering a three-degree-of-freedom linear-interceptor dynamic model and by following the procedure used to design a dynamic sliding-mode controller, the interceptor autopilot is designed as a simple static controller to track the lateral acceleration generated by the guidance law. Numerical simulation results are presented to validate the proposed guidance laws and the autopilot design for different initial engagement geometries and impact angles.

Morphing Aerospace Vehicles and Structures

Abstract: Numerical Techniques Fundamentals of Tactical Missile Guidance Method of Adjoints and the Homing Loop Noise Analysis Covariance Analysis and the Homing Loop Proportional Navigation and Miss Distance Digital Fading Memory Noise Filters in the Homing Loop Advanced Guidance Laws Kalman Filters and The Homing Loop Tactical Zones Strategic Considerations Boosters Lambert Guidance Strategic Intercepts Miscellaneous Topics Ballistic Target Properties Extended Kalman Filtering and Ballistic Coefficient Estimation Multiple Targets Weaving Targets Representing Missile Airframe with Transfer Functions Introduction to Flight Control Design Three-Loop Autopilot Trajectory Shaping Guidance Filtering and Weaving Targets Alternative Approaches to Guidance Law Development Filter Bank Approach to Weaving Target Problem Engagement Simulations in Three Dimensions Advanced Adjoint Applications Miscellaneous Tactical Missile Guidance Topics Comparison of Differential Game Guidance with Optimal Guidance Kinematics of Intercepting a Ballistic Target Boost-Phase Filtering Options Kill Vehicle Guidance and Control Sizing for Boost-Phase Intercept Appendix: Additional Examples Index Supporting Materials

Sliding Mode and Inertial Delay Control Based Missile Guidance

Abstract: A novel formulation of sliding mode control (SMC) based proportional navigation (PN) guidance law is presented. Unlike conventional SMC-based guidance laws, the law presented here does not need any knowledge of bounds of target acceleration. The target acceleration is estimated using the so-called inertial delay control (IDC). Closed-loop stability for the guidance loop is established. Simulations are carried out by considering highly-maneuvering targets and constant as well as varying missile velocity and the results are presented to demonstrate the effectiveness of the proposed formulation.

Guidance Strategies Against Defended Aerial Targets

Abstract: The three-body engagement scenario is considered where an aerial missile, homing onto a target aircraft, encounters a defender missile launched by the aircraft. A given set of defender-missile guidance laws, namely, proportional navigation and line-of-sight guidance, are considered, and analysis is carried out for proportional navigation and pure pursuit attacking missile strategies. Analytic launch envelopes and lateral accelerations ratios for the four resulting scenarios are derived. Closed-form expressions for attackingmissile initial position and launch angles are derived for a successful evasion from the defender. Numerical simulations are carried out which comply with the analytical findings. The resulting launch envelopes are highly sensitive to the adversary’s guidance strategy, and thus, concerning practical implementation, a game theoretic analysis is carried out. Game solutions are obtained in pure and mixed strategies, resulting in the missile’s evasion probability envelope, that is, based on the four individual launch envelopes.

Optimality of Linear Time-Varying Guidance for Impact Angle Control

Abstract: The work presented here demonstrates optimality of linear time-varying guidance laws for controlling impact angles as well as terminal misses using the inverse problem of optimal control theory. Under the assumptions of a stationary target and a lag-free missile with a constant speed and a small flight path angle, it is shown that a generalized linear time-varying guidance law with an arbitrary pair of guidance coefficients can minimize a certain quadratic performance index subject to the terminal constraints. Feasible sets of the guidance coefficients for capturability are investigated and explicitly stated by yielding the closed-form solutions of the guidance loop in this paper. These results imply that it is possible for more realistic settings of the guidance coefficients either to improve the guidance performance or to achieve some specific guidance objectives in practical environments. Furthermore, the time-to-go calculation methods, which consider the closed-loop solutions for arbitrary guidance coefficients, are included for implementation of the guidance law. Linear and nonlinear simulations are performed to validate the proposed approach.

Hybrid Guidance Law for Approach Angle and Time-of-Arrival Control

Abstract: A new derivation of an optimal missile guidance law is presented. The guidance law derived is an extension of previously published versions of explicit guidance in that gravity compensation is incorporated in an optimal fashion. Performance of the extended guidance law is demonstrated via simulation comparisons with generalized vector explicit guidance andKappa guidance, both of which have appeared previously in the literature. The new law is then incorporated into a hybrid guidance scheme that allows the specification and satisfaction of a desired time-of-arrival constraint. Performance of this capability is also demonstrated via simulation examples. Finally, an approach to coordinating desired times of arrival among a group of cooperatingmissiles to achieve simultaneous target intercepts or missile rendezvous is presented.

Extended state observer-based robust pitch autopilot design for tactical missiles

Abstract: In this article, a new design based on the extended state observer (ESO) technique for the pitch autopilot of a tail-controlled, skid-to-turn missile is proposed. The pitch-plane dynamics with the ...

Aircraft and missile afterbody flow control device and method of controlling flow

Abstract: An afterbody flow control system is used for aircraft or missile flow control to provide enhanced maneuverability and stabilization. A method of operating the flow control system is also described. The missile or aircraft comprises an afterbody and a forebody; at least one activatable flow effector on the missile or aircraft afterbody; at least one sensor having a signal, the at least one sensor being positioned to detect forces or flow conditions on the missile or aircraft afterbody; and a closed loop control system; wherein the closed loop control system is used for activating and deactivating the at least one activatable flow effector based on at least in part the signal of the at least one sensor.

Non-singular terminal sliding mode guidance and control with terminal angle constraints for non-maneuvering targets

Abstract: In this paper guidance laws to intercept stationary and constant velocity targets at a desired impact angle, based on sliding mode control theory, are proposed. The desired impact angle, which is defined in terms of a desired line-of-sight (LOS) angle, is achieved in finite time by selecting the missile's lateral acceleration (latax) to enforce non-singular terminal sliding mode on a switching surface designed using this desired LOS angle and based on non-linear engagement dynamics. Numerical simulation results are presented to validate the proposed guidance laws for different initial engagement geometries and impact angles.

System, device and method of protecting aircrafts against incoming missiles and threats

Abstract: The present invention includes a system for protecting an aircraft against one or more incoming threats. The system includes one or more electro-optic sensors to scan an area around the aircraft for one or more possible incoming threats, and to generate an indication signal once an incoming threat is detected; an integrated unit combining a Missile Approach Confirmation Sensor (MACS) with Directed Infra-Red Counter Measure (DIRCM), to verify the incoming threat and to activate a countermeasure against the verified incoming threat; and a processor to receive data from said one or more electro-optic sensors and the integrated MACS-DIRCM unit, and to select a countermeasure technique for deployment against the incoming threat.

Exo-Atmospheric Guidance of an Accelerating Interceptor Missile

Abstract: An exo-atmospheric interception scenario between an accelerating missile and its target is investigated. It is assumed that the maneuvering acceleration is obtained by instantaneous rotation of the missile's body to the required attitude. Two different guidance laws are derived for such an interceptor using the sliding mode control methodology. The difference is in the definition of the sliding surface enforcing different trajectories for the interceptor. It is shown that if this surface is chosen as the zero-effort-miss of the well-known proportional navigation guidance law, then the missile is commanded to point its acceleration vector along the line-of-sight and consequently fly along a curved trajectory. For the second guidance law, a unique sliding surface is chosen enforcing the missile to fly on a straight line towards collision, after the initial heading error is nulled. The performance of the guidance laws is analyzed and compared using a nonlinear two dimensional simulation. It is shown that on top of enforcing a different flight geometry for the interceptor, the use of the new guidance to collision sliding mode guidance law can enhance the capture zone of the interceptor.

Robust autopilot design of uncertain bank-to-turn missiles using state-space disturbance observers

Abstract: Robust autopilot design for bank-to-turn (BTT) missiles under disturbances and uncertainties is investigated in this article using the disturbance observer concept. It is well known that the BTT missile dynamics undergo substantial change during its flight. In this disturbance observer-based control (DOBC) setting, the influences caused by parameter variations are merged into disturbance terms and regarded as parts of the lumped disturbances. Disturbance observers are employed to estimate the lumped disturbances, and then a disturbance observer-based robust control (DOBRC) method is proposed in this article to compensate the influences of parameter variations and the disturbances from the output channels. Similar to the baseline linear quadratic regulator design, the DOBRC is analysed and designed using linear techniques. Very promising performance has been achieved for the BTT missile as shown in simulation. It is demonstrated that DOBC approach provides a simple, intuitive, and practical solution for many challenging control problems where systems are subject to significant external disturbances, and uncertainties such as BTT missiles.

Predictive Functional Control-Based Missile Autopilot Design

Abstract: Predictive functional control is applied to design a missile autopilot considering control constraints. The missile nonlinear dynamics are first transformed into a linear structure with state-dependent coefficient matrices. At each sampling instant, the internal state-space model for prediction is obtained through a normal discretization procedure. Based upon this model, a predictive functional controller for the nonlinear missile autopilot is proposed. Compared with the conventional predictive control algorithms, which usually involve an online quadratic programming in the practical implementation, the new controller demands less online computation resources. Simulation results validate the effectiveness and robustness of the proposed algorithm.

Adaptive-sliding-mode guidance law design for missiles with thrust vector control and divert control system

Abstract: This approach addresses an adaptive-sliding-mode guidance law for missiles equipped with thrust vector control and divert control system, for the task of intercepting of a theatre ballistic missile with manoeuvring acceleration. The aim of the present study is to achieve a bounded target interception under the three degrees of freedom (DOF) control and the zero-effort-miss phase to improve the intercepting accuracy of the missile by narrowing the distance between the intercepting missile and the target missile to within effective range for triggering the missile’s explosion. Considering the external disturbances and the variation in the missile’s mass, the 3 DOF adaptive-sliding-mode guidance law is designed using adaptive control and sliding-mode control to minimise the distance between the centre of the intercepting missile and that of the target missile asymptotically. Also, the overall system stability is verified using the Lyapunov stability theory. Furthermore, successful simulation results are conducted to validate the effectiveness of the proposed adaptive-sliding-mode guidance law.

L1 adaptive augmentation of a missile autopilot

Abstract: An L1 control augmentation is developed based on a novel autopilot structure for a highly agile, tail-controlled missile. The autopilot is designed to control the pitch and yaw accelerations in the body-fixed frame while maintaining a desired roll angle. In order to compensate for the undesirable effects of modeling uncertainty, an L1 adaptive controller with piece-wise constant adaptation laws is developed that explicitly takes into account the multi-loop dynamic structure of the autopilot. The benefits of the augmented design are demonstrated via Monte-Carlo simulations of a high-fidelity 6DOF missile model, in which a wide spectrum of uncertainty combinations is considered.

Adaptive Augmentation of a New Baseline Control Architecture for Tail-Controlled Missiles Using a N onlinear Reference Model *

Abstract: An adaptive augmented Nonlinear Dynamic Inversion (NDI) control structure based on a nonlinear reference model is developed for a highly agile, tail-controlled surface-to-air missile. Due to a modular control design, the missi le is able to perform skid-to-turn (STT) and bank-while-turn (BWT) maneuvers. The focus of this paper lies on the nonlinear reference model and the respective change in the NDI error feedback architecture. In order to overcome the non-minimum phase property of tail-controlled missiles, a conversion of the pitch- and yaw-acceleration, which renders the syst em minimum phase is conducted. A nonlinear reference model featuring all the main nonlinear effects of the missile dynamics is developed. With such a nonlinear command filter, the closed-loop system fully exploits the physical capabilities of the plant. Based on a clas sical two-loop inversion strategy, the design of the NDI error feedback baseline controller is ad justed in accordance to the modification of the reference model. By making use of NDI, the t racking problem is transformed into a stabilizing problem with an almost linear error dyn amics. This makes the system perfectly suitable for using a Model Reference Adaptive Control (MRAC) approach as an adaptive layer, which is able to cope with modeling errors a nd sensor biases. The autopilot design proved its robustness and performance under a large spectrum of uncertainty combinations using Monte Carlo simulations.

An Adaptive Backstepping Controller for a Hypersonic Air-Breathing Missile

Abstract: This paper presents the development of an adaptive controller for a hypersonic air-breathing missile with terminal constraints. The controller is designed to regulate the longitudinal dynamics of a hypersonic vehicle model via a backstepping approach. The backstepping approach is used to compensate for uncertainties in the dynamics that do not satisfy the matching condition while ensuring asymptotic tracking of a desired velocity profile and asymptotic regulation of the vehicle position, angle of attack, body angle, and angular rates. A Lyapunov-based stability analysis is used to prove the asymptotic regulation of the controlled states. Simulation results are presented to verify the performance of the controller.

Missile acceleration controller design using proportional–integral and non-linear dynamic control design method:

Abstract: In this article, a method to make the normal acceleration of medium angle of attack missile to track some desired trajectories is presented. The dynamics which describe a missile acceleration are k...

Proportional Guidance and CDM Control Synthesis for a Short-Range Homing Surface-to-Air Missile

Abstract: This paper presents a high-performance and simple guidance method for a short-range missile. Missile control systems are designed along the estimated flight envelope of the missile based on coefficient diagram method. The linear time invariant (LTI) models are obtained by linearizing missile nonlinear model at specific operating points. The governing equations of motion are accurately modeled by taking into account the effect of the mass rate and the center of gravity shift. A novel algebraic approach called coefficient diagram method (CDM) is used to design normal acceleration and roll angle control systems. The proportional navigation law is considered as the guidance law. One case of engagement scenarios is studied and the missile performance is evaluated. The robustness of the proposed controller is tested against parameter uncertainties and wind disturbance.

Missile Homing-Phase Guidance Law Design using Reinforcement Learning

Abstract: A new approach to missile guidance law design is proposed, where reinforcement learning (RL) is used to learn a homing-phase guidance law that is optimal with respect to the missile’s airframe dynamics as well as sensor and actuator noise and delays. It is demonstrated that this new approach results in a guidance law giving superior performance to either PN guidance or enhanced PN guidance laws developed using Lyapunov theory. Although optimal control theory can be used to derive an optimal control law under certain idealized modeling assumptions, we discuss how the RL approach gives more flexibility and higher expected performance for real-world systems. Nomenclature

Sliding mode control based terminal impact angle constrained guidance laws using dual sliding surfaces

Abstract: In this paper, sliding mode control theory based guidance laws to intercept non-maneuvering targets at a desired impact angle are presented The desired impact angle, defined in terms of a desired line-of-sight (LOS) angle, is achieved by selecting the missile's lateral acceleration (latax) to enforce sliding mode on a sliding surface based on this LOS angle As will be shown, this guidance law does not ensure interception for all states of the missile and the target during the engagement Hence, to satisfy the requirement of interception at the desired impact angle, a second sliding surface is designed and a switching logic, based on the conditions necessary for interception, is presented that allows the latax to switch between enforcing sliding mode on one of these surfaces so that the target can be intercepted at the desired impact angle The guidance laws are designed using non-linear engagement dynamics

Aerodynamic Shape Optimization System of a Canard-Controlled Missile Using Trajectory-Dependent Aerodynamic Coefficients

Abstract: This paper describes a shape optimization study to maximize the range of a guided missile. To design a guided missile havingmaximum range, a shape optimization system is incorporated with a trajectory analysis program and an optimization technique. In particular, trajectory-dependent aerodynamic coefficients are fully considered. In the trajectory analysis step, a component buildup method is directly connected to the equation of motion to calculate aerodynamic coefficients at every time step. In the optimization step, a real-coded adaptive range genetic algorithm is adopted to determine the optimal shape of the global maximum range. The shape optimization system of a guided missile can maximize the range of the missile and yield the optimal shapes of canards and tailfins. Finally, the effects of trajectory-dependent aerodynamic coefficients, guidance, and control on the range of a missile are illustrated.

On the powered flight trajectory of an earth satellite

Abstract: : The actual choice of a powered flight trajectory depends upon details of booster design and upon the particular orbit selected, the latter being determined by instrumentation and data requirements. Independent of these considerations, however, it is of interest to ask what type of trajectory will be optimal from the standpoint of missile efficiency. The approach used assumes that the characteristics of the booster, i.e., its thrust and mass as functions of time, are specified. The problem is then to determine what powered flight trajectory will put the satellite into an orbit of maximum altitude.