Showing papers on "Missile published in 2009"

Guidance and Control of Missile Interceptor using Second-Order Sliding Modes

Abstract: A new smooth second-order sliding mode control (SSOSM) is proposed and proved for a system driven by uncertain sufficiently smooth disturbances. The main target application of this technique, the missile interceptor guidance-control system against targets performing evasive maneuvers, is considered to demonstrate benefits of this design for a two-loop integration of guidance and flight control systems. The designed guidance-control system performance is verified via computer simulations using a miniature hypervelocity kinetic energy endo-atmospheric interceptor planar model.

Integrated Higher-Order Sliding Mode Guidance and Autopilot for Dual Control Missiles

Abstract: An integrated autopilot and guidance algorithm, robust to target maneuvers and missile’s model uncertainties, is developed using higher-order sliding mode control for interceptors steered by a combination of aerodynamic lift, sustainer thrust, and center-of-gravity divert thrusters. A smooth higher-order sliding mode guidance that is robust to target maneuvers generates flight-path trajectory angular rates and attitude rate commands. The attitude rate missilemaneuvers are aimed at producing desired aerodynamic lift and/or orienting the sustainer thrust. The lateral acceleration created by themissile attitudemaneuver is treated as a cooperative disturbance andaccounted for by the trajectory control. Robust to themissile’smodel uncertainties, higher-order (second-order) slidingmode autopilot is designed based on a nonlinear dynamic slidingmanifold. The proposed integrated autopilot and guidance algorithm also includes seeker/tracker bore-sight stabilization and estimation of target lateral acceleration. The algorithm is tested using computer simulations against a ballistic maneuvering target.

L1 Adaptive Output-Feedback Controller for Non-Strictly-Positive-Real Reference Systems: Missile Longitudinal Autopilot Design

Abstract: This paper presents an extension of the L 1 adaptive output-feedback controller to systems of unknown relative degree in the presence of time-varying uncertainties without restricting the rate of their variation. As compared with earlier results in this direction, a new piecewise continuous adaptive law is introduced, along with the low-pass-filtered control signal that allows for achieving arbitrarily close tracking of the input and the output signals of the reference system, the transfer function of which is not required to be strictly positive real. Stability of this reference system is proved using a small-gain-type argument. The performance bounds between the closed-loop reference system and the closed-loop L 1 adaptive system can be rendered arbitrarily small by reducing the step size of integration. Missile longitudinal autopilot design is used as an example to illustrate the theoretical results.

Numerical State-Dependent Riccati Equation Approach for Missile Integrated Guidance Control

Abstract: A x , B x = state-dependent system matrices of size n n and n m Q x , R x = state-dependent weighting matrices of sizes n n and m m r, _ r = range and range rate of the target with respect to the missile S = solution to the Riccati equation T = rocket motor thrust per unit mass acting along the longitudinal axis of the missile u = control vector of size m 1 upert = control perturbation vector of size m 1 x = state vector of size n 1 xpert = state perturbation vector of size n 1 X, Y, Z = relative position components of the target with respect to the missile y, z = position of the moving masses along the pitch and yaw axes with respect to the body yc, zc = moving-mass position commands , = pitch and yaw Euler angles of the missile y, z = line-of-sight angles

Missile Autopilot Design: Gain-Scheduling and the Gap Metric

Abstract: This paper presents a systematic methodology for the synthesis of global gain-scheduled controllers for nonlinear time-varying systems. A controller of this type is used to compute the pitch-axis autopilot of an air-air missile. The missile model used is considered a benchmark for testing autopilot controllers in the academic and industrial communities. The missile's dynamics are linearized at a small set of operating points for which proportional― integral/proportional-type controllers are designed, to shape the frequency response of the linear plants' dynamics. A new set of operating points is computed afterward using the connection between the gap metric and the H ∞ loop-shaping theory. Then, reduced-order, static, H ∞ loop-shaping controllers are designed for this set of points using linear matrix inequality optimization techniques. Finally, the global gain-scheduled controller is obtained by interpolating the proportional―integral/proportional and the loop-shaping controllers' gains over the missile's flight envelope. The simulation results given show the generality and effectiveness ofthe proposed control strategy in terms of the operating point selection, stability, performance and robustness, of the closed loop.

Guidance law switching logic considering the seeker's field-of-view limits

Abstract: The impact time control guidance (ITCG) method, which has been proposed recently, can be applied successfully to a salvo attack of multiple missiles. Compared to the proportional navigation guidance law, this guidance method makes additional manoeuvres to synchronize the impact times. However, such manoeuvres do not consider the manoeuvrability and the seeker's field-of-view (FOV) of a missile and may cause the target to move out of the missile seeker's FOV; maintaining the seeker lock-on condition during the engagement is critical for missile guidance. To solve this problem, two methods are presented in this article: one is based on the calculation of minimum and maximum flight times considering the missile's manoeuvring limit and the seeker's FOV limit to check the available impact time. The other is based on guidance law switching logic that keeps the target look angle of the seeker constant. These methods can provide the boundary limit of the impact time of the salvo attack and prevent the loc...

Roll-Pitch-Yaw Integrated Robust Autopilot Design for a High Angle-of-Attack Missile

Abstract: This paper explores the feasibility of roll-pitch-yaw integrated autopilots for a high angle-of-attack missile. Investigation of the aerodynamic characteristics indicates strong cross-coupling effects between the motions in longitudinal and lateral directions. Robust control techniques based on H ∞ synthesis are employed to design roll- pitch-yaw integrated autopilots. The performance of the proposed roll-pitch-yaw integrated controller is tested in high-fidelity nonlinear 5-degree-of-freedom simulations. The proposed controllers are scheduled as a function of total angle of attack in a linear parameter varying framework with proportional navigation guidance laws. The integrated controller demonstrates satisfactory performance that cannot be achieved by the controller designed in a decoupled manner.

Cooperative Interception in a Multi-Missile Engagement

Abstract: Two estimators are presented, that enable cooperative target tracking of several missiles intercepting a single maneuvering target. The flrst estimator is a nonlinear adaptation of an interacting multiple model fllter, whereas the second estimator is a multiple model particle fllter. The paper develops the fllters for the cooperative and non-cooperative estimation modes, and investigates their individual estimation performance, using a nonlinear two-dimensional simulation. An extensive Monte Carlo study is used to demonstrate the viability of the cooperative estimation concept, for both estimators. It is shown that the closed loop interception performance of two cooperating missiles, guided by an optimal guidance law, improves, when compared to that of non-cooperating missiles. The particle fllter based estimator demonstrates hit-to-kill closed loop interception performance in the cooperative mode, but requires higher computational load than the extended Kalman fllter based estimator, making the choice of estimator a tradeofi between performance and computational power.

Genetic-Algorithm Optimization of Liquid-Propellant Missile Systems

Abstract: The use of computer programs driven by genetic algorithms has become an increasingly popular method for optimizing engineering designs. This paper focuses on the modeling and optimization of liquid-rocket-enginepropelled missiles with a computer program that is composed of a series of codes that simulate the performance of liquid-propelled rockets andare controlledby a genetic algorithm.Because the entiremissile design is considered, the complete system performance must be modeled accurately and efficiently. The methodology described in this paper has been extended to both the preliminary design and reverse-engineering of liquid-propelled missiles. The performance model has been validated against the performance of a liquid-propelled missile and the results are provided. Two different fast-aerodynamic-prediction codes are used and their results are compared. A complete preliminary design of a liquid-propelled missile system is considered with a variety of goals and constraints. Results from the preliminary design optimization are also shown and discussed in detail.

Theoretical Predictions of F-16 Fighter Limit Cycle Oscillations for Flight Flutter Testing

Abstract: A computational investigation of the flutter onset and limit cycle oscillation behavior of various F-16 fighter weapons and stores configurations is presented. A nonlinear harmonic balance compressible Reynolds-averaged Navier–Stokescomputational fluiddynamic flowsolverisusedtomodeltheunsteadyaerodynamicsoftheF-16wing. Slender body/wing theory is used as an approximate method for accounting for the unsteady aerodynamic effects of wing-tip launchers and missiles. Details of the computational model are presented along with an examination of the sensitivity of computed aeroelastic behavior to characteristics and parameters of the structural and fluid dynamic model. Comparisons with flight-test data are also shown. I. Introduction T HE SEEK EAGLE Office at Eglin Air Force Base performs an essential task in clearing new aircraft/stores configurations through flight tests for safe and effective operation. Many of these flighttestsarefortheF-16aircraftwhichcontinuestobeaworkhorse for the U.S. Air Forcewith continually new stores (missiles, bombs, and fuel tanks) being considered for aircraft operations. Similar aeroelastic flight tests are expected for future fighter aircraft as they go into service in the coming years. The number of needed flight tests is projected to be well beyond the financial and staff resources available. Hence there is a pressing need to identify the most critical aircraft/store configurations for the limited flight-test resources available and also insofar as possibly reduce the number of flight tests needed. Virtual flight testing may be the answer. Using new improved computational capability that provides much more rapid solutions, computational simulation can help identify the most critical aircraft/ store configuration and also hasthe potential of reducingthe number ofneeded flighttestsifconfidencecanbeestablishedinthecapability of simulations to correlate with flight-test data. A new methodology has been developed to produce these computer simulations based upon the notion that because the response is periodic in time, the solution need only be obtained over a single period of oscillation in time. By avoiding the traditional time marching solution which computes the long transient before a steady-state periodic oscillation is reached, computational times are reduced by a factor of 10–100. This enables a sufficiently rapid solutiontomakesuchsimulationsapracticalrealityforthe flight-test engineer and support team. Future developments of this methodology hold the promise of further substantial reductions in computational cost and are being vigorously pursued. Also further refinements in the physical fidelity of the simulation models are being considered.

System and method for guiding and controlling a missile using high order sliding mode control

Abstract: Higher Order Sliding Mode (HOSM) control techniques are applied to the Guidance Control (G&C) of interceptor missile in which velocity may be steered by combination of main thrust, aerodynamic lift and lateral on-off divert thrusters, and attitude may be steered by continuous or on-off actuators. Methods include the pointing of the seeker, its associated estimation processes, a guidance law that uses concurrent divert mechanisms, and an attitude autopilot. The insensitivity of the controller to matched disturbances allows the concurrent usage of the divert mechanisms without adverse effect on the accuracy. The controller also allows the de-coupling of the control of roll, pitch and yaw channels, and usage quaternions to represent body attitude and it provides control perfect robustness. While it conceivable to design separately the components of the G&C method, it is widely accepted that designing them in an integrated fashion usually produces a better result.

Discrete element modelling of missile impacts on a reinforced concrete target

Abstract: A three-dimensional Discrete Element Method (DEM) is used to study the penetration and perforation process of a concrete target subjected to rigid flat-nose-shaped missile impacts. The evolution of the missile velocity is compared with real test cases made by the French Atomic Energy Agency (CEA) and the French Electrical Power Company (EDF). The perforation limits observed in the experimental data are well predicted by the three-dimensional discrete element model. Parametric studies are then carried out to show the respective roles of the mechanical components, such as the dependence of the perforation process on the percentage of reinforcement.

Trajectory Prediction for a Typical Fin Stabilized Artillery Rocket

Abstract: This paper investigates the trajectory prediction and dispersion for unguided fin stabilized artillery rocket in order to explain the importance of the rocket production accuracy and the benefit of using guided rockets. The total dispersion results mainly from three effects. The first is the dispersion due to rocket production inaccuracy, which includes propellant mass, composition inaccuracy, rocket total mass, axial and lateral moments of inertia and resultant center of gravity. The second dispersion during boosting phase which includes launcher deflection, missile tip-off from the launcher, thrust and fin misalignments, and atmospheric disturbances such as tail wind, cross wind, and gusts. While the third is thedispersion during free-flight phase that is due to the fluctuations in wind profile. In this study, a trajectory calculation using a 6-DOF model was developed and applied for a typical artillery rocket, the 122 mm artillery rocket, at different mass and flight properties to predict the trajectory parameters and dispersion.

Trajectory optimization for hypersonic boost‐glide missile considering aeroheating

Abstract: Purpose – The purpose of this paper is to optimize the downrange for hypersonic boost‐glide (HBG) missile under near‐real condition, and to validate the suitability of proposed wall cooling materials.Design/methodology/approach – The trajectory optimization problem is characterized by a boost phase followed by a glide phase. A multi‐phase trajectory optimization tool is adopted to optimize the downrange. The associated optimal control problem has been solved by selecting a direct shooting method. The dynamics has been transcribed to a set of nonlinear constraints and the arising nonlinear programming problem has been solved through a sequential quadratic programming solver. An aerothermodynamics analysis method is introduced to calculate the aerodynamic heating at nose, leading edge, and ventral centerline regions.Findings – HBG missile is suitable for long‐range attack, and the optimal trajectory solved is a novel boost‐glide‐skip trajectory, i.e. boost firstly, glide secondly, and skip at last. The prop...

Missile avoidance manoeuvres with simultaneous decoy deployment

Abstract: Anti-missile defence scenarios for aircraft involves an evasive manoeuvre and the simultaneous deployment of countermeasures to deceive the threat. In this paper both aspects are considered and an original analytical expression for the miss distance at intercept is derived. The evader aircraft has control over (1) its lateral acceleration in order to perform the evasive strategy, and (2) the deployment of its decoys (time of launch and direction of launch). The pursuer missile uses proportional navigation and it needs a fixed time delay to discriminate between the decoy and the target. Planar linear engagement geometries and a first-order guidance system are considered. As an example, simulation cases are presented in order to study the optimal deployment of the decoy(s) with, and without, performing the so-called vertical-S manoeuvre.

Spin-Yaw Lock-In of a Rotationally Symmetric Missile

Abstract: Spin―yaw lock-in has usually been associated with small aerodynamic asymmetries that produce a trim angle that rotates with the missile. Recent trajectory calculations using wind-tunnel measurements ofthe aerodynamic moment acting on a rotationally symmetric missile predicted spin―yaw lock-in. Thus, an asymmetry-induced trim moment is not necessary for lock-in. For a side moment that varies with the roll angle between the angle-of-attack plane and a fixed plane on the missile, this paper shows that lock-in of the coning motion and the rolling motion can occur. The conditions on the aerodynamic moments for occurrence of lock-in are derived and appropriate stability criteria are obtained. A simple case of a hypersonic finned missile is considered and the existence of lock-in is demonstrated.

Microwave radar sensors for active defense systems

Abstract: Modern anti-armor missiles represent enormous threat for any military vehicle. Simple hand-held missiles are able to penetrate 300 mm, more sophisticated missiles up to 1000 mm of the best steel armors. Active defense methods seem to be promising ways how to face this problem. Systems of active defense are based on sensors able to detect and measure approaching threat missile and generate signals that activate a suitable counter-measure. The paper describes several different designed radar sensors tested in active defense configurations. This concerns narrowband sector sensor, “microwave curtain” and a new wideband sensor with a distance measurement capability. All practical tests were performed using real missiles at army shooting ranges.

Research on missile formation control system

Abstract: The control system of missile formation flight is described. In this paper, the missile formation flight controller and BTT intelligence missile outer-loop controller are interested. A combination of proportion and differential control of the position error is used in the designing of the missile formation flight controller to generate the control command for the single missile. The missile outer-loop controller is designed based on the total energy. Simulation demonstrate that the missile formation controller are capable of controlling the missile formation and the BTT missile controller can track the command generated by the missile formation controller exactly.

Target tracking enhancement using a Kalman filter in the presence of interference

Abstract: In this paper we present a new target tracking enhancement system that uses a Kalman filter in the presence of interference. If the radar (seeker) is affected by different types of interference, this will affect the missile trajectory towards the target and may cause inaccurate tracking. In the new system a six-state Kalman filter is utilized to perform the tracking task and to carry out smoothing to the corrupted trajectory. This also provides good information about the target velocity in three dimensions which is very important information about the target. A three dimensional scenario between target (with high manoeuvre) and missile is used to illustrate the performance of the system in the case when (i) no interference is present and (ii) interference is present. The performance of the filtered trajectory using the Kalman tracker will be assessed for different guidance methods: including (i) proportional navigation (ii) pure pursuit and (iii) constant bearing. The Kalman improvement for the tacking for the three guidance method will be analysed.

Method and system for controlling solid propellant thrust

Abstract: The present invention relates to a solid rocket motor control system that uses state of the art electrically ignited, extinguishable and throttleable propellants. In embodiments, the control system is implemented by a software algorithm embedded in missile flight controllers and contains all control system elements and control compensation that performs thrust control for many defense and space rocket motor applications, for example.

A Method Based on Genetic Algorithm for Anti-ship Missile Path Planning

Abstract: This paper presented a novel approach to search and optimize path points for anti-ship missile path planning. We utilized the method of MAKLINK graph to construct free space, and then, a global state connected graph is built up for searching for all possible routes. Genetic Algorithm is used to search and optimize path points severally in these local routes. According to flight rules and technique characteristics of anti-ship missile, a coding measure and the fitness function are proposed. By introducing constraints selection strategy, the method in this paper guaranteed that all paths resulted could satisfy constraints. The results by simulation revealed that the method proposed in this paper is effective in solving the path planning problem with constraints.

An efficient imaging algorithm for missile-borne side-looking SAR

Abstract: This paper presents an efficient imaging algorithm for missile-borne SAR. In the course of missile's diving flight, it is difficult to use the traditional imaging principle because of the strong coupling between range and azimuth signal. Through the fourth order range model, the echo model is established and the two-dimensional point target spectrum is derived by the method of series reversion. Then, the corresponding range and azimuth match filtering functions are given. Finally, simulation results demonstrate the accuracy and validity of the proposed algorithms.

Analytical analysis of lock-on range of infrared heat seeker missile

Abstract: The lock-on range, R, of infrared (IR) heat seeker missile has been studied in details. An analytical expression form of R has been derived in term of target (assumed as aircraft hot metal tailpipe), atmosphere and infrared (IR) detector parameters. The R is represented in term of “Lambert W function” and as a function of target temperature, T, and atmosphere extinction coefficient, α. Assuming clear sky background, the simulation results show: firstly, R increases as the target temperature increases. High value of R is achieved when the temperature of aircraft engine is high especially in takeoff case where the engine thrust is large. Secondly, R can be decreased by (1) increasing the attenuation coefficient of the propagation medium, i. e. by cooling the aircraft hot engine or any other obscurants that absorb or scatter the IR radiation (2) or by decreasing emissivity , ε, of the target surface which depends on the applied coatings / paintings. The decreasing of R will increase the probability of aircraft survivability assuming effective aircraft early warning system where the detecting range of the latter should be larger than the missile lock-on range.

Integrated guidance and control for a missile in the pitch plane based upon subspace stabilization

Abstract: Based upon the model of the relative motion between the missile and the target and the model of the missile moving in the pitch plane, a novel integrated guidance and control model under the error tracking meaning is established. It is different from the state-equation-based integrated guidance and control models, it emphasizes particularly on attitude angles tracking. Especially, the realization of the error tracking properties is turned into the corresponding parameters adjusting by subspace stabilization technology. The globally asymptotical stabilization of the error tracking system is realized by an observer with the adaptive adjusting function about the model error and about the target maneuverability. The simulation results indicate that the design is effective.