Showing papers on "Missile published in 2013"

Augmented Polynomial Guidance With Impact Time and Angle Constraints

Abstract: In this paper, a new impact time and impact angle control guidance law for homing missiles against a stationary target is proposed. To derive a state feedback command for the proposed law with terminal constraints on the impact time and angle, we introduce a polynomial function of the guidance command with three unknown coefficients, one of which is determined to achieve the impact time requirement. The others are determined to satisfy the terminal impact angle constraint, as well as the zero miss distance. Because the proposed guidance law has arbitrary guidance gains, it is possible to shape the intercept trajectory and acceleration command profile by choosing proper gains in relation to the missile's capability, operational conditions, and so on.

Missile Guidance Law Based on Extended State Observer

Abstract: The problem of missile interception is investigated in this paper. Sliding-mode controllers with finite-time convergence are proposed for the design of guidance laws. In particular, the control design consists of the estimation of the target acceleration by the extended state observer (ESO), and thus, it achieves the decrease of undesired chattering effectively. Also, simulation results are presented to illustrate the effectiveness of the control strategies.

Bias-Shaping Method for Biased Proportional Navigation with Terminal-Angle Constraint

Abstract: F OR decades, many advanced guidance laws with terminalimpact-angle constraints have been devised to maximize the warhead effect of antiship or antitank missiles and to ensure a high kill probability. The proposed suboptimal guidance with a terminal body-attitude constraint for reentry vehicles in [1] appears to be the first attempt to design an impact-angle-constrained guidance. In [2], an energy-optimal impact-angle control lawwas proposed by solving the linear quadratic optimal control problem with arbitrary missile dynamics. As an extension of this study, the authors also proposed an optimal impact-angle controller that can minimize the time-to-go weighted energy-cost function [3]. Using the Schwartz inequality and differential game theory, terminal-impact-angle-constrained guidance laws for maneuvering targets were developed in [4,5], respectively. To intercept a stationary target with zero terminal acceleration as well as a specified impact angle, the guidance law called the time-to-go polynomial guidance law was suggested in [6], where the acceleration command was assumed initially as a polynomial function of the time-to-go with two unknown coefficients. In [7], a modified proportional navigation (PN) guidancewith a time-varying bias, which is a function of relative range, was proposed using a nonlinear planar engagement and Lyapunov-like function. Although various guidance laws to control the impact angle have been developed so far, most of these laws are difficult to implement, especially for a passive homing missile equipped with an infrared seeker, because an accurate time-to-go estimation or range information is required. The authors of [8–10] proposed two-phase guidance schemes with terminal-angle constraints on the basis of the conventional PN guidance. The guidance scheme suggested in [8] comprises PN guidance with a navigation gainN < 2 for covering all impact angles from 0 to −π and PN with N 2 for intercepting stationary targets with a desired impact angle in surface-to-surface engagements. This guidance scheme was further extended to the case of moving targets in [9]. Using the biased PN (BPN) guidance, the authors of [10] developed a similar two-phase scheme in which the missile follows BPN with a constant bias for the initial homing phase and then switches to PN (i.e., BPN with zero bias) when the integral value of the bias satisfies a certain value calculated from initial engagement conditions and desired impact angle. Because these two-phase guidance schemes only use the line-of-sight (LOS) rate information for the impact-angle control, they can be applied to passive homingmissile systems. However, these guidance schemes have some drawbacks. 1) If the limitation of missile acceleration capability exists, a large miss distance or impact-angle error is generated. 2) A higher look angle, which may result in seeker lock-on failure or instability, is produced in the first-phase guidance. To overcome the drawbacks resulting from the look angle and acceleration limits, we first propose a bias-shaping method based on the two-phase BPNguidance scheme of [10], which can achieve both the terminal-angle constraint and look-angle limitation to maintain the seeker lock-on condition. Next, we investigate analytically the guidance performance of BPN with the proposed bias-shaping method in consideration of the limited acceleration capability. The proposed bias-shaping method consists of two time-varying biases and switching logic similar to the proposed logic of [11] and only requires the LOS rate information to generate the guidance command, thereby easily implementing the proposed law in practical passive homing missiles.

Optimal impact angle control guidance law considering the seeker’s field-of-view limits:

Abstract: A new optimal guidance problem with impact angle constraint and seeker’s field-of-view limits is investigated for a missile with a strapdown seeker. Impact angle control to satisfy the terminal fli...

Linear Quadratic Optimal Cooperative Strategies for Active Aircraft Protection

Abstract: Aircraft protection from a missile employing a known linear guidance strategy is considered The aircraft protects itself by using a cooperating defensive missile Depending on the available cooperation scheme, three different cooperative linear quadratic guidance strategies are derived and investigated for arbitrary-order linear-adversaries dynamics: 1) two-way cooperation, where the target–defender team employs its optimal cooperative strategy, 2) one-way cooperation, realized by the defender employing a classical one-on-one guidance law while the target helps by luring in the missile, and 3) one-way cooperation, realized by the target employing an arbitrary evasive strategy while the defender attempts to reach the predicted interception point The performance of the three different cooperative strategies is compared and analyzed via simulation using the notion of Pareto fronts It is shown that the performance of the target and the defender is highly dependent on the cooperation scheme As expected, th

Partial Integrated Missile Guidance and Control with Finite Time Convergence

Abstract: A novel method of adaptive nonsingular terminal sliding mode control is proposed for a class of nonlinear systems subject to disturbances and uncertainties. The obtained results are applied to integrated guidance and control design for a missile. Using the Lyapunov stability theory, it is shown that the finite time convergence in both reaching and sliding phase is achieved by virtue of a terminal sliding surface and a switching controller that avoids singularity. With the assumption that the disturbances and uncertainties are bounded and the bounds are unknown, the adaptive method is incorporated with the controller design, which makes the proposed adaptive nonsingular terminal sliding mode control with better robustness. Due to the inherent property of time-scale separation between the translational and rotational dynamics, partial integrated guidance and control with a two-loop controller structure is designed for a missile based on the proposed adaptive nonsingular terminal sliding mode control. The ou...

Modified Command to Line-of-Sight Intercept Guidance for Aircraft Defense

Abstract: P ROTECTING an aircraft from missile attack is a challenging issue. Most missile guidance algorithms [1–11] use information basically about the missile and the target but not that of the protected system. Considering only the kinematics with the aid of the collision triangle, Boyell [12] calculated a minimum range to intercept the attacking missile over a specified distance to protect a moving aircraft/torpedo and obtained a closed-form expression of the range for the missile to be successful against attacking missiles. Shneydor [13] developed conditions for the applicability of Boyell’s technique and simplified the expression for the operational range for the defense missile. However, those analyses were based on an ideal guidance assuming a perfect collision course. Rusnak [14] applied the differential game theory (DGT) to such a three-player game, a protected aircraft, an attacking missile, and a defense missile, by changing the three-player game into a two-team game by way of grouping the same cooperative players: a protected aircraft and a defensivemissile. Rusnak [14] showed that the resulting expression of the guidance lawwas represented by two line-of-sight (LOS) rate terms multiplied by variable gains derived from the differential game theory. One of the two LOS rates is the LOS rate of the aircraft with respect to the attacking missile, and the other is the LOS rate of the attacking missile with respect to the defense missile. The Rusnak method demands the recursive backward computations from the end. Perelman et al. [15] further modified Rusnak’s method to represent the two LOS rate terms in a closed form instead of using the recursive solution. Both simulation studies showed that the required acceleration could be reduced as compared with that of the conventional two-player differential game theory. Further recent work on such DGT can be found in [16]. Shima [17] proposed an optimal guidance methodology for a defense missile assuming prior knowledge of an attacking missile’s guidance law. Shaferman and Shima [18] applied a multiple model adaptive estimator algorithm with the multiple model adaptive control for such a three-player game to estimate the attacking missile’s guidance law. The proposed work here takes a basic approach to the three-player guidance law, similar to the proportional navigation (PN), in the way that it uses geometrical information of the moving object, that is, the protected aircraft’s LOS and the LOS rate. This type of guidance law can be classified as a three-point guidance [19–21]; examples are the beam rider (BR) guidance [2,19–21] and the command to LOS (CLOS) guidance [2,19–21]. The “three points” in the three-point guidance usually means amissile, a target, and a reference point from which the target LOS is drawn or the target is observed. In this study, the protected aircraft is selected as one of the three points instead of the reference point. Ratnoo and Shima [22] applied the CLOS guidance for aircraft protection. Their analysis shows that the defense missile requires less lateral acceleration (latax) than that of the attacking missile. They also proposed a guidance law for the protected aircraft that would help the defense missile against the target. However, the problem is that as the range of the target becomes larger, the resolution of the target LOS angle at a tracking pointwill be lower. This fact degrades the system performance or demands a highresolution radar to track a moving target from a distance. This Note proposes a different approach as compared with the BR or the CLOS guidance though it falls under the three-point guidance classification. The proposed approach manipulates two LOS rates associated with three vehicles and has benefits over the conventional three-point guidance law in two major aspects: 1) a simple form with one gain for two LOS rates, and 2) high sensitivity to an attacking missile’smaneuvers in the proximity of the attackingmissile, but low sensitivity to the attacking missile’s maneuvers in the proximity of the protected aircraft. This novel guidance law, called the airborneCLOS guidance (A-CLOSG) law, is derived using optimal control theory.

Compound Control Methodology for Flight Vehicles

Abstract: Overview of Sliding Mode Control.- Overview of Active Disturbance Rejection Control.- Overview of Flight Vehicle Control.- The Descriptions of Flight Vehicle.- SMC for Missile Systems Based on Back-Stepping and ESO Techniques.- Adaptive SMC for Attitude Stabilization in Presence of Actuator Saturation.- Adaptive Nonsingular Terminal Sliding Mode Control for Rigid Spacecraft.- Attitude Tracking of Rigid Spacecraft with Uncertainties and Disturbances.- SMC for Attitude Tracking of Rigid Spacecraft with Disturbances.- Missile Guidance Law Based on ESO Techniques.- Missile Guidance Laws Based on SMC and FTC Techniques.- Cooperative Attack of Multiple Missiles Based on Optimal Guidance Law.

Composite guidance laws based on sliding mode control with impact angle constraint and autopilot lag

Abstract: For the terminal phase of tactical missiles intercepting maneuvering targets, the terminal guidance problem is studied. Based on an integral sliding mode (ISM) control method and nonlinear disturba...

Intercept Angle Missile Guidance Under Time Varying Acceleration Bounds

Abstract: A linear quadratic guidance law for a missile with a time varying acceleration constraint is presented. By introducing the constraint into the running cost, the optimization produces time varying gains that shape the missile’s trajectory for avoiding no-capture zones. The guidance law is derived for a missile with high-order autopilot dynamics and a terminal intercept angle constraint against a maneuvering target. The acceleration constraint of aerodynamic steering missiles is usually trajectory dependent rather than time dependent. Transforming the constraint into a time-dependent function by analytical means might not be possible, due to the nonlinear nature of the constraint. The problem is alleviated using a simple iterative calculation. For practical implementation reasons, and in order to improve the guidance performance under model uncertainties and disturbances, the guidance command is decomposed into two separate optimizations: one for the acceleration constraint, for which the guidance gains are...

Terminal Impact Angle Constrained Guidance Laws Using Variable Structure Systems Theory

Abstract: In this brief, variable structure systems theory based guidance laws, to intercept maneuvering targets at a desired impact angle, are presented. Choosing the missile's lateral acceleration (latax) to enforce sliding mode, which is the principal operating mode of variable structure systems, on a switching surface defined by the line-of-sight angle leads to a guidance law that allows the achievement of the desired terminal impact angle. As will be shown, this law does not ensure interception for all states of the missile and the target during the engagement. Hence, additional switching surfaces are designed and a switching logic is developed 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 nonlinear engagement dynamics for the general case of a maneuvering target.

A Guidance Law With Terminal Impact Angle Constraint Accounting for Missile Autopilot

Abstract: A sliding-mode guidance (SMG) law is designed to intercept maneuvering targets with impact angle constrained flight trajectories under the assumption of ideal missile autopilot. Furthermore, accounting for the autopilot as second-order dynamics, a new guidance law with terminal impact angle constraint is designed using the dynamic surface control method. Some first-order low-pass filters are introduced into the designing process to avoid the occurrence of high-order derivatives of the line of sight (LOS) angle in the expression of the guidance law such that the guidance law can be implemented in practical applications. The proposed guidance law is effective in compensating for the second-order autopilot lag. Simulation results show that it is able to guide a missile to impact a maneuvering target with a desired angle and a small miss distance.

Analytic Solutions of Generalized Impact-Angle-Control Guidance Law for First-Order Lag System

Abstract: To examine the effects of system lag on performance of a generalized impact-angle-control guidance law, analytic solutions of the guidance law for a first-order lag system are investigated. Under the assumptions of a stationary target and a first-order missile system with constant speed and small flight-path angle, the analytic solutions are obtained by solving a third-order linear time-varying ordinary differential equation. The solutions are expressed by combinations of polynomial, logarithmic, and infinite power series functions. The analytic solutions provide an insight into the behavior of the missile near the target: the guidance command, the acceleration of the missile, and the velocity component perpendicular to the collision course tend to diverge as the missile approaches the target. Terminal misses due to the system lag are discussed using the analytic solutions, and effects of guidance coefficients on the terminal misses are examined. Linear and nonlinear simulations are performed to verify th...

Adaptive back‐stepping sliding mode attitude control of missile systems

Abstract: SUMMARY This paper is devoted to the attitude control of a quaternion missile model, which is nonlinear in aerodynamics with atmospheric moment uncertainties, inertia uncertainties, bounded disturbances and actuator failures. By employing the back-stepping technique, the corresponding sliding mode controller is designed to guarantee the state variables of the closed loop system to converge to a small region of the reference states with the help of the adaptive law by estimating the total uncertainties, and be chatter-free. Also, simulation results are presented to illustrate the effectiveness of the control strategy. Copyright © 2013 John Wiley & Sons, Ltd.

Linear Quadratic Integrated Versus Separated Autopilot-Guidance Design

Abstract: Two types of guidance and control design concepts are explored: a traditional separated approach and an integrated one. Using the integrated approach, two different guidance systems are presented: a single-loop guidance law and a two-loop autopilot-guidance law. In the two-loop case, the autopilot loop is designed separately from the guidance one, but all the states are fed back into the guidance loop. It is proven that the integrated guidance laws achieve the same performance for single-input single-output systems. The performance of the three guidance laws is evaluated and compared via a thrust vector control missile. It is shown that the performance of the separated guidance law is inferior to that of the integrated laws.

Blazing Gyros: The Evolution of Strapdown Inertial Navigation Technology for Aircraft

Abstract: I NERTIAL navigation is the process of autonomously calculating the position and velocity of a moving vehicle frommeasurements of angular rotation and linear acceleration provided by vehiclemounted inertial sensors (gyros and accelerometers). The first inertial navigation system (INS) was developed at the Massachusetts Institute of Technology (MIT) Instrumentation Laboratory (eventually becoming the Charles Stark Draper Laboratory) for ballistic missile guidance [1]. (The INS includes velocity, attitude, heading, etc. outputs. In commercial application parlance, it also includes guidance steering outputs based on an input waypointdefined flight profile.) Soon thereafter, the technology was applied to aircraft navigation, with four companies eventually dominating the U.S. aircraft inertial navigation industry in the 1960s: Honeywell Aerospace and Defense Group with gyro design and manufacturing in Minneapolis, Minnesota, and INS design, development, and manufacturing in Clearwater, Florida; Kearfott in Wayne, New Jersey; Litton Guidance and Control Division in Woodland Hills, California; and Delco Electronics Division of General Motors in Milwaukee, Wisconsin. Honeywell specialized in high-accuracy systems and introduced a new electrostatically suspended gyro (ESG) technology for precision applications. Delco concentrated on transoceanic commercial and military cargo/tanker aircraft applications using the Carousel IV system (a variation of the Titan II ballistic missile inertial guidance set). Litton and Kearfott focused on medium-accuracy military tactical aircraft and airborne missile applications. To achieve required gyro accuracy, each of the aforementioned inertial navigation systems was configured with gimbaled platforms to isolate the inertial sensors from aircraft angular rates. INS advanced development at Litton and Kearfott in the 1960s centered on improving accuracy, reducing size, weight, and cost, and improving reliability of gimbaled INS products (important requirements for expanding military aircraft and airborne missile applications). A key contribution was the introduction of dry tuned-rotor-gyro (TRG) technology. Delco focused on reliability improvement for transport applications. Honeywell focused on improved accuracy and reliability of ESG gimbaled systems. For future cost, reliability, and size reduction based in part on projected advances in computer technology, several companies (most prominently, Honeywell) focused a significant portion of company resources on a radically new strapdown approach to inertial navigation: replacing the gimbaled platform with a computerized analytical equivalent and mounting (“strapping down”) the inertial sensors directly to the user vehicle. Based on technical books, journals, internet archives, discussions with past colleagues, but mostly from direct experience and personal records, this paper describes the curious and sometimes convoluted path by which the Honeywell strapdown program eventually led to development of the ring laser gyro (RLG) strapdown INS and conversion from gimbaled to strapdown technology throughout the airborne inertial navigation industry. For technical background, the paper first discusses the concept of inertial navigation using gimbaled versus strapdown system

Optimal Guidance-to-Collision Law for an Accelerating Exoatmospheric Interceptor Missile

Abstract: A guidance-to-collision law is presented. The guidance law is derived, using the optimal control methodology, for an accelerating exoatmospheric interceptor. It is dependent on a unique zero-effort heading angle error and on a variable denoted as the zero-effort angle of attack. The guidance law enables an accelerating interceptor to fly toward the interception point along a straight line, after the initial heading error is nulled. In addition, the guidance-to-collision law enables interception of the target with a terminal body angle. Moreover, the trajectory imposed by this guidance law enables an accelerating interceptor to estimate the target’s acceleration by using bearings-only measurements. This is possible due to the imposed line-of-sight rotation, making the range between the two vehicles observable. Using a nonlinear two-dimensional simulation, the performance of the guidance law is analyzed, showing superior performance compared to that of classical proportional navigation.

Missile Design and System Engineering

Abstract: In his latest book, "Missile Design and Systems Engineering", Eugene Fleeman presents a comprehensive review of the missile design and systems engineering process pulling from his decades of experience in the development of missiles and their technologies. Aimed toward the needs of aerospace engineering students and professors, systems analysts and engineers, program managers, and others working in the areas of missile systems and missile technology development, the book provides readers with an understanding of missile design, missile technologies, launch platform integration, missile system measures of merit and the missile system development process. This book has been adapted from Fleeman's earlier title, "Tactical Missile Design, Second Edition" to include a greater emphasis on systems engineering. Topics discussed include: top components in the missile design and system engineering process; critical tradeoffs, methods and technologies in aerodynamic, propulsion, structure, seeker, warhead, fuzing, and subsystems sizing to meet flight performance and other requirements; launch platform - missile integration; robustness, lethality, guidance, navigation & control, accuracy, observables, survivability, reliability, and cost considerations; missile sizing examples for missile systems and missile technologies; and, missile system and missile technology development process.

Optimal feedback guidance for nonlinear missile model with impact time and angle constraints

Abstract: This study addresses the optimal guidance for nonlinear missile model which enables missile to attack the target at the designated time and angle while integral square control efforts are minimized. This guidance is an optimal control problem and can be reduced to a two point boundary value problem (TPBVP). The optimal input is given by a function of state variables and parameters determined by solving the TPBVP. This TPBVP can be easily solved by using one of shooting methods. In the case where there is no impact angle constraint, optimal input can be obtained as in the case where both impact time and angle are designated. The validity of the proposed feedback guidance is verified by simulations.

Ionospheric disturbances induced by a missile launched from North Korea on 12 December 2012

Abstract: [1] Ionospheric disturbances caused by a missile launched from North Korea on 12 December 2012 were investigated by using the GPS total electron content (TEC). The spatial characteristic of the front edge of V-shaped disturbances produced by missiles and rockets was first determined. Considering the launch direction and the height of estimated ionospheric points at which GPS radio signal pierces the ionosphere, the missile passed through the ionosphere at heights of 391, 425, and 435 km at 0056:30, 0057:00, and 0057:30 UT, respectively. The observed velocities of the missile were 2.8 and 3.2 km/s at that time, which was estimated from the traveling speed of the front edge of V-shaped disturbances. Westward and eastward V-shaped disturbances propagated at 1.8–2.6 km/s. The phase velocities of the westward and eastward V-shaped disturbances were much faster than the speed of acoustic waves reported in previous studies, suggesting that sources other than acoustic waves may have played an important role. Furthermore, the plasma density depletion that is often observed following missile and rocket launches was not found. This suggests that the depletion resulting from the missile's exhaust was not strong enough to be observed in the TEC distribution in the topside ionosphere.

An Aerodynamic Analysis of a Spinning Missile with Dithering Canards

Abstract: A generic spinning missile with dithering canards is used to demonstrate the utility of an overset structured grid approach for simulating the aerodynamics of rolling airframe missile systems. A grid convergence study andassessment of viscous effects shows that a medium-resolution viscous grid with 8 million grid points provides a good compromise between solution accuracy and throughput. Viscous effects should be included in detailed studies to capture the interaction between the inboard canard vortex and the fuselage and tail boundary layers. The computed results agree well with experimental data. A database of cases with variation of angle of attack and the strength of control authority of a pitch-up maneuver is computed to evaluate missile performance. The case throughput rates are sufficient to contemplate population of aerodynamic databases with Navier-Stokes computations. Overall, the structured overset grid approach enables accurate and efficient simulation of rolling airframe missile configurations that involve relative motion between system components.

Application of partial sliding mode in guidance problem

Abstract: In this paper, the problem of 3-dimensional guidance law design is considered and a new guidance law based on partial sliding mode technique is presented. The approach is based on the classification of the state variables within the guidance system dynamics with respect to their required stabilization properties. In the proposed law by using a partial sliding mode technique, only trajectories of a part of states variables are forced to reach the partial sliding surfaces and slide on them. The resulting guidance law enables the missile to intercept highly maneuvering targets within a finite interception time. Effectiveness of the proposed guidance law is demonstrated through analysis and simulations.

A Hybrid Guidance Law for a Strapdown Seeker to Maintain Lock-on Conditions against High Speed Targets

Abstract: This paper proposes a new guidance law, which considers the Field of View (FOV) of the seeker when a missile has a strapdown seeker mounted instead of a gimbal seeker. When a strapdown seeker, which has a narrow FOV, is used for tracking a target, the FOV of the seeker is an important consideration for guidance performance metrics such as miss distance. We propose a new guidance law called hybrid guidance (HG) to address the shortcomings of conventional guidance laws such as proportional navigation guidance (PNG), which cannot maintain lock-on conditions against high speed targets due to the narrow FOV of the strapdown seeker. The aim of the HG law is to null miss distance and to maintain the look angle within the FOV of the strapdown seeker. In order to achieve this goal, we combine two guidance laws in the HG law. One is a PNG law to null the LOS rate, and the other is a sliding mode guidance (SMG) law derived to keep the look angle within the FOV by employing a Lyapunov-like function with a sliding mode control methodology. We also propose a method to switch these two guidance laws at certain look angles for better guidance performance.

Hierarchical closed-loop flow control system for aircraft, missiles and munitions

Abstract: The present invention relates to a missile or aircraft with a hierarchical, modular, closed-loop flow control system and more particularly to aircraft or missile with a flow control system for enhanced aerodynamic control, maneuverability and stabilization. The present invention further relates to a method of operating the flow control system. Various embodiments of the flow control system of the present invention involve different elements including flow sensors, active flow control device or activatable flow effectors and logic devices with closed loop control architecture. The sensors of these various embodiments are used to estimate or determine flow conditions on the various surfaces of a missile or aircraft. The active flow control device or activatable flow effectors of these various embodiments create on-demand flow disturbances, preferably micro-disturbances, at different points along the various aerodynamic surfaces of the missile or aircraft to achieve a desired stabilization or maneuverability effect. The logic devices are embedded with a hierarchical control structure allowing for rapid, real-time control at the flow surface.