Showing papers on "Missile published in 1996"

Practical guidance for homing missiles with bearings-only measurements

Abstract: Observability is analyzed for the target tracking with bearings-only measurements (BOM) applied to a passive homing missile system using augmented proportional navigation guidance (APNG). Based on the analysis, a practical homing guidance law is proposed to enhance observability for tactical missile applications. The new guidance law is tailored to provide oscillation of line-of-sight (LOS) angle induced by initial missile heading errors without sacrificing terminal effectiveness. Simulation studies indicate that the proposed guidance law provides convergence of the state estimates essential to homing guidance applications as well as terminal effectiveness.

Moving mass trim control for aerospace vehicles

Abstract: A moving mass trim controller is proposed to increase the accuracy of axisymmetric, ballistic vehicles. The moving mass trim controller differs from other moving mass schemes because it generates an angle of attack directly from the mass motion. The nonlinear equations of motion for a ballistic vehicle with one moving point mass are derived and provide the basis for a detailed simulation model. The nonlinear equations are linearized to produce a set of linear, time-varying autopilot equations. These autopilot equations are analyzed and used to develop theoretical design tools for the creation of moving mass trim controllers for both fast and slow spinning vehicles. A fast spinning moving mass trim controller is designed for a generic artillery rocket that uses principal axis misalignment to generate a trim angle of attack. A slow spinning moving mass trim controller is designed for a generic re-entry vehicle that generates a trim angle of attack with a center of mass offset and aerodynamic drag. The performance of both moving mass trim controllers are evaluated with the detailed simulation. VER the years, techniques for controlling the flight character- istics of missiles and re-entry vehicles (RV) have gravitated to systems that deliver relatively large amounts of control authority. For certain missions, such as an air-to-air missile or an RV designed to evade defenses, a large lateral acceleration capability was required. The technologies used to perform these missions ranged from ac- tuated canards, elevons, and flaps to jet interaction, thrust vector control, and a variety of other techniques.1 Because of the mission requirements for large maneuvers, systems that provided modest amounts of control capability were of little or no value. However, a new mission for accurate artillery rockets and RVs that utilize existing assets has prompted a renewed interest in simpler control techniques that produce small maneuvers. One such control technology is moving mass control. This tech- nique has previously been evaluated in conjunction with other con- trol methods such as the moving mass roll control of an aerodynam- ically asymmetric RV.2'3 A more direct application of moving mass control technology is the moving mass trim controller (MMTC). The MMTC generates a trim angle of attack (AOA) on an axisym- metric, ballistic vehicle directly from the motion of the mass. It is a novel, lightweight, low-cost retrofit to spinning ballistic vehi- cles that require modest flight-path corrections to obtain increased accuracy. Over 10 years ago, initial studies of the MMTC were performed by Regan and Kavetsky4 at the U.S. Naval Systems Warfare Cen- ter. Regan and his co-workers devised a single-shot MMTC that would provide modest range corrections near the target. At Sandia National Laboratories (SNL), the MMTC was an outgrowth of the deconing device test (DOT) described by White and Robinett.5 The DDT provided an initial glimpse of the effects of principal axis mis- alignment (PAM), roll rate, and center of mass offset. The MMTCs developed at SNL address the issue of roll rate, static margin (SM), PAM, and center of mass offset. The trim AOA for a fast spinning vehicle is generated by a PAM, whereas a slow spinning vehicle with a small SM relies on a center of mass offset to create a trim AOA resulting from aerodynamic drag. This paper derives the gen- eral nonlinear equations of motion for a one-moving mass system,

Gain-scheduled bank-to-turn autopilot design using linear parameter varying transformations

Abstract: A gain-scheduled autopilot design for a bank-to-turn missile is presented. The approach follows previous work for a longitudinal missile autopilot. The method is novel in that the gain-scheduled design does not involve linearizations about operating points. Instead, the missile dynamics are brought to a linear parameter varying form via a state transformation. A linear parameter varying system is defined as a linear system whose dynamics depend on an a priori unknown but measurable exogenous parameter. This framework is applied to the design of a coupled longitudinal/lateral bank-to-turn missile autopilot. The pitch and yaw/roll dynamics are separately transformed to linear parameter varying form, where the cross axis states are treated as exogenous parameters. These are actually endogenous variables, and so such a plant is called quasilinear parameter varying. Once in quasilinear parameter varying form, a family of robust controllers using p, synthesis is designed for both the pitch and yaw/roll channels, using angle of attack and roll rate as the scheduling variables. The closed-loop time response is simulated using the original nonlinear model and also using perturbed aerodynamic coefficients.

Root-mean-square miss distance of proportional navigation missile against sinusoidal target

Abstract: The performance of a proportional navigation homing missile against a target performing a sinusoidal weave maneuver is evaluated. The missile's effectiveness is measured in terms of the root-mean-square miss distance over a set of engagements in which the initial phase of the target weave is uniformly distributed. Closed form solutions for the root-mean-square miss are derived for the case where the missile guidance system is modeled by a first-order lag and the lateral acceleration is unlimited. The analysis is then extended to include the effects of acceleration saturation and higher order missile dynamics. Comparisons are made between a first-order and a fifth-order guidance system, and the root-mean-square miss is determined numerically as a function of the interceptor's effective navigation gain, time constant and acceleration limit, and the target's weave amplitude and frequency. UTURE homing interceptor missiles will face new and unique challenges as the sophistication of the threat spectrum in- creases. Engagements against air targets can occur at both very low and very high altitudes, with the threats accidentally or intentionally performing weaving or spiraling maneuvers during their midcourse and terminal phases.1"4 The lateral displacement, acceleration ca- pability, and weave frequency of the target maneuver can greatly enhance the threat's ability to survive a counterattack. To counter this, the defensive missile must have sufficient lateral acceleration, guidance system time constant, and terminal homing time to achieve a high probability of intercept. Whereas the per- formance of the interceptor will be influenced by many scenario- dependent factors,5"8 a major consideration will be the fundamental response of the proportional navigation (PN) guidance system to the postulated target weave motion.9'10 In the general case, the target dynamics may involve arbitrary periodic motion in three dimensions. A useful starting point for analysis, however, is the response of the PN homing system to a single plane sinusoidal maneuver of constant amplitude and fre- quency. The phase angle of target weave, which is associated with initial conditions at the start of the missile's terminal guidance, can be treated as a random variable, uniformly distributed between 0 and 2n over a set of engagements. The missile's dynamics are approxi- mated by a simple first-order transfer function, and unlimited lateral acceleration capability is assumed. The miss distance can then be parameterized in terms of the effective PN navigation gain N, the missile time constant T, and the amplitude AT and frequency co of the target weave. This paper focuses on root-mean-square (rms) miss distance as a recommended measure of effectiveness in analyzing missile perfor- mance against weaving targets. This measure allows uncertainties in target phase characteristics to be accounted for in the terminal per- formance results. The weaving target problem was first addressed by Chadwick, 9 who determined analytical expressions for the rms miss distance of the single lag PN missile for values of N =2 and 3. Zarchan 10 employed adjoint theory and transfer function tech- niques to determine formulas for the peak miss distance against a weaving target for values of N between 3 and 6. The present paper derives general closed-form expressions for the rms miss distance against a sinusoidal target. New results are obtained for arbitrary

An efficient scheduling algorithm for a multifunction radar

Abstract: The new generation of air defence frigates has to be equipped with a high performance sensor suite to cope with the threat in future maritime scenarios. A multifunction radar (MFR) with a single rotating or multiple fixed phased array antennas is often the key element in this sensor suite because it can perform not only surveillance functions but also missile support. Since the MFR must be able to execute multiple functions concurrently, a scheduling mechanism is required that allocates the time and energy resources of the MFR to the radar functions in such a way that the overall performance of the sensor suite is optimized. This paper describes an efficient scheduling algorithm that is able to schedule the dwell requests of each function on the basis of their priority and transmission window. Some results are shown of the scheduling algorithm in a typical scenario in which a number of seaskimmers is engaged by semi-active surface-to-air missiles.

Design of Nonlinear Autopilots for High Angle of Attack Missiles

Abstract: Two nonlinear autopilot design approaches for a tail-controlled high angle of attack air-to-air missile are described. The research employs a highly nonlinear, time varying pitch plane rigid- body dynamical model of a short range missile. Feedback linearization technique together with linear control theory are then used for autopilot design. In order to manage the difficulties associated with "zerodynamics" that arise in tail controlled missiles, two distinct approaches for approximate feedback linearization are advanced. The first approach imposes a time-scale structure in the closed-loop dynamics, while the second technique redefines the output. Performance of these autopilots are illustrated in a nonlinear simulation.

Missile flight control using active flexspar actuators

Abstract: A new type of subsonic missile flight control surface using piezoelectric flexspar actuators is presented. The flexspar design uses an aerodynamic shell which is pivoted at the quarter-chord about a graphite main spar. The shell is pitched up and down by a piezoelectric bender element which is rigidly attached to a base mount and allowed to rotate freely at the tip. The element curvature, shell pitch deflection and torsional stiffness are modeled using laminated plate theory. A one-third scale TOW 2B missile model was used as a demonstration platform. A static wing of the missile was replaced with an active flexspar wing. The 1 in 2.7 in active flight control surface was powered by a bimorph bender with 5 mil PZT-5H sheets. Bench and wind tunnel testing showed good correlation between theory and experiment and static pitch deflections in excess of . A natural frequency of 78.5 rad with a break frequency of 157 rad was measured. Wind tunnel tests revealed no flutter or divergence tendencies. Maximum changes in lift coefficient were measured at which indicates that terminal and initial missile load factors may be increased by approximately 3.1 and 12.6 g respectively, leading to a greatly reduced turn radius of only 2400 ft.

Missile jet vane control system and method

Abstract: A jet vane control system and method for a missile in which the system is compact, rugged, lightweight and detachably connected to the aft end of a missile. The system provides for very quick pitch over and roll control during launch. The system then detaches from the missile so as not to burden the missile during its flight to target. The vanes of the system are divided into quadrants, each having its own mounting support and gear train assembly. Each vane is also connected through a detachable coupling to the steering control system of the missile such that actuation of the steering control system simultaneously actuates the jet vane control system.

The design, development and testing of the THAAD (Theater High Altitude Area Defense) solid state phased array (formerly ground based radar)

Abstract: This paper describes the Theater High Altitude Area Defense (THAAD) solid state phased arrays design, development and testing. It is the largest most complex solid state phased array built in the world, employing more than 25,000 active radiators each with a corresponding gallium arsenide (GaAs) transmit/receive (T/R) module. The array operates in several aperture illumination modes, is self calibrating in the field, transportable and designed to be "fail soft". The array's performance has been verified on a near-field range and during radar search and track missions of satellites and missiles at White Sands Missile Range (WSMR).

Short-time stability of proportional navigation guidance loop

Abstract: Stability characteristics of proportional navigation (PN) guidance are analyzed by using the short-time stability criterion which is extended here to accommodate time-varying state weights and time-varying bounds of the state norm. As short-time stability is defined over a specified time interval, its application to the stability analysis of a homing guidance loop that operates up to a finite time gives more accurate results than previous studies. Furthermore, within the framework of short-time stability, zero effort miss and acceleration command, which are the most important variables determining guidance performance, can be directly related with guidance loop stability. An application to a PN guidance loop with a 1st-order missile/autopilot time lag shows that the stability condition based on short-time stability is less conservative than the previous results based on hyperstability and Popov stability.

Optimal Planar Evasive Aircraft Maneuvers Against Proportional Navigation Missiles

Abstract: A comparative study of evasive strategies of an aircraft against a missile with fixed, gravity-limited, proportional navigation is conducted for both subsonic and supersonic confrontations. A complete point-mass aircraft model and a variable-mass missile model that includes missile dynamics are used. No linearization is required in the analysis, and all motion is constrained to a horizontal plane. Sequential quadratic programming is used to solve the optimal control problem. Numerical results are presented for an early model of the F-4 fighter aircraft. In particular, the effects of varying the aircraft/missile initial velocity ratio, the missile initial heading angle, and the missile guidance time constant are determined.

High performance supersonic missile inlet design using automated optimization

Abstract: A multilevel design strategy for supersonic missile inlet design is developed. The multilevel design strategy combines an efŽ cient simple physical model analysis tool and a sophisticated computational  uid dynamics (CFD) Navier – Stokes analysis tool. The efŽ cient simple analysis tool is incorporated into the optimization loop, and the sophisticated CFD analysis tool is used to verify, select, and Ž lter the Ž nal design. The genetic algorithms and multistart gradient line search optimizers are used to search the nonsmooth design space. A geometry model for the supersonic missile inlet is developed. A supersonic missile inlet that starts at Mach 2.6 and cruises at Mach 4 was designed. SigniŽ cant improvement of the inlet total pressure recovery has been obtained. Detailed  owŽ eld analysis is also presented.

Ballistic Missile Proliferation: The Politics and Technics

Abstract: The current concern that ballistic missile technologies are spreading throughout the world is addressed in this book It examines the missile and missile-armament programmes and technologies, and the ability of countries to acquire such technologies The concluding chapter investigates the international efforts to control ballistic missile proliferation

Analysis of optimal midcourse guidance law

Abstract: An optimal midcourse guidance law is presented that maximizes the final speed for missiles against a target at far distance or at low attitude in which the final speed is a prime factor. An explicit acceleration command is derived analytically in which the trajectory-dependent optimal control gains are written in terns of thrust, lift, drag, and intercept boundary condition. The optimal guidance law can be implemented either in airframe coordinates or inertia coordinates. It is shown that the acceleration commands with constant control gain are adequate when the range is relatively short; during midcourse guidance, however, the optimal control gains are required to enhance the performance.

Nonlinear H-infinity optimal control for agile missiles

Abstract: Nonlinear #00 optimal control is used to design a pitch plane flight control system for a high-angle-of-attack agile missile. The nonlinear H^ control law is obtained by approximating the solution to the Hamilton-JacobiIsaacs equation. The solution to the partial differential equation is formed using the method of characteristics and is numerically approximated using successive approximations. Nonlinear simulation results using the nonlinear control law are presented. Simulation results testing the algorithms demonstrated excellent performance.

Missile fuzing system

Abstract: A fuzing system adapted for use by a guided missile to generate a detonation signal for a warhead carried by the missile. The missile has a seeker/tracker to track, and direct the missile towards, a target. The seeker/tracker has: a seeker, gimballed with respect to the body of the missile, for producing a signal representative of the angular deviation between the target and the missile; and, a ranging system for producing a signal representative of a range between the target and the missile. The fuzing system, in response to the range signals, produces a time-to-go signal, t go , where t go , related to the range between the missile and the target divided by the rate of change in such range. The seeker produced signal represents the line of sight angle between the missile and the target and wherein the processor, in response to the range signal and the line of sight angle signal produces a miss distance signal representative of a predicted distance, normal to the line of sight, at the time the remaining before the missile intercepts the target. A fragment velocity signal is produced representative of the velocity of fragments of the warhead divided by the predicted warhead miss distance.

Adaptive matched augmented proportional navigation

Abstract: Two new adaptive homing guidance laws are developed which are variants of Augmented Proportional Navigation (PRONAV). The first guidance law commands flight path angle rate and is referred to as adaptive matched augmented PRONAV (AMAAP). The second guidance law commands linear acceleration and is referred to as adaptive matched augmented linear PRONAV (AMALP). The major attributes of these guidance laws are (1) they are the solutions of a linear quadratic control problem, (2) they do not require an estimate of time-to-go (tgo), (3) they are matched to a nonlinear model of the target's motion, (4) they adapt in real time to provide optimal guidance over each small segment of the intercept trajectory, and (5) they optimally account for missile deceleration.

Passive aircraft and missile detection device

Abstract: This invention offers the capability to passively detect and classify missiles, aircraft, and units with gas turbine engines and rocket motors. It does so by measuring the electromagnetic radiation from ions in the exhaust gas. These emissions are low in frequency and propagate beyond the visual horizon with little attenuation other than through wavefront spreading in the earth-ionosphere waveguide. Acoustic and mechanical processes in combustion chambers and turbines modulate the ion signals in a deterministic manner that makes the engine type classifiable. The proposed passive detection and classification device can sense and identify aircraft, missiles, and potentially ground vehicles and gunfire. This capability will reduce vulnerability of forces to threats and reduce fratricide among friendly units.

The Legality of the 1993 US Missile Strike on Iraq and the Right of Self-Defence in International Law

Abstract: In the early evening hours of Saturday, 26 June 1993, the United States launched a missile attack on Iraq. Twenty-three Tomahawk sea-to-ground missiles were fired from two US warships, the USS Chancellorsville and the USS Peterson, located in the Persian Gulf and the Red Sea respectively.1 Sixteen of those launched hit their desired military target, the Military Intelligence Headquarters, situated just outside the Iraqi capital of Baghdad. A further four missiles fell within the compound of the intelligence service complex. Conflicting reports put the death toll at between six and eight civilians, with 20 injured, when the remaining three missile warheads went astray.2 The Venezuelan Embassy was also reported to have been damaged.3

Apparatus for unfolding and fixing missile fins

Abstract: An improved apparatus for unfolding and fixing missile fins which is capable of automatically unfolding the fins of a missile when launching the missile loaded in a missile launch tube, in which the fins of the missile are folded, which includes a plurality of fins fixed to a missile body, a plurality of rotation fins rotatably supported by the fixed fin, rotation stoppers elastically supported in the direction the rotation fins are unfolded, and unfolding and fixing member forwardly and rearwardly movable with respect to the fixed fins and each having a straight movement stopper elastically supported toward the rotation stopper.

Neural-adaptive nonlinear autopilot design for an agile anti-air missile

Abstract: Previous research has shown that neural networks can be used to improve upon approximate dynamic inversion controllers in the case of uncertain nonlinear systems. In the one possible architecture, the neural network adaptively cancels linearization errors through on-line learning. Learning may be accomplished by a simple weight update rule derived from Liapunov theory, thus assuring the stability of the closed-loop system. In this paper, the authors apply this methodology to design a bank-to-turn autopilot for an agile anti-air missile. First, a control scheme based on approximate inversion of the vehicle dynamics is presented. This nonlinear control system is then augmented by the addition of a feedforward neural network with on-line learning. Finally, the resulting control law is demonstrated in a nonlinear simulation, and its performance is evaluated relative to a more traditional gain-scheduled linear autopilot. (Author)

Residual advantage learning applied to a differential game

Abstract: An application of reinforcement learning to a differential game is presented. The reinforcement learning system uses a recently developed algorithm, the residual form of advantage learning. The game is a Markov decision process (MDP) with continuous states and nonlinear dynamics. The game consists of two players, a missile and a plane; the missile pursues the plane and the plane evades the missile. On each time step each player chooses one of two possible actions; turn left or rum right 90 degrees. Reinforcement is given only when the missile hits the plane or the plane reaches an escape distance from the missile. The advantage function is stored in a single-hidden-layer sigmoidal network. The reinforcement learning algorithm for optimal control is modified for differential games in order to find the minimax point, rather than the maximum. As far as we know, this is the first time that a reinforcement learning algorithm with guaranteed convergence for general function approximation systems has been demonstrated to work with a general neural network.

Method of computing divert velocity for the ground-based interceptor using numerical partial derivatives

Abstract: A ground-based interceptor missile, part of a larger ballistic missile defense system, is provided for intercepting enemy intercontinental ballistic missiles The interceptor employs a method based on numerical partial derivatives for computing required divert velocity corrections to remove predicted ground-based interceptor miss errors for engagement during the midcourse of the enemy missile flight

Angle only range estimator for homing missile

Abstract: Missile motion information from an inertial measurement set and the angle-to-target information from a seeker/sensor is used to estimate range and/or time to go on a homing missile without a direct measurement of range or range rate. Up to six comparisons of spectrally matched linear and angular measurements, corrected by target state estimates and analytical coupling terms, are used to determine the error in the range estimate. A nonlinear sorting and filtering procedure rejects the unusable error sources among the six potential error sources, and thresholds against too little information. The stability margin in the missile guidance loop is used as a sensor indicator for excessive estimated range.

Performance enhancement methods for unstable bank-to-turn (BTT) missiles with saturating actuators

Abstract: In this paper, three methods are presented to enhance the performance of an unstable missile system in the presence of multiple saturating actuators. Each method involves the design of a saturation detection system, which modifies a nominal autopilot via online optimization and maintains, to the extent possible, the multivariable properties of the nominal design. To prevent saturation due to exogenous signals, disturbances and noises, in particular, an error governor is proposed. This system processes the error signals within the feedback loop and the state of the compensator, so that appropriately co-ordinated control surface deflections may be generated. Traditionally, such a system has been applied only to stable plants. In this paper, it is shown how to design such a system for unstable plants. To prevent saturation due to reference (acceleration) commands from the guidance system, a reference governor is proposed. Such a system has traditionally required access to the entire state of the plant as wel...

Missile launch and flyout simulator

Abstract: A Missile Launch and Flyout Simulator (10) for simulating the UV and IR flight characteristics of an incoming missile throughout its launch, powered flight and post burnout phases, as would be viewed by a missile launch detection and tracking system. The simulator (10) produces a UV output to simulate the launch of a missile, and an IR output to simulate the powered flight and post burnout phases of the missile's flight. The simulator (10) is also portable and capable of being remotely triggered so that it can be used in isolated locations or on moving platforms.

Advanced guidance laws for acceleration-constrained missile, randomly maneuvering target and noisy measurements

Abstract: Explicit, closed-form formulae of advanced guidance laws for a linear, time-invariant, acceleration-constrained arbitrary-order missile, and a linear, time-invariant, arbitrary-order, randomly maneuvering target with noisy position measurements are derived. Two approaches are presented. The first approach derives the optimal guidance law for a quadratic objective. The solution is the guidance law for deterministic system with limiting on the commanded acceleration applied on the estimated state. The limiting function in this case is the saturation function. The second approach derives a control law called the average input guidance law. This approach is based on the idea of applying the average of the input that would have been applied to the plant if the noises were known. The solution has similar structure. It is the guidance law for deterministic system with limiting on the commanded acceleration applied on the estimated state. The limiting function in this case is from the family of describing functions of the saturation function. The formulas of the different guidance laws are given in terms of the transfer function of the missile and acceleration constraint, the shaping filter of the maneuver of the target, responses to initial conditions, error variance matrix of the estimated state and weights in the criterion. It is demonstrated by simulations that although the optimal guidance law has improved performance in terms of the miss distance, the suboptimal average input guidance law consumes less energy.