Divert Thrusters Research Articles

Integrated Controller for Divert Thruster Flight Propulsion Control Systems

Future generation missile calls for development of fast response propulsion based control systems to achieve hit-to-kill miss distance against highly manoeuvrable targets covering high-endo to exo-atmosphere regions. Control systems based on solid propulsion are developed connecting multiple pintle nozzles and control thrust vector is generated with distribution hot gasses in pitch/yaw directions. Differential distribution of hot gasses among multiple nozzles is complex due to time varying coupled non-linear chamber pressure dynamics with propellant combustion and nozzles’ flow phenomena and methods based on linear control may not ensure stability in wide operating conditions. There is need to develop "Unified Controller" based on non-linear formulation to ensure control stability in wide operating conditions of various mission applications, with advanced features for new generation missiles to counter various futuristic targets, and achieve hit-to-kill capabilities following integrated approach in order to improve the total system performance. Non-linear algorithms proposed in the recent literature do not exploit synergism between plant and actuation systems dynamics. Present paper presents novel non-linear feedback linearization technique, Dynamic Inversion, a relatively straightforward approach in Integrated Control Framework (ICF) implementing inherent time scale separation between solid motor plant and hot gas flow control valve actuation system. To ensure stability in various operating conditions and assure robust performance in the presence of parameter inaccuracies, controller is augmented with model-following neuro-adaptive design. Prototype systems: Thrust Controlled Motor (TCM) and Divert Thruster Flight Propulsion Control Systems (DTFPCS) are developed to demostrate/validate "Unified Controller" for missile control system. Closed loop control of chamber pressure is demonstrated with single pintle nozzle system and control thrust vector generation with multiple nozzle systems. Model update takes place adaptively online and unlike a typical indirect adaptive control approach, there is no need to update the individual parameters explicitly. Instead the inaccuracy in the system output dynamics is captured directly and then used in modifying the control. Based on present work, missilecontrol systems for various applications can be developed with limited controller changes and thereby minimizing number of developmental tests and cost.

Flow Field Simulation Tool of Solid Propulsion Control Systems for Missile Applications

In order to achieve hit-to-kill miss distances against manoeuvrable ballistic missile targets, future generation missiles calls for fast response non-conventional propulsion control systems during terminal homing period in high-endo to exo-atmosphere regions augmenting aero control. Based on advances in thrust modulation controlling throat area in solid propulsion pintle motor, control systems are developed for missile Divert and Attitude Control Systems (DACS) applications using multiple nozzles in pitch/yaw directions connected to common gas generator. Control thrust vector for lateral and attitude corrections is generated with proportional distribution of total thrust in pitch/yaw directions through real time thrust modulation of pintle nozzles
 Towards development of propulsion control systems based on solid propulsion, prototype systems: Thrust Controlled Motor (TCM) and Divert Thruster Flight Propulsion Control System (DTFPCS) are developed to demonstrate thrust modulation/ control thrust vector generation in solid propulsion and finalize unified controller for various missile control applications. Design & development process of prototype systems calls for flow field simulation tool for parametric design analysis/optimization and generate performance maps/control parameters based on detailed flow distributions. Present paper presents flow field simulation tools with Steady Flow Field (SFF) model and Dynamic Flow Field (DFF) model integrating dynamics of pintle movement with steady flow solver. Parametric analysis of geometrical profiles of pintle/throat regions and effect of pintle speed on system response are presented for system design optimization. Performance maps and controller maps of prototype systems are generated based on flow field distributions for nozzle thrust modulation. Simulated results are compared with open loop hot tests and unified controller parameters are updated for control thrust vector generation. Finally flow field distributions of DTFPCS operating modes: neutral-thrust and lateral-thrust are presented demonstrating the utility of simulation tool in design analysis and system development. Present simulation tool is critical in system design optimization and unified controller development for various flight control applications with minimum number of developmental tests. Further work of integrating propellant grain geometry variation during motor hot tests, with flow field model is in progress to predict control system performance for various missile control applications.

Development of high burn rate propellant and testing in miniature rocket motor for control applications

For aerospace vehicles, notably missiles or satellites for attitude and divert control, impulsive thrusters are most desirable if they are quantized as exact instantaneous pulses available over and after extended periods of inactive storage. These thrusters operate for a few milliseconds (ms), typically between 10 ms to 100 ms duration. Compared to its liquid or gaseous counterpart, an impulsive thruster powered by solid propellants is a simple, and reliable solution. Therefore, they are suitable as reaction control systems and divert thrusters for satellite and unmanned aerial vehicle applications.A typical solid propellant would need an extremely thin web thickness for port burning arrangement due to its brief burn duration, which makes propellant manufacturing and its structural anchoring in motor challenging. A high burn rate and high-density propellant based on ammonium perchlorate, aluminum, and Teflon is processed to enable end-burning configuration thus solving the problems, simplifying propellant manufacturing, structural design and integration. It also permits stacking propellants for multi-pulse applications. The high burn rate (40 mm/s at 120 bar pressure) propellant is formulated and tested for ballistic performance evaluation in proof motor configuration, and evaluation of its advantages over conventional propellant. A case of identical design requirements with both double base propellant and high burn rate pressed propellant is compared, tested and evaluated to compare the relative performance merits. The volumetric load is increased from 47% to 82% due to the development and adoption of high burn rate propellant. The overall weight of the thruster is also reduced by 20%.

Reinforcement learning for angle-only intercept guidance of maneuvering targets

We present a novel guidance law that uses observations consisting solely of seeker line-of-sight angle measurements and their rate of change. The policy is optimized using reinforcement meta-learning and demonstrated in a simulated terminal phase of a mid-course exo-atmospheric interception. Importantly, the guidance law does not require range estimation, making it particularly suitable for passive seekers. The optimized policy maps stabilized seeker line-of-sight angles and their rate of change directly to commanded thrust for the missile's divert thrusters. Optimization with reinforcement meta-learning allows the optimized policy to adapt to target acceleration, and we demonstrate that the policy performs better than augmented zero-effort miss guidance with perfect target acceleration knowledge. The optimized policy is computationally efficient and requires minimal memory, and should be compatible with today's flight processors.

Integrated Estimation/Guidance Law against Exoatmospheric Maneuvering Targets

An integrated guidance integrated estimation/guidance law is designed for exoatmospheric interceptors equipped with divert thrusters and optical seekers to intercept maneuvering targets. This paper considers an angles-only guidance problem for exoatmospheric maneuvering targets. A bounded differential game-based guidance law is derived against maneuvering targets using zero-effort-miss (ZEM). Estimators based the extended Kalman filter (EKF) and the unscented Kalman filter (UKF) are designed to estimate LOS rates that are contaminated by noise and target maneuver. Furthermore, to improve the observability of the range, an observability enhancement differential game guidance law is derived. The guidance law and the estimator are integrated together in the guidance loop. The proposed integrated estimation/guidance law has been tested in several three-dimensional nonlinear interception scenarios. Numerical simulations on a set of Monte-Carlo simulations prove the validity and superiority of the proposed guidance law in hit-to-kill interception.

Numerical Prediction of Surface Heat Flux During Multiple Jets Firing for Missile Control

Numerical simulations are carried out to obtain the flowfield and heat flux arising out of the flow interactions of different control thrusters viz. vernier, pitch, yaw, roll and divert thruster among themselves as well as with free stream at different altitudes of operation. Three critical points on a typical trajectory of a missile are chosen and combinations of the thrusters operating at those conditions are considered. Simulations have also been performed to simulate DT motor interaction with free stream at different altitude. The interaction of different motor flowfield with the free stream presents a very complex flowfield. Flow gradients are very high close to the nozzle exit because of high altitude operation of motors. 3-Dimensional RANS equations are solved along with k-e turbulence model on unstructured tetrahedral grid using commercial CFD software. The flow properties along with the surface heat flux distribution for four isothermal wall temperatures 350, 450, 550, 650 K are computed and provided for surface temperature prediction. It is observed that with increase in altitude, the high heat flux region reduces and heat transfer coefficient is independent of wall temperature.

Logic-Based Guidance Law for Interceptor Missiles Steered by Aerodynamic Lift and Divert Thruster

A low-altitude endoatmospheric interceptor, which is steered by aerodynamic lift and divert thruster, is investigated in this paper. The divert thrust is provided by a solid rocket engine with a single public combustion chamber and a group of nozzles around the missile. As a result, it can be used only once. To utilize the divert thruster effectively and to determine the width of the divert thruster pulse, a logic-based guidance strategy is proposed. Moreover, the decomposition of the game space is employed to analyze the appropriate initiating time of the divert thruster. Finally, the validity of the guidance strategy is tested using a realistic 2-D ballistic missile defense scenario. It is shown that, due to the sufficient utilization of the divert thruster, there is a significant improvement in the homing performance.

Collision-Geometry-Based Pulsed Guidance Law for Exoatmospheric Interception

I N THE absence of aerodynamic control forces, divert thrusters are used to generate lateral acceleration required by interceptors in exoatmospheric engagements. The thrusters can be fired as pulses of constant lateral acceleration magnitude. The pulse width or duration is determined by the guidance law used. Several guidance laws based on the zero-effort miss (ZEM) have been proposed in the literature for exoatmospheric engagements. A ZEM-based midcourse guidance law for intercepting strategic targets and spacecraft is proposed by Newman [1,2]. Later, Hablani [3] proposed a ZEM-based pulsed guidance law with image processing delays for the endgame phase. Hablani [4] also demonstrated that zero-effort-miss guidance is superior to proportional navigation guidance in the endgame phase of an exoatmospheric interception using divert thrusters. Miss distance analysis of ZEM guidance with respect to various measurement errors is carried out by Hablani and Pearson [5]. Zarchan [6] has derived the ZEM-based pulsed guidance implemented with equally spaced pulses. ZEMcan be calculated from the instantaneous line-ofsight rate, the closing velocity, and the time to go. Lateral distance traveled by the interceptor depends on the width of the divert thruster pulse, which can be evaluated by equating the lateral distance to the ZEM. Because of the assumptions involved, interception cannot be guaranteed by firing a single pulse even for exact ZEM calculations. Subsequent pulses are fired to achieve a satisfactory miss distance. In this Note, an interceptor endgame pulsed guidance law is derived by attempting to attain the collision heading rather than negating ZEM. It is shown that this strategy is more effective than ZEM-based guidance for intercepting targets with higher heading angles off the nominal head-on case. The interceptor and the target are assumed to be constant speed lag free point masses. A planar engagement scenario is considered, assuming that the interceptor can perform uncoupled guidance commands in two perpendicular planes. The proposed guidance law, to determine the pulse width, is derived in steps by 1) obtaining the collision heading based on the collision triangle engagement geometry and parameters, and 2) computing the width of the pulse fired by the divert thruster to attain the collision heading. In the existing pulsed guidance laws, the pulses are fired either at fixed intervals [6] or every time the line-of-sight rate (or zero-effort miss) crosses a predefined threshold value [3]. In the present work, it is shown that firing pulses in quick succession results in minimum pulse widths and hence minimum control effort for interception. For comparative study, the derivation of pulsed guidance [6], based on ZEM steering, is described briefly in Sec. II. The proposed guidance law is derived in Sec. III, followed by some results on pulse firing sequence in Sec. IV. Simulation studies are carried out in Sec. V.

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

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.

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

An integrated Autopilot-and Guidance Algorithm is developed using Higher Order Sliding Mode Control, for interceptors steered by combination of aerodynamic lift, sustainer thrust and center of gravity divert thrusters. A smooth HOSM guidance generates flight path trajectory angular rates and attitude rate commands. The attitude rate maneuvers are aimed at producing desired aerodynamic lift and / or orienting sustainer thrust. The lateral acceleration created by the attitude maneuver is treated as a “cooperative” disturbance and accounted for by the trajectory control. SOSM autopilot design is based on nonlinear dynamic sliding manifold. Proposed 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.

Missile Controlled by Lift and Divert Thrusters Using Nonlinear Dynamic Sliding Manifolds

A new approach to the design of a kinetic-kill longitudinal autopilot steering the missile trajectory by the combination of aerodynamic lift and divert thrusters (dual-thrusters control) and with its attitude oriented by attitude thrusters is presented. The proposed approach may increase the total divert acceleration capability by up to 100%, improving the end-game intercept accuracy, when such capability is required. The pitch plane autopilot design is based on second-order sliding mode control using a nonlinear dynamic sliding manifold technique. A robust high-accuracy tracking of the missile normal acceleration guidance command is achieved in the presence of considerable model uncertainties created by the interactions between the airflow and the thrusters jets. Results of the computer simulation demonstrate excellent, robust, high-accuracy tracking performance of the proposed design.

Endgame Guidance and Relative Navigation of Strategic Interceptors with Delays

This paper is concerned with endgame guidance and relative navigation of strategic missiles equipped with divert thrusters to intercept nonmaneuvering reentry vehicles. The paper has three objectives: 1) demonstrate superiority of predictive guidance to pulsed proportional navigation guidance; 2) devise compensation for a onesample delay in arrival of line-of-sight (LOS) angle measurements at Kalman-filter module and another one-sample delay in communication of LOS rate estimates to the guidance module; and 3) develop a Kalman filter insensitive to modeling errors for relative navigation. A hybrid Kalman filter is developed for relative navigation with angle measurements—hybrid because relative position and velocity vectors are propagated in Cartesian coordinates, whereas the error covariance is propagated, and updated at the time of measurement, in spherical coordinates. For spherical computations, a time-to-go dependent transition matrix and a time-to-go dependent process noise matrix are formulated. It is shown that these time-to-go dependent matrices reduce the miss distance and its sensitivity to modeling errors. With this Kalman filter, it is shown that the closed-loop predictive guidance yields smaller miss distance, consumes less fuel, and requires fewer divert firings than the pulsed proportional navigation guidance does. The delay compensation is accomplished in two steps: first, the high-frequency stream of ∆V measurements from an accelerometer is stored and synchronized with the delayed LOS angle measurement from the seeker; second, the delayed LOS rate estimate from the Kalman filter is propagated forward using the measured ∆V that transpires during the delay. The miss distance of the zero-delay level and robustness to modeling errors is shown to be almost recovered.

Modeling of Chemically Reacting Flowsfrom a Side Jet at High Altitudes

The interaction of a jet from a 60-lbf (267-N) thruster positioned on the side of a small rocket, with the rarefied atmosphere between altitudes of 80 to 160 km, is studied numerically. A multistep approach is employed, which combines the successive computation of the flow inside the thruster using a Navier‐Stokes solver, the axisymmetric plume core flow, and chemically reacting three-dimensional plume-atmosphere interaction using the direct simulation Monte Carlo method, and obtains the UV radiation fields based on the flow solution. The impact of the boundary layer inside the nozzle as well as the rocket speed (5‐8 km/s) and flight altitude (80‐160 km) on the plume-atmosphere interaction is examined. I. Introduction A TMOSPHERIC interceptor (AI) vehicles currently being developed use divert and attitude reaction control systems (RCS) to perform quick maneuvers during flight. The forward and aft RCS engines provide the thrust for attitude (rotational) maneuvers (pitch, yaw, and roll) and for small velocity changes along the rocket axis (translation maneuvers). In the past few years extensive experimental and theoretical studies have been undertaken to predict RCS jet interactions with the ambient atmosphere accurately. Although significant progress has been achieved in understanding the phenomenology of the jet-atmosphere interaction, several important problems still need to be resolved, particularly with regard to supersonic and hypersonic flows at low and high altitudes. Experimental measurements have been made to characterize the jet-atmospheric interaction in wind-tunnel facilities. However, it is difficult to scale these results 1 to flight conditions because of freestream/plume chemistry, Reynolds number, surface, and wall interference effects. Computational studies are therefore necessary to supplement and expand experimental efforts and obtain credible information on performance characteristics and flowfield structure of RCS thrusters at different altitudes. At low altitudes (up to about 60 km) the freestream-jet interaction is characterized by thin plume and bow shock waves and a jet-induced separation region in front of the divert thruster side jet. 2 The flow separation is responsible for the elevated pressure in this region and corresponding thrust amplification and might also affect the operation of optical sensors located upstream from the nozzle. 3,4 The problem of sensor contamination was considered in Ref. 5. AI vehicle aerodynamics and thruster performance at low altitudes were examined numerically by several workers, mostly using solutions of the Navier‐Stokes equations. 6−9 The jet-atmosphere interaction at high altitudes is qualitatively different from that at low altitudes. The flow separation in front of the jet becomes less pronounced and diminishes at altitudes 80 km and higher, and the effect of the atmosphere on the plume is much weaker. Because of the rarefaction, the jet interaction impact on the vehicle aerodynamics is significantly smaller. The direct simulation Monte Carlo (DSMC) method was used in Ref. 10 to compute the three-dimensional jet interaction for a corner flow configuration,

Miss Distance Error Analysis of Exoatmospheric Interceptors

A miss distance error analysis of exoatmospheric interceptors with liquid divert thrusters in endgame is presented. The analysis relates miss distance uncertainty to 1) infrared/electrooptic seeker noise, 2) seeker latencies caused by image processing and motion compensation and integration to detect the target on the focal plane, 3) seeker scan angle measurement errors of gyro, 4) range error or time-to-go error, and 5) filtration of all these errors through a Kalman filter for estimating the line-of-sight (LOS) rates. With the relationships developed, the missile designer/analyst can predict a probable miss distance with the hardware parameters of the selected divert thrusters and sensor suite, as well as software parameters of the selected guidance policy. The analyst can also arrive at the LOS rate estimation accuracy requirements about the azimuth and elevation axes, to be achieved by a tracking Kalman filter for specified miss distance statistics. The analysis provides insight and predicts reliably the miss distances expected from Monte Carlo simulations and hardware-in-the-loop laboratory tests. I. Introduction M ISS distance error analysis of air missiles seems to be well understood, 1,2 but that of exoatmospheric interceptors equipped with infrared sensors and liquid divert thrusters appears to be lagging behind. To be sure, Refs. 3‐6 provide a system engineer design tools to arrive at hardware/software requirements for exoatmospheric interceptors to meet a specific miss or hit accuracy without the use of an elaborate trajectory simulation. However, these references do not specifically address unique features of exoatmospheric interceptors with liquid propulsion systems. This paper attempts to fill this gap.

Architecting a Missile System for Navy Theater Ballistic Missile Defense

ABSTRACT While most of the theater ballistic missiles (TBM) in threat countries' inventories are of the shorter range SCUD varieties, mid‐ to long‐range versions are currently in development in a number of third world countries. Threat potential exists in the following three battle spaces: endo‐atmosphere (0–30 km), high endo‐atmosphere (30–70 km) and exo‐atmosphere (greater than 70 km). The inherent short range and low speed of endo‐atmospheric threats match well with capabilities of SM‐2 Block IVA, which equips the Navy with an area defense capability. The exo‐atmospheric TBMs are longer range and can threaten more targets which may be widely dispersed. Their higher velocities reduce response times dramatically. Therefore, exo‐atmospheric TBMs create the need for Standard Missile‐3 (SM‐3), which provides the Navy with theater wide defense capability.Defining its area and theater wide systems as clearly endo‐atmospheric and exo‐atmospheric systems allows the Navy to use derivatives of the Standard Missile Block IV to take full advantage of the conditions associated with each of these operating zones. Use of an existing missile and ship system baseline also allows use of the existing interface structure to minimize cost.To counter the endo‐atmospheric TBMs, the SM‐2 BLK IVA upgrades include an advanced imaging infrared (IIR) seeker, an improved fast‐reaction auto pilot and a forward looking RR all in the same volume as the existing missile. The highly responsive SM‐2 Block IVA missile, complemented with Aegis weapons systems modifications, provides capability against enemy aircraft and cruise missiles, as well as TBMs. Standard Missile‐3 (SM‐3) replaces the SM‐2 Block IV warhead, radar and guidance section with a boosted third stage and an advanced kinetic warhead (KW). Operation in the exo‐atmospheric region permits a KW design with autonomous guidance control and divert thrusters for high maneuverability and has the capability of achieving very high interceptor velocities.

Postflight data-reduction techniques for hovered kinetic energy weapons

One of the major concerns in attitude control of kinetic-energy-weapon vehicles proposed for exoatmospheric engagements is that error sources such as vehicle center-of-gravity variations and thrust misalignments of divert thrusters may induce excessive parasitic torques on the vehicle. These torques may result in significant attitude perturbations that are undesirable for target tracking, especially when using strapdown seekers. In an effort to identify and characterize these error sources in a quantitative way, it has been proposed to explicitly estimate them by using the data obtained from kinetic-energy-weapon hovered interceptor test flights. An estimation procedure based on extended Kalman filtering is developed for the postflight data reduction that includes the estimation of the center-of-gravity location and thrust misalignments. Some modifications to the basic extended Kalman filter algorithm are made to reduce the sensitivity of the estimator to uncertainties in divert thrust timing. The proposed estimation scheme is then used to process simulated flight data and actual test data. The data-reduction results are provided in this paper.