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This paper proposes an integrated cooperative guidance strategy for multiple unmanned aerial vehicles to simultaneously attack multiple targets. To solve the challenging cooperative attack problem coupled with the target assignment problem simultaneously in a unified one-phase framework, a novel auxiliary index denoted as the integrated relative distance is designed and analysed from two opposite perspectives. Because of its special properties, the proposed integrated relative distance plays a crucial role in evaluating dynamic engagement situations between attackers and targets, and it further provides a criterion for whether the cooperative attack mission is completed. Then all integrated relative distances are minimised to directly derive a cooperative guidance law for attackers based on optimal control theory. The simultaneous minimisation of integrated relative distances from both forward and backward perspectives ensures that each target is destroyed by at least one attacker and all attackers are fully exploited to obtain optimal attack performance. By this means, the target assignment problem is automatically solved and the cooperative simultaneous attack is achieved. Several simulation cases are presented to illustrate the effectiveness of the proposed integrated cooperative guidance strategy. [ABSTRACT FROM AUTHOR]

This paper discusses the optimal control law, in a three-dimensional (3D) heliocentric orbit transfer, of a spacecraft whose primary propulsion system is a Solar Wind Ion Focusing Thruster (SWIFT). A SWIFT is an interesting concept of a propellantless thruster, proposed ten years ago by Gemmer and Mazzoleni, which deflects, collects, and accelerates the charged particles of solar wind to generate thrust in the interplanetary space. To this end, the SWIFT uses a large conical structure made of thin metallic wires, which is positively charged with the aid of an electron gun. In this sense, a SWIFT can be considered as a sort of evolution of the Janhunen’s E-Sail, which also uses a (nominally flat) mesh of electrically charged tethers to deflect the solar wind stream. In the recent literature, the optimal performance of a SWIFT-based vehicle has been studied by assuming a coplanar orbit transfer and a two-dimensional scenario. The mathematical model proposed in this paper extends that result by discussing the optimal guidance laws in the general context of a 3D heliocentric transfer. In this regard, a number of different forms of the spacecraft state vectors are considered. The validity of the obtained optimal control law is tested in a simplified Earth–Venus and Earth–Mars transfer by comparing the simulation results with the literature data in terms of minimum flight time.

Abstract Satellite Image classification provides information about land use land cover (LULC) and this is required in many applications such as Urban planning and environmental monitoring. Recently, deep learning techniques were applied for satellite image classification and achieved higher efficiency. The existing techniques in satellite image classification have limitations of overfitting problems due to the convolutional neural network (CNN) model generating more features. This research proposes the optimal guidance-whale optimization algorithm (OG-WOA) technique to select the relevant features and reduce the overfitting problem. The optimal guidance technique increases the exploitation of the search technique by changing the position of the search agent related to the best fitness value. This increase in exploitation helps to select the relevant features and avoid overfitting problems. The input images are normalized and applied to AlexNet–ResNet50 model for feature extraction. The OG-WOA technique is applied in extracted features to select relevant features. Finally, the selected features are processed for classification using Bi-directional long short-term memory (Bi-LSTM). The proposed OG-WOA–Bi-LSTM technique has an accuracy of 97.12% on AID, 99.34% on UCM, and 96.73% on NWPU, SceneNet model has accuracy of 89.58% on AID, and 95.21 on the NWPU dataset.

The miss distance analysis of the first-order explicit guidance law (EGL) is carried out using linearized equation of motion in the normalized form in order to obtain normalized miss distance curves. The initial heading error, constant target, acceleration limit, radome refraction error, and fifth-order binomial control system are considered. Moreover, body rate feedback is added to the explicit guidance law as a well-known classical compensation method of the radome effect as in proportional navigation. The analysis is performed for different values of the power of the alpha function, defined as the time decrease rate of the zero-effort miss to unit control input. As a special case, the EGL with unit power gives the first-order optimal guidance strategy for minimizing the integral of the square of the commanded acceleration during the total flight time. For the performance/stability analysis, the rms miss distance versus turning rate time constant and radome slope can be plotted for different values of the power of alpha function.

In order to improve the adaptability and robustness of gliding guidance under complex environments and multiple constraints, this study proposes an intelligent gliding guidance strategy based on the optimal guidance, predictor-corrector technique, and deep reinforcement learning (DRL). Longitudinal optimal guidance was introduced to satisfy the altitude and velocity inclination constraints, and lateral maneuvering was used to control the terminal velocity magnitude and position. The maneuvering amplitude was calculated by the analytical prediction of the terminal velocity, and the direction was learned and determined by the deep Q-learning network (DQN). In the direction decision model construction, the state and action spaces were designed based on the flight status and maneuvering direction, and a reward function was proposed using the terminal predicted state and terminal constraints. For DQN training, initial data samples were generated based on the heading-error corridor, and the experience replay pool was managed according to the terminal guidance error. The simulation results show that the intelligent gliding guidance strategy can satisfy various terminal constraints with high precision and ensure adaptability and robustness under large deviations. [ABSTRACT FROM AUTHOR]

This paper presents an impact-angle-control guidance law with terminal constraints on the curvature of the missile trajectory. The formulation takes into account nonlinear kinematics and time-varying velocity, allowing for more general cases in which the flight path angle may not be small throughout the entire trajectory. The proposed optimal guidance law aims to minimize the change rate of pseudo-curvature, which is defined as a curvature with weighting factors based on the flight-path angle and range to go. The analysis shows that the trajectories generated by this nonlinear guidance law have a simple polynomial form with respect to downrange. Numerical simulations demonstrate that the cost function parameter can be used to shape the trajectories along the downrange. [ABSTRACT FROM AUTHOR]

Satellite Image classification provides information about land use land cover (LULC) and this is required in many applications such as Urban planning and environmental monitoring. Recently, deep learning techniques were applied for satellite image classification and achieved higher efficiency. The existing techniques in satellite image classification have limitations of overfitting problems due to the convolutional neural network (CNN) model generating more features. This research proposes the optimal guidance-whale optimization algorithm (OG-WOA) technique to select the relevant features and reduce the overfitting problem. The optimal guidance technique increases the exploitation of the search technique by changing the position of the search agent related to the best fitness value. This increase in exploitation helps to select the relevant features and avoid overfitting problems. The input images are normalized and applied to AlexNet–ResNet50 model for feature extraction. The OG-WOA technique is applied in extracted features to select relevant features. Finally, the selected features are processed for classification using Bi-directional long short-term memory (Bi-LSTM). The proposed OG-WOA–Bi-LSTM technique has an accuracy of 97.12% on AID, 99.34% on UCM, and 96.73% on NWPU, SceneNet model has accuracy of 89.58% on AID, and 95.21 on the NWPU dataset. [ABSTRACT FROM AUTHOR]

Trajectory optimization is a familiar method for most of re-entry and Re-usable vehicles. This is because of the ability to include almost all of the problem constraints without facing restrictions such as time & Calculation issues. Adding or removing constraints in trajectory optimization problem has significant effects on overall optimization performance which even can upgrade the method to an on-line process. Most of optimization Algorithms such as nonlinear-programming need an initial guess and are also sensitive to it. Hence in this research management of initial guess is done to remove some constraints from optimization problem and transfer them to initial phase. Accordingly an effort is conducted through using a classic guidance method to satisfy constraints of distance error and angle of impact command. The output of guidance initial guess is then fed to the optimization problem. 6Dof Simulation results show the increase of optimization performance via reduced number of iterations and Optimization time and increased solution accuracy.

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Launcher reusability is the most effective way of reducing access to space costs, but remains a great technical challenge for the European aerospace industry. One of the challenges lies in the recovery GNC strategy and algorithms, in particular those of the powered-landing phase, which must enable a precise landing with low fuel margins and significant dispersions. While state-of-the-art solutions for Navigation and control problems can be applied, namely, hybrid Navigation techniques and robust control, for the powered descent guidance problem novel techniques are required to enable on-board optimization, that is necessary to achieve the landing accuracy required to recover the first stage of a launcher. This paper presents the GNC solution currently in development by DEIMOS Space for RETALT (Retro Propulsion Assisted Landing Technologies), an EU Horizon 2020 funded project for studying launch system re-usability technologies for different classes of vertical take-off vertical-landing vehicles. At first, the architecture of the GNC solution identified for the return mission of the launcher is presented. Then, the paper focuses on the landing phase guidance solution, whose performance is critical to enable the recovery and, therefore, the reusability of the launcher making use of retro-propulsion. The guidance strategy is based on direct optimal control methods via on-board optimization, which is necessary to satisfy the pinpoint landing requirement in a high uncertain dynamic system, such as a booster recovery mission. Online convex optimization and successive convexification are explored for the design of the guidance function. The proposed guidance solution was integrated and tested in a high-fidelity simulator and the performance was preliminary assessed. The guidance assessment allowed selecting the best algorithms to be further consolidated and integrated in an end-to-end GNC solution for the return mission. [ABSTRACT FROM AUTHOR]

In the last decade, the rapid and successful development of reusable launch systems such as SpaceX' Falcon 9 demonstrated both the operational feasibility of reusable launchers and their economic viability. The objective of recovering a launcher or a launcher's booster requires to safely return the launch vehicle from orbital or sub-orbital conditions to a soft landing. To increase the reusability, decrease the turnaround time and reduce costs, a precise touchdown on a pre-defined landing site or on a floating barge on the Ocean is preferred to splashdown in the water, due mainly to the highly detrimental effect of the salted water on the launcher components and equipment. The project RETALT (Retro Propulsion Assisted Landing Technologies) was funded by the EU Horizon 2020 program to study and develop critical technologies for launcher recovery based on retro-propulsion. In this context, and based on in-house experience and tools, DEIMOS Space carried out the mission engineering of the RETALT1 vehicle concept to assess the feasibility of a return mission, from a wide range of Main Engine Cut-Off (MECO) conditions, when the stage is separated from the rest of the launch vehicle, in line with the available propellant budget, and while maintaining the peak entry conditions within acceptable limits. Either a DownRange Landing (DRL) on a drone ship at sea or a Return To Launch Site (RTLS) to land in the proximity of the launch pad is performed based on the velocity and distance at MECO from the launch site. For the landing burn, a safe splashdown approach has been implemented to avoid damaging the ground infrastructure in case of anomalies during the flight. Based on the mission feasibility assessment, the needs for the vehicle recovery have been identified, leading to the definition of preliminary mission requirements at the system and subsystem level. Consequently, the consolidation of the return mission design was possible and optimised trajectories have been defined for the DRL and RTLS scenarios. [ABSTRACT FROM AUTHOR]

This paper presents an impact-angle-control guidance law with terminal constraints on the curvature of the missile trajectory. The formulation takes into account nonlinear kinematics and time-varying velocity, allowing for more general cases in which the flight path angle may not be small throughout the entire trajectory. The proposed optimal guidance law aims to minimize the change rate of pseudo-curvature, which is defined as a curvature with weighting factors based on the flight-path angle and range to go. The analysis shows that the trajectories generated by this nonlinear guidance law have a simple polynomial form with respect to downrange. Numerical simulations demonstrate that the cost function parameter can be used to shape the trajectories along the downrange.

This article investigates the problem of scale factor error estimation for missiles with a strapdown seeker. We first theoretically prove that the seeker scale factor error will introduce a parasitic loop in the guidance loop and, hence, result in instability of the guidance loop. This fact reveals that it is necessary to extract the required guidance information and estimate the scale factor error simultaneously in real time. To this end, the observability of scale factor error estimation is analyzed, and the theoretical results show that this information is weakly or poorly observable under conventional proportional navigation guidance. To address this problem, a new optimal guidance law that can be utilized to improve the estimation performance is proposed by optimizing terminal miss distance, control energy, and observability in an integrated manner. The characteristics of the proposed optimal guidance law are also theoretically analyzed to provide better insights of the proposed approach. Extensive numerical simulations are conducted to support the proposed algorithm. [ABSTRACT FROM AUTHOR]

A new design method for trajectory shaping guidance laws with the impact angle constraint is proposed in this study. The basic idea is that the multiplier introduced to combine the equations for the terminal constraints is used to shape a flight trajectory as desired. To this end, the general form of impact angle control guidance (IACG) is first derived as a function of an arbitrary constraint-combining multiplier using the optimal control. We reveal that the constraint-combining multiplier satisfying the kinematics can be expressed as a function of state variables. From this result, the constraint-combining multiplier to achieve a desired trajectory can be obtained. Accordingly, when the desired trajectory is designed to satisfy the terminal constraints, the proposed method directly can provide a closed form of IACG laws that can achieve the desired trajectory. The potential significance of the proposed result is that various trajectory shaping IACG laws that can cope with various guidance goals can be readily determined compared to existing approaches. In this study, several examples are shown to validate the proposed method. The results also indicate that previous IACG laws belong to the subset of the proposed result. Finally, the characteristics of the proposed guidance laws are analyzed through numerical simulations. [ABSTRACT FROM AUTHOR]

The orbital pursuit-evasion game (OPE) is a topic of research that has been attracting increasing attention from scholars. However, most works based on the relative dynamics under a short-distance assumption is not applicable when the distance between two spacecrafts is too large. Accordingly, there should be two phases in the OPE, a long-distance OPE (LDOPE) as well as a short-distance one. This article concerns on the optimal guidance problem for the LDOPE. Two different models are introduced in this article to formulate the LDOPE, namely, the Cartesian model, and the spherical model. Then, to overcome the unacceptable solution computation time of traditional algorithms, such as the differential evolution (DE), a well-designed algorithm called “mixed global-local optimization strategy” (MGLOS), which consists of the global optimization phase, and the local optimization phase, is introduced in this article. The MGLOS is nearly two orders of magnitude more efficient than the DE. Moreover, simulations under different initial conditions demonstrate the robustness of the algorithm, and the accuracy, and efficiency of the Cartesian, and spherical models, respectively. Finally, the robustness of two models is analyzed by Monte Carlo simulation, which provides a quantified way to make a choice between two models depending on the measurement accuracy, and permitted maximum error. [ABSTRACT FROM AUTHOR]

The atmospheric reentry phase is one of the most important mission steps in space missions, therefore, the guidance and control of reentry vehicles in this phase of mission is important. In this article, a reentry vehicle guidance algorithm is proposed which has suitable robustness in the presence of initial reentry parameters uncertainty. To use any conductive method, first the motion equations must be obtained. In this paper, quadratic nonlinear control method is used to guide the vehicle. In this regard, the equations of motion of reentry vehicles are developed in form of state space and the system and control matrices depending on the state and control variables are extracted. In this article, it is tried to minimize the landing errors at terminal point using Nonlinear Quadratic Tracking (NQT) and chasing a reference trajectory. In order to define a trajectory with different initial states using evolutionary genetic algorithm with changes in weighting matrices Q and R, it is tried to reduce the errors of landing at terminal point. Monte Carlo analysis is used to evaluate the performance of the proposed algorithm. According to the results, the proposed algorithm can reduce the errors more than 90% in the presence of reentry initial parameter uncertainties.

The atmospheric re-entry phase is one of the most significantmission steps in the space missions;hence, theguidance and control of reentry vehicles in this phase of mission is important. In this article, a reentry vehicle guidance algorithm has been proposed which has suitable robustness in the presence of initial reentry parameters uncertainties. Here,it has been tried to minimize the landing errors at terminal point using Nonlinear Quadratic Tracking (NQT) and chasing a reference trajectory. In order to define several trajectories with different initial states using evolutionary genetic algorithm with changes in weighting matrices Q and R, it hasbeen tried to reduce the errors of landing at terminal point. The reentry position of the reentry vehicles may be different from the desired ones with respect to several events. In this situation, reentry vehicles start to move in a new trajectory which is not suitable. Therefore, the reentry vehicles should be guided to come back into the desired trajectory or a new optimum trajectory needs to be redesignedto have the same target position on the ground. To do this, we need optimum weighting matrices R and Q for every new trajectory. In this article, this problem has been resolved using partial least squares regression; meanwhile, obtaining the optimal matrices by genetic algorithms needed many times. Also,it is shown that using this method, in the presence of reentry uncertainties, weighting matrices for each new initial condition hasbeen quickly derived. Additionaly,through the matrices obtained and the nonlinear quadratic tracking controller, reentry vehicle was directedto the target with a good accuracy. The Monte Carlo analysis has been used to evaluate the performance of the proposed algoritm. According to the results, the proposed algoritm has a suitable accuracy level and it can generate the online optimum trajectory.

This article presents an optimal guidance approach for unmanned aerial vehicle navigation between two given points in 3D considering the wind influence. The proposed cost function to be minimized involves the weighting of the travel time and the control energy. An analytical expression is derived for the optimal cost yielding a fourth-order polynomial, whose positive real roots correspond to the optimal travel times. The optimization problem is shown to be equivalent to the zero-effort-miss/zero-effort-velocity optimal guidance approach for the case of a constant wind acceleration. Case studies for rendezvous and intercept problems are shown through simulation examples for different wind conditions. [ABSTRACT FROM AUTHOR]

In this article, a minimum-fuel powered-descent optimal guidance algorithm that incorporates obstacle avoidance is presented. The approach is based on convex optimization that includes the obstacles using nonconvex functions. To convert these nonconvex obstacle constraints to convex ones, a simple linearization procedure is employed. It is proved that the optimal solution of the convex relaxation problem is also optimal for the original nonconvex one. The sensitivity of the multiobstacle avoidance method to the relaxation factor and its effectiveness under different conditions are also investigated through simulations. [ABSTRACT FROM AUTHOR]

This paper proposes an optimal impact angle control guidance law for homing missiles with a narrow field-of-view of the seekers. As groundwork for designing a guidance law, we first present a general guidance structure that can achieve any terminal constraint of the line-of-sight rate based on the optimal control theory. We configure the desired profile of the line-of-sight rate using a saturation function whose exact form is determined to satisfy the required boundary conditions. By combining the line-of-sight rate profile with the optimal guidance structure, we develop a guidance law that achieves an impact angle interception with the field-of-view constraint. Herein, as the entire guidance structure is derived based on exact kinematics without any approximation, the proposed law ensures the accurate impact angle interception for various engagement scenarios. This precise consideration of the engagement kinematics also accurately ensures the energy optimality of preventing the excessive use of control inputs when homing. To evaluate the performance of the proposed method, numerical simulations with various engagement scenarios are conducted, and the results demonstrate that the proposed law allows missiles to accurately intercept their targets with the desired impact angles and without violating the prescribed field-of-view constraint.

An optimal explicit guidance law that maximizes terminal velocity is developed for the reentry of a vehicle to a fixed target. The equations of motion are reduced with differential flatness approach and acceleration commands are related to the parameters of trajectory. An optimal trajectory is determined by solving a real-coded genetic algorithm. For online trajectory generation, optimal trajectory is approximated. The approximated trajectory is compared with the pure proportional navigation and genetic algorithm solutions. The near optimal terminal velocity solution compares very well with these solutions. The approach robustness is examined by Monte Carlo simulation. Other advantages such as trajectory representation with minimum parameters, applicability to any reentry vehicle configuration and any control scheme, and Time-to-Go independency make this guidance approach more favorable.

This paper presents the design of a singular-perturbation-based optimal guidance with constraints on terminal flight-path angle and angle of attack. By modeling the flight-control system dynamics as a first-order system, the angle of attack is introduced into the performance index as a state variable. To solve the resulting high-order optimal guidance problem analytically, the posed optimal guidance problem is divided into two sub-problems by utilizing the singular perturbation method according to two time scales: range, altitude, and flight-path angle are the slow time-scale variables while the angle of attack is the fast time-scale variable. The outer solutions are the optimal control of the slow-scale subsystem. Thereafter, by applying the stretching transformation, the fast-scale subsystem establishes the relationships between the outer solutions and acceleration command. Then, the optimal command can be obtained by solving the fast-scale subsystem also using the optimal control theory. The proposed guidance can achieve a near-zero terminal acceleration as well as a small miss distance. The superior performance of the guidance is demonstrated by adequate trajectory simulations. [ABSTRACT FROM AUTHOR]

Highlights • Analytical expression of docking guidance with rotating target is researched. • Flying-around approaching method maintains a stable relative position for docking. • Optimal docking method is energy saving and collision avoidance method is given. • Collision avoidance using two-phase guidance is presented. Abstract The guidance and control strategy for spacecraft rendezvous and docking are of vital importance, especially for a chaser spacecraft docking with a rotating target spacecraft. Approach guidance for docking maneuver in planar is studied in this paper. Approach maneuver includes two processes: optimal energy approach and the following flying-around approach. Flying-around approach method is presented to maintain a fixed relative distance and attitude for chaser spacecraft docking with target spacecraft. Due to the disadvantage of energy consumption and initial velocity condition, optimal energy guidance is presented and can be used for providing an initial state of flying-around approach process. The analytical expression of optimal energy guidance is obtained based on the Pontryagin minimum principle which can be used in real time. A couple of solar panels on the target spacecraft are considered as obstacles during proximity maneuvers, so secure docking region is discussed. A two-phase optimal guidance method is adopted for collision avoidance with solar panels. Simulation demonstrates that the closed-loop optimal energy guidance satisfies the ending docking constraints, avoids collision with time-varying rotating target, and provides the initial velocity conditions of flying-around approach maneuver. Flying-around approach maneuver can maintain fixed relative position and attitude for docking. [ABSTRACT FROM AUTHOR]

Abstract Descent & landing (D&L) on small planetary bodies are scientifically rewarding exploration missions but they are technically challenging due to the complex and poorly-known environment around those bodies. The standard guidance synthesis approach considers nominal conditions and applies optimal control theory to obtain guidance law gains, followed by intensive verification and validation. In this article, it is shown that the standard approach may yield gains that are not optimal once dispersions (and/or other optimality metrics) are taken into account and a tuning approach is then proposed based on a priori methodological system assessment. The proposed approach employs systematic high-fidelity simulations to generate trade-off maps. These maps can be generated by on ground operators based on the best estimated conditions and uploaded to the spacecraft as it approaches the target. The proposed systematic guidance tuning and resulting maps also provide a valuable understanding of the system dynamics towards the application of other industry-oriented tools such as structured ℋ ∞ optimisation. It is shown that the proposed tuning enables propellant consumption reductions of around 40% compared to state-of-practice gain selections. [ABSTRACT FROM AUTHOR]

Abstract A novel cruise guidance method that can meet the terminal latitude and longitude, time and angle constraints independent of time-to-go prediction is studied in this paper. The impact time and angle control guidance (ITACG) goal is achieved through two steps. The first step is the construction of the complete motion model, the constant altitude and velocity guidance laws, and the optimal proportional guidance law which can satisfy the terminal position constraint based on the in-plane relative motion equation. The impact time control is the second step, which can be achieved by the integrated utilization of optimal guidance and maneuvering flight. Additional maneuvering trajectories used to control the impact time are designed in both cases with and without angle constraint, and the maneuvering parameters are directly solved online without time-to-go prediction. Simulation results verify the effectiveness of the guidance method, as well as the adaptability to different initial conditions and terminal constraints. [ABSTRACT FROM AUTHOR]

Abstract Maximizing the terminal altitude for Mars atmospheric entry has long been investigated on trajectory design to allow a sufficient timeline margin for subsequent operations and the scientific requirements of exploring Mars ancient highland. The purpose of this paper is to design an onboard Mars atmospheric entry guidance algorithm, which can achieve the optimal terminal altitude at the predetermined terminal flight range. Two important characters that the optimal bank angle profile are revealed in this paper, offering the gateway to the application of the optimal guidance by onboard parameter searching, which will be accomplished by the numerical predictor-corrector strategy. Moreover, suboptimal situations are also investigated considering the performance restriction of reactive control system (RCS). Effectiveness of the proposed guidance algorithm is demonstrated using scenarios of the Mars Science Laboratory (MSL) mission. Highlights • Optimal guidance achieves terminal position accuracy and optimal terminal altitude. • Two important characters that the optimal bank angle profile are revealed. • The optimal guidance algorithm does not rely on offline trajectory optimization. [ABSTRACT FROM AUTHOR]

Abstract A renewed scientific interest has been growing in the exploration of small asteroids in addition to larger planetary bodies such as Mars, since their weaker gravitational field makes them more easily accessible. However, such exploratory missions are very challenging from an engineering perspective, particularly when striving for optimal propellant consumption. This is mostly due to the perturbed and poorly known characteristics of small planetary bodies but also, as shown by the European Rosetta mission, to the long-time degradation of spacecraft subsystems. In order to address this challenge, it has been long recognised the need for robust descent algorithms. However, Space guidance and control communities have different understandings, restricting the integration of scientific advances and even constraining their capabilities. To incite such an integration and guide engineers in the development of planetary descent algorithms, this survey gathers state-of-practice guidance and control techniques and presents them in an instructive fashion. In addition, it clarifies and reconciles different concepts from both guidance and control perspectives. The survey and reconciliation of concepts then lead to the identification of an underlying parametric generalisation of guidance techniques, suitable for the application of systematic optimisation tools. Albeit simple, this structural identification is very important as the latter tools have shown great promise and have already been employed, for example, for Rosetta's critical control update. Finally, special emphasis is placed on the robustness of those techniques against uncertainties. [ABSTRACT FROM AUTHOR]

Aircraft point-mass equations of motion have been largely adopted to calculate optimal trajectories with local aerodynamic models, i.e. valid in a restricted domain. However, some optimal maneuvers may need aerodynamic models valid for a broader range of flight conditions. For this purpose, global aerodynamic models are attractive but their nonlinear structure can preclude obtaining optimal trajectories by an indirect method together with the point-mass equations of motion. To solve this impasse without resorting to direct methods the authors propose a new set of aircraft equations of motion. When compared to the point-mass equations, the proposed set permits the inclusion of the angular velocity in the evaluation of aerodynamic forces, making them more accurate. Another advantage of the proposed model over the point-mass one is that it allows a qualitative estimate of the control surface deflections after the trajectory is obtained, which enables to discard solutions with infeasible deflections. To verify consistency, the proposed equations of motion are compared by simulation to the point-mass and to the rigid-body equations. The use of the proposed set of equations is demonstrated by three optimizations of a 360 ∘ roll problem. [ABSTRACT FROM AUTHOR]

The paper presents a laser guided bomb guidance law based on the linear quadratic differential game theory, where a case of two perpendicular planes with two state variables in each plane has been considered. The Kalman filtering method has been used for noise removal from the measured signals and for estimation of the missing state variable values needed for the optimal guidance law. Optimisation has been conducted with respect to minimisation of the performance index. Comparative analysis of different guidance laws is done. A statistical analysis is performed to obtain the terminal miss distance in dependence on total flight time. [ABSTRACT FROM AUTHOR]

A novel multiple constrained adaptive gliding guidance method which is independent of quasi-equilibrium gliding condition (QEGC) and standard trajectory is proposed in this paper. The gliding guidance task is decomposed into longitudinal and lateral directions. In longitudinal direction, an altitude control model is established independent of QEGC, a hierarchical adaptive guidance strategy is introduced to control the vehicle to achieve equilibrium flight state and to meet the terminal altitude and flight-path angle constraints. In lateral direction, a heading error control model is constructed and the optimal control is employed to eliminate the heading error in real time with minimum energy consumption. In addition, the terminal velocity magnitude is predicted and corrected analytically based on lift–drag ratio, and the coordination strategy between guidance and velocity control is proposed to realize multi-constraint gliding guidance. This algorithm can generate angle-of-attack and bank angle commands which can meet the given terminal constraints with high precision based on the current flight states analytically, and has strong robustness to the initial deviation and environmental deviation. [ABSTRACT FROM AUTHOR]

This paper deals with the closed-loop form of mid-course guidance law design for accelerating missile system, whose acceleration is approximately constant. A midcourse guidance algorithm of feedback form is proposed to satisfy the engagement geometry conditions at the burn-out time for terminal homing performance enhancement. The effect of velocity change due to missile acceleration is explicitly considered in the derivation of the guidance law. The terminal constraint update algorithm is proposed under the assumption that the target trajectory is predicted precisely. Simulation results are provided to show the performance and characteristics of the proposed algorithm. [ABSTRACT FROM AUTHOR]

Besides enhancing the warhead effectiveness, the terminal lead angle constraint ensures that the target remains within the field of view of the radio proximity fuse, which is very critical for the success of a mission. This paper presents substantial enhancement of the generalized explicit guidance to include the terminal lead angle constraint in three dimension (3-D).. With the availability of the desired total lead angle in 3-D, it computes the necessary terminal flight path and heading angles so that it not only satisfies the lead angle constraint, but also leads to the minimum among all minimum control effort solutions for various possible terminal flight path and heading angle combinations. With no lead angle constraint, it also gracefully degenerates to the standard proportional navigation guidance. [ABSTRACT FROM PUBLISHER]

In this paper, an explicit optimal line-of-sight guidance law for second-order binomial control systems is derived in closed-loop without acceleration limit. The problem geometry is assumed in one dimension and the final time and final position are fixed. The formulation is normalized in three forms to give more insight into the design and performance analysis of the guidance law. The computational burdun of the guidance law is reasonable for now-a-day microprocessors; however curve fitting or look-up table may be used for the implementation of the second-order optimal guidance law. The performance of the second-order optimal guidance law is compared in normalized forms with zero-lag and first-order optimal guidance laws using third-, fourth-, and sixth-order binomial control systems with/without acceleration limit. Moreover, the effect of the final time, the equivalent time constant of the vehicle control system, the vehicle-to-target line-of-sight weighting factor in cost function, and acceleration limit are investigated. Normalized miss distance analysis shows that the miss distance of the second-order guidance law is smaller than the two mentioned schemes for small total flight times, especially with large maneuvering capability. [ABSTRACT FROM AUTHOR]

A novel highly constrained optimal gliding guidance algorithm, which can satisfy multiple path and terminal constraints in the presence of no-fly zones for hypersonic vehicle, is investigated in this paper. It employs line-of-sight angle to describe the relative location relationship between flight path and no-fly zones, and proposes boundaries selection algorithm for each no-fly zone, which is called avoidance strategy, with minimum energy consumption. The boundaries determined by avoidance strategy divide the flight process and guidance task into several stages. In each stage, it constructs guidance models based on quasi-equilibrium glide condition in longitudinal and lateral directions respectively, and employs optimal control to generate guidance law to satisfy terminal location, altitude, and flight path angle constraints. In addition, all the path constraints, including heating rate, dynamic pressure, and overload, are converted into the angle-of-attack constraint to ensure the guidance mission can be accomplished successfully. This algorithm is independent of standard trajectory, and all the guidance commands, including the bank angle and the angle of attack, can be calculated analytically in real time and adaptively. Finally, the simulation results of CAV-H indicate that the proposed strategy can guide the vehicle to satisfy multiple different constraints with high precision and avoid no-fly zones effectively. [ABSTRACT FROM AUTHOR]

One of the main challenges posed by the next space systems generation is the high level of autonomy they will require. Hazard Detection and Avoidance is a key technology in this context. An adaptive guidance algorithm for landing that updates the trajectory to the surface by means of an optimal control problem solving is here presented. A semi-analytical approach is proposed. The trajectory is expressed in a polynomial form of minimum order to satisfy a set of boundary constraints derived from initial and final states and attitude requirements. By imposing boundary conditions, a fully determined guidance profile is obtained, function of a restricted set of parameters. The guidance computation is reduced to the determination of these parameters in order to satisfy path constraints and other additional constraints not implicitly satisfied by the polynomial formulation. The algorithm is applied to two different scenarios, a lunar landing and an asteroidal landing, to highlight its general validity. An extensive Monte Carlo test campaign is conducted to verify the versatility of the algorithm in realistic cases, by the introduction of attitude control systems, thrust modulation, and navigation errors. The proposed approach proved to be flexible and accurate, granting a precision of a few meters at touchdown. [ABSTRACT FROM AUTHOR]

In this paper, new optimal guidance laws for a stationary or a slowly moving target are proposed by solving the optimal control problem with the energy performance index weighted by sinusoidal functions. This weighting function is adopted to shape the specific guidance command: introducing a small acceleration at the initial and final time. This property is desirable for reducing the sensitivity against the initial heading error and improving terminal guidance performance. Numerical simulations are performed to determine the performance of the proposed guidance laws. [ABSTRACT FROM PUBLISHER]