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Thin-walled tubular deployable composite boom (DCB) has gained significant attention due to its lightweight, simple structure and high packaging efficiency. The experimental and numerical simulation methods were employed in this paper to investigate the folding behavior of the DCB. Using T700/epoxy unidirectional reinforced prepreg as raw material, DCB specimens were prepared by the vacuum bag method. To complete the folding experiments of DCB specimens, a folding mechanism was designed and manufactured. Folding experiments of DCB specimens were conducted and folding moment versus rotational displacement curves were measured. In addition, a Finite Element Model (FEM) was established to predict the folding behavior of the DCB. Different failure criteria were considered in the numerical analysis, including the Tsai‐Hill criterion and maximum stress criterion. Prediction results using the FEM were compared with experimental results, and both sets of results were in good agreement. It is shown that the DCB can achieve the folding function without failure. The research results in this paper are of great significance to the practical engineering application of the DCB. [Display omitted] • Six tubular deployable composite boom specimens and a precise folding mechanism were designed and prepared. • Folding experiments were conducted to measure the folding moment and rotational displacement curves. • A finite element model for predicting the folding behavior of the tubular deployable composite boom was established. [ABSTRACT FROM AUTHOR]

Advancements in bulk metallic glass thermoplastic forming unleash the potential to fabricate microscale metallic features with unparalleled precision and durability, i.e. microspine gripping structures for mobility platforms on planetary or in-space robots. Four designs of grippers were fabricated and rigorously tested. Coefficient of friction on four test surfaces were measured and compared to SiC paper and Velcro. The hardness and elasticity of the bulk metallic glasses make them a game-changer in the field, offering a superior alternative to conventional nonmetallic grippers. • Microspine grippers were manufactured via thermoplastic forming of bulk metallic glasses using silicon-on-insulator molds. • The microspines excelled on textured surfaces, where sharp points anchor on to pores or softer regions. • Mechanical behavior of microspines were studied using nanoindentor-based micro-compression and computer simulations. [ABSTRACT FROM AUTHOR]

This paper presents an innovative prescribed-time moving path following control strategy for a stratospheric airship subject to time-varying error constraints and external disturbances. With the help of a new time-varying constraining function, the strict constraint requirements on initial conditions of the system are relaxed. A novel error transformation mechanism is incorporated into the guidance method to tackle the error constraint. The constructed prescribed-time disturbance observer can provide estimates of unknown disturbances and model dynamic couplings within the predefined time. The proposed scheme guarantees that the predetermined precision requirement can be implemented within the prespecified settling time regardless of initial conditions. Numerical simulations demonstrate the effectiveness of the presented control strategy. • A prescribed-time MPF control method is proposed for a stratospheric airship with time-varying error constraints and disturbances. • The initial constraint requirements are removed by a new prespecified error constraint function. • A novel error transformation mechanism is integrated into the control algorithm to address the error constraint. • Estimates of model dynamic couplings and disturbances are obtained from the PTDO. [ABSTRACT FROM AUTHOR]

The processes occurring in near and far space have a significant impact on the safety of space flights. Processes such as explosions of supernovae and relativistic jets have a particularly great impact. They arise in quasars, black holes, protostars. Often such structures are observed during the formation of new stars. Such jets have a great impact on the safety of space flights. For example, there is currently no radiation shelter on the International Space Station, so astronauts are at serious risk of getting radiation sickness. In addition, the skin of the orbital station modules consists of aluminum, which accumulates secondary radiation. The formation process of new star systems is the result of complex processes that occur in interstellar gas. These processes include nonlinear interactions of turbulence, gravity, collisions of molecular clouds with each other, as well as rotation and several other factors. The evolution of the formation of superdense substance begins from the moment when it gathers in turbulent flows or is formed by supersonic collisions of molecular clouds with each other, or are formed during the intereaction of molecular clouds with supersonic waves, which are formed during the explosions of supernovae, until the moment when these superdense areas reach the prestellar density. Further, depending on several factors, these superdense formations either collapse or disintegrate and return to the interstellar medium. The analysis of observations allows to conclude that a significant part of the molecular clouds is not gravitationally bound. These conclusions are made based on the virial theorem, which expresses the connection between gravitational and kinetic energy. In this paper, a simulation of large molecular clouds collision in a three-dimensional formulation on ultra-high-resolution grids is carried out. Execution of such calculations requires a lot of computer capacity. The paper presents the results of computer modeling of large-scale processes of filament formation and superdense, gravitationally coupled structurally parallel computation clusters with hybrid architecture. Parallel simulation on supercomputers was conducted with author's software, which uses a modified second-order Godunov method of accuracy of the Total Variation Diminishion type. Calculations were conducted on grids that contain more than one billion cells (1024 × 1024 × 1024). The gravitational potential is calculated on graphic processors. To refine the calculations, the method of adaptive refinement of the AMR grid is used. Simulation results are presented for cases of frontal collision of two molecular clouds, which density is distributed along the radius according to certain laws. • The processes of formation of new stellar systems are studied in this work in 3D simulation. • Appearance of clumps and filaments is discussed. • Morphing of high compression clumps set in possible stars formation zone is discussed. • The calculation program is based on a modified scheme of a high order of accuracy. [ABSTRACT FROM AUTHOR]

Chemical ablation and mechanical erosion of hybrid rocket nozzles can cause loss of nozzle thermal protection material, resulting thrust reduction. In this paper, the chemical ablation and mechanical erosion mechanisms of nozzle thermal protection materials are studied through numerical simulation. For chemical ablation, a coupled thermal-mechanical-chemical model is established, considering three mechanisms: wall regression caused by non-homogeneous chemical reactions at the wall, convective heat transfer and fluid-structure interaction. For mechanical erosion, a Fluent DPM model and an Abaqus DEM model are established, which are investigated from the perspectives of empirical formula and solid mechanics mechanisms respectively, and validated against each other. The comprehensive effect of chemical ablation and mechanical erosion on the morphology of the nozzle is considered. The results show that both chemical ablation and mechanical erosion dominate the wall regression rate at the downstream of the converging section and upstream of the throat while the other location is only dominate by chemical ablation. After 5s of operation, thrust of rocket engine decreases by 5.8%. The Abaqus DEM model presented by this paper breaks the limitations of the empirical formulae due to the introduction of the material constitutive model and the failure criterion. • A coupled thermal-mechanical-chemical model is established. • Mechanical erosion is studied from perspective of empirical formula and the solid mechanics mechanism. • Comprehensive effect of chemical ablation and mechanical erosion on nozzle inner wall morphology is studied. [ABSTRACT FROM AUTHOR]

This paper presents an approach to ensuring the safety of a landing module on a threat during its splashdown in the World Ocean. We propose mathematical methods and software to predict the main hydrophysical disturbances from underwater vehicles, which are necessary to estimate artificial anomalies on the sea surface. • Detection of underwater objects ensures safety of the spacecraft splashdown. • Underwater object creates identifiable hydrophysical anomalies. • Computer technologies can predict hydrophysical anomalies. • Analysis of radar images of anomalies allows to detect an underwater object. [ABSTRACT FROM AUTHOR]

Various celestial bodies contain cracks. Also, cracks can be in the skin of aircraft, while their growth is one of the main mechanisms for the breakup of structures. This paper presents a study of one and two rectangular cracks. The influence coefficient (the ratio of the stress intensity factors for the fracture system to the corresponding value for a single crack of the same size) was chosen as the quantity under investigation. The final shape of the crack in the case of its growth is also studied. The following methods were used to model this process: the three-dimensional boundary elements method of discontinuous displacements, in which stress and displacement field is a finite series of expansions of analytically found functions, as well as the two-dimensional boundary elements method of discontinuous displacements, which takes into account the twisting of a growing crack along the stress field at the tip. • Single rectangular crack and system of two parallel cracks were investigated. • For rectangular crack growth is likely to happen from the centers of the sides. • The danger of two parallel cracks depends significantly on their relative position. • Pair of uniaxial cracks is less dangerous then a single crack of the same size. • Non-overlapping cracks in close planes are the most dangerous. [ABSTRACT FROM AUTHOR]

Understanding the formation of impact craters on planetary surfaces has fundamental importance in geology, impact dynamics, and planetary sciences. The finite element-smoothed particle hydrodynamics (FE-SPH) adaptive method combines the advantages of smoothed particle hydrodynamics (SPH) and finite element method (FEM) in solving hypervelocity impact (HVI) problems. Employing FE-SPH adaptive method is a new idea for the analysis of planetary impact cratering which is different from impact-simplified arbitrary Lagrangian Eulerian (iSALE). In this paper, we establish a complete numerical model of planetary impact cratering through the FE-SPH method, including material failure criterion, thermodynamic analysis, gravity preloading, etc. We reproduce the numerical results by Collins et al. (2012) and highly consistent results are obtained through comparative analysis. This study simulates the formation process of simple and complex impact craters, and the temperature, pressure, density, and geometric shape of the impact zone can be quantitatively obtained. We further analyze the distribution of polymorphs and rock fragmentations in the impact zone based on the theories of shock metamorphism. The proposed analysis method may assist geologists in conducting geological studies. • We establish a complete numerical model of planetary impact cratering different from iSALE. • We analyze the distribution of high-pressure polymorphs based on shock metamorphism. • The FE-SPH model has higher numerical precision for planetary cratering impacts. [ABSTRACT FROM AUTHOR]

This paper provides a valuable method to perform transient numerical simulations for hybrid rocket motors (HRMs). A two-dimensional transient numerical simulation model based on Butterworth filter and dynamic mesh is established. The Butterworth filter adopted in this project is a second-order low-pass filter. Then a firing test is arranged to prove the accuracy of the numerical simulation model of the HRMs. This rocket motor adopts 98% hydrogen peroxide (HP) and hydroxyl-terminated polybutadiene (HTPB) fuel containing aluminum as the propellants. The results demonstrate that the Butterworth filter can significantly improve the smoothness of the combustion surface compared with non-filter. And the combustion chamber pressure and thrust of simulation and test results are in good agreement. The average deviation of combustion chamber pressure between the simulation and test is 0.9%, and the average deviation of thrust between the simulation and test is 3.3%, which verifies the feasibility of the numerical simulation model for this HRM. Comparing the regression rate and filtered regression rate along with axial distance over time, it is indicated that the high-frequency oscillations are filtered by the Butterworth filter. In order to compare the port of simulation and test, the combustion chamber is scanned by computed tomography (CT) after the test. The results demonstrate that the simulation and test results are in good agreement, and the average deviation of simulation and test results is 4.5% mainly due to the end-burning of grain and nitrogen blowing. Then, the long-time simulations that adopted the Butterworth filter and non-filter were conducted. The results demonstrate that the use of the Butterworth filter can significantly smooth the combustion surface and the model has the ability to perform long-time simulations. This project can make the simulation for HRMs closer to the test results and provide a valuable guide for HRMs simulation. • A simulation model of hybrid rocket motor based on Butterworth filter is established. • The Butterworth filter is adopted to smooth the combustion surface. • A firing test of hybrid rocket motor with fuel containing aluminum is carried out. • Simulated combustion surface is compared with the computed tomography scan results. • A long-time simulation is conducted to validate model accuracy. [ABSTRACT FROM AUTHOR]

This paper aims to study the detonation re-initiation through multiple reflections when the detonation enters the horizontal channel holding liquid fuel from the vertical branch. Based on the Eulerian-Lagrangian method, the detonation diffraction, reflection and re-initiation of gas phase (C 7 H 16 /Air) and two-phase (n-heptane/Air) are compared, and the effects of droplet diameters on re-initiation and the interaction between shock wave and droplets are analyzed. The results show that compared with the detonation propagating in pure gaseous mixture, detonation is easier to be decoupled in two-phase mixtures during the diffraction stage due to the droplet fragmentation and evaporation process, and lots of unburned fuels are left behind the detonation wave. As the droplet diameter increases, the detonation wave propagates slower and the cell size increases. The number of transverse waves formed after the first reflection gradually decreases, and the transverse detonation in mixture with larger droplets is decoupled during the downstream propagation. The shock wave interacts with the droplet, leading to the occurrence of droplet breakup. At this stage, the total mass of the droplets remains basically unchanged. After that, the droplet evaporation process occurs, and the droplet mass decreases rapidly. • Detonation re-initiation process of gaseous fuel and liquid fuel are compared. • Effect of droplet diameter on detonation re-initiation is studied. • The interaction between droplets and leading shock is illustrated. [ABSTRACT FROM AUTHOR]

In order to efficiently perform the direct numerical simulation for shock wave boundary layer interference, a large grid computing scale is needed, and a single graphics processing unit cannot meet the urgent needs of computing. This paper proposes a parallel algorithm based on the architecture of multiple graphics processing units to effectively deal with the direct numerical simulation for shock wave boundary layer interference. The algorithm achieves a high overlap between graphics processing unit computation and boundary data exchange to improve the parallel efficiency when the grid computing scale is large. The working principle of the multi-graphics-processing-unit parallel algorithm is introduced, and the accuracy and feasibility of the algorithm are verified. Finally, through the application of the algorithm in the research of the compression-expansion slope boundary layer, fine flow structures such as large-scale vortex structure, shocklet, and attached shock wave can be captured. As a conclusion, the multi-graphics-processing-unit parallel algorithm proposed in this paper can accurately display the flow field of shock wave boundary layer interference. • A parallel algorithm based on multiple GPUs is proposed. • The accuracy, feasibility, and efficiency of the parallel algorithm are verified. • The DNS of the SWBLI based on the algorithm is tested and analyzed. [ABSTRACT FROM AUTHOR]

The National Aeronautics and Space Administration (NASA) is currently developing the next generation of spacesuits for use in future exploration missions. This effort, referred to as the Exploration Extravehicular Mobility Unit (xEMU), has been underway since 2015 with the goal of demonstrating the new technologies during a mission to the International Space Station (ISS). Due to the complexity of the xEMU system, and particularly its portable life support system (PLSS), computer simulations are heavily relied on in the development process, from dynamic loads analysis to software verification. The models of the physical systems that are currently being used for system-level simulations are mostly empirical, where the underlying formulas are mathematical fits to test or simulation data. An advantage of this approach is the fact that these models are very performant and can be used in real-time applications such as user interface testing and operator training. The research presented in this paper describes a different approach to modeling and simulation. A simulation tool focused on life support systems for human spaceflight in general and PLSSs in particular has been under development at the Technical University of Munich (TUM) since 2006. The spacesuit specific variant of this tool is called The Virtual Spacesuit (V-SUIT) and is now being utilized in the current spacesuit development process at the NASA Johnson Space Center (JSC). In contrast to the previously mentioned empirical models, V-SUIT uses a first-principles-based, bottom up modeling approach, where basic physical, chemical or biological effects are modeled at the lowest level and the component and system models are created using these fundamental building blocks. This enables the simulation to account for effects, and especially off-nominal situations, that may not be within the range of validity of the empirical models. This paper presents the structure of and the rationale behind V-SUIT itself as well as a status report on the ongoing effort to integrate it with the existing simulation systems in use at JSC, which is based on Trick/GUNNS. As a proof of concept, a single component model from V-SUIT, that of the spacesuit water membrane evaporator (SWME), has been integrated with NASA's simulation tool. Results include comparisons of the performance of the two simulation systems and a description of challenges that were encountered during implementation. The paper closes with an outlook towards the future developments that are planned for V-SUIT within the xEMU development program. • A novel approach to dynamic system-level simulations is applied to a spacesuit. • Multi-disciplinary, bottom-up modeling instead of domain-specific, top-down models. • Performance shown to be fast enough to provide value during system development. [ABSTRACT FROM AUTHOR]

In this paper, the attitude control of spacecraft using a pyramid-type variable-speed control moment gyro (VSCMG) system is considered. For successive attitude maneuvers, it is desirable that the wheel angular momentums are regulated to nominal values after an attitude maneuver. In addition, avoidance of reference aligned singularity (RAS) is also important. RAS is a singularity point where the singular direction of the constant-speed control moment gyro (CSCMG) system is aligned to a reference torque. Regarding the motivations, this paper proposes a steering law which achieves attitude control and regulates the wheel angular momentums while successfully avoiding RAS. Both numerical simulations and experimental verification are performed to demonstrate not only the validity of the proposed method, but also a certain robustness against unmodeled dynamics and the gravity torque. • Spacecraft attitude is controlled by variable speed control moment gyros. • Reference aligned singularities of CMGs are particularly focused in the control. • Wheel angular momentums are usually regulated to nominal values. • Effectiveness of the control law is validated in all maneuver directions. • Ground experiments are conducted in order to examine the proposed control. [ABSTRACT FROM AUTHOR]

Force control provided by a small-scaled free-flying space robot on a floating or spinning target will be required for future on-orbit servicing missions, such as detumbling a target, performing an assembly or a maintenance operation. Regarding to different contact geometries for a space robot interacting with a spinning target, this paper addresses the contact model considering its multi-dimensional characteristics. A hybrid motion and force controller is developed to acquire desired contact forces and space robot configuration despite the floating feature of the system. On this basis, the control strategy to implement the operations of three typical phases, namely approaching phase, contact phase and post-contact phase, is proposed. The simulations have been performed to verify the control approaches and visualize the contact scenarios. • Multi-dimensional contact models regarding to two contact geometries are formulated. • Hybrid motion and force controllers for two contact scenarios are developed. • A novel control strategy for typical on-orbit servicing phases is proposed. • Numerical simulations based on Simscape Multibody verify the approaches. [ABSTRACT FROM AUTHOR]

The paper continues the numerical investigation for the problem of hypervelocity impact of a spherical projectile against thin-walled metallic containment that is a key process in the space vehicles shielding by composite honeycombs. The features of initial outward penetration process are studied for wide variation of projectile and containment material properties. The statements of the impact problem presented in the paper involve triple contact points. • Numerical investigation for the problem of hypervelocity impact. • Spherical projectile against thin-walled metallic containment made of different materials. • Space vehicles shielding by composite honeycombs concept. • The features of initial outward penetration process are studied for wide variation of material properties. • The statements of the impact problem involve triple contact points. [ABSTRACT FROM AUTHOR]

The paper deals with the attitude dynamics of CubeSats with flexible stabilizing panels in free molecular flow taking into account the aerodynamic damping. In addition to the operating position, characterized by zero angle of attack, aerodynamically stabilized satellites may have intermediate equilibrium positions. The presence of unstable equilibrium positions and small perturbations such as the oscillations of the flexible panels is the cause chaos in the attitude motion. An analysis of the chaotic motion is carried out using Poincare sections and Lyapunov exponents. Numerical simulations show that the chaos intensity is sensitive to the geometric and environmental parameters of the system. • The attitude motion of the CubeSats with aerodynamic instability is considered. • The chaotic behavior of the satellite with flexible panels is proven. • The chaos intensity is sensitive to the geometric and environmental parameters of the system. [ABSTRACT FROM AUTHOR]

This paper addresses the problem of powered descent control of a lunar lander during its final landing phase, where soft and precise landing at the target cite is required. In the literature, this problem has been solved considering a fully actuated lunar lander. Recently, landers are designed as complex systems with several onboard payloads and subsystems for completing demanding mission tasks. There are increasing constraints on mass budget and space availability onboard a lander. In order to meet these requirements, a novel controller using a single thruster is proposed. The proposed controller has three control inputs (thrust and two gimbaled angles) to control the six degrees of motion of a lander; these control inputs are non-affine. The proposed controller is designed based on variable structure control technique augmented with a high order filter, an immersion and invariance-based mass estimator and a sliding mode angular velocity observer. The stability analysis using Lyapunov theory and the results of the numerical simulations along with Monte Carlo simulations of the system show that the precise landing of the lunar lander using a single thruster is feasible. • The position and attitude control of a lunar lander using a single thruster is examined. • A novel controller along with lander mass and angular velocity observers are designed. • The proposed controller is based on variable structure control technique. • The Lyapunov theory is applied to prove the stability of the proposed control scheme. • Results of numerical simulations show the feasibility of the proposed control scheme. [ABSTRACT FROM AUTHOR]

This paper is concerned with the transfer trajectory design among multiple planets using multiple flybys. The proposed trajectory design method consists of a semi-analytical tuning strategy and iterative patched-conic method. The tuning strategy refers to a method realizing the smooth connection between a hyperbolic trajectory near a planet and a trajectory obtained by Lambert's problem between the planet and another one. The proposed method iteratively solves Lambert's problem and the tuning strategy so that the arrival or departure velocity obtained by Lambert's problem coincides with the velocity at the corresponding point in the sphere of influence (SOI), assigned by the tuning strategy. Consequently, a smooth and continuous trajectory with multiple flybys can be obtained, where the SOI of each flyby planet and the flyby time are explicitly taken into account. Moreover, analytical conditions with respect to the orbital elements of the arrival and departure hyperbolas are provided to minimize velocity mismatch at the periapsis of a flyby planet. Through numerical simulation using the Cassini example, it is shown that the resultant trajectory by the proposed method is close to the optimal trajectory obtained by the n -body problem. Thus, a suitable approximate trajectory can be effectively generated without solving the n -body problem. • Interplanetary trajectories with multiple flybys are designed. • The proposed method consists of analytical tuning strategy and iterative patched-conic. • The SOI and flyby time for each flyby planet are explicitly taken into account. • Analytical conditions of the orbital elements are provided to minimize velocity mismatch. • A suitable approximate trajectory is effectively generated without solving the N-body problem. [ABSTRACT FROM AUTHOR]

The problem of task-space tracking control of a free-floating space robot is addressed in this paper. Based on a fixed-time extended state observer (FXESO), a fixed-time position and attitude control (FXPAC) method is developed. The attitude of the end-effector is represented in modified Rodrigues parameters (MRPs). The FXESO provides the estimation of the joint velocity and the lumped disturbance for the control system. Based on the backstepping technique and a power integrator, the FXPAC method is designed. This control scheme can guarantee tracking errors converging to a neighborhood of the origin within a fixed time in the presence of the external disturbance. Moreover, since the inverse of the generalized Jacobian matrix is involved in the controller, it is necessary to consider dynamic singularities. To handle these singularities, a method of combining the singularity separation and damped reciprocal (SSDR) is applied to the space robot. Numerical simulations are presented to demonstrate the effectiveness of the proposed method. • Fixed-time tracking control in task space is proposed for space robots. • Fixed-time extended state observer is designed to obtain velocities and disturbance. • Dynamic singularities are handled during the control design. [ABSTRACT FROM AUTHOR]

To overcome the singularity problem of control moment gyros (CMGs) and simultaneously provide a lightweight, agile, and simple dynamical system, a CMG system called a double-gimbal scissored-pair control moment gyro (DGSPCMG) was recently proposed. The DGSPCMG is a hybrid mechanism that combines scissored-pair CMGs and a double-gimbal mechanism; further, the outer singularities in this system (saturation singularities) can be escaped by steering the scissored-pair gimbals. This system generates less perturbation torque and has almost no inner singularities except at the origin and the line along the outer gimbal axis. This paper proposes model predictive control (MPC) based steering laws for the mentioned CMG system. The proposed MPC based steering laws can directly provide the optimal gimbal rate under the nonlinear constraints in real time without using the inverse of the CMG Jacobian matrix, and hence, the singularities do not have to be considered when determining the gimbal steering rate of this system. Three objective functions are considered in the steering law design: fast maneuver-weighted type, minimum control energy type, and a hybrid of these two types. The control performances of these three MPC-based steering laws are compared via numerical simulations. Results demonstrate that the hybrid type can achieve both a short rise time of the angular velocity at the initial stage of control, and smooth stabilization to the target attitude; moreover, its settling time and control effort is between those of the other two types. • Double-gimbal scissored-pair control-moment gyros (DGSPCMG) is considered. • Model-predictive control (MPC) based steering laws are designed for this system. • Fast maneuver, minimum control energy types, and hybrid of these two are considered. • Hybrid type achieved a short rise time of angular velocity and smooth stabilization. • As setting a long receding horizon length, the settling time was shortened. [ABSTRACT FROM AUTHOR]

In this paper, supersonic unsteady flows around the spiked cylinder are investigated through two-dimensional (2D) and three-dimensional (3D) unsteady Reynolds-averaged Navier–Stokes (URANS) approaches. Three types of spikes are used to provoke pulsating/oscillating flow in front of the cylinder. Numerical simulations of the flow past the cylinder with a pointed spike are validated against the experimental data of Feszty et al. (2004). Results indicate that the pressure imbalance in the dead-air region can be significantly reduced by increasing the aerodome size, which represents an effective method for controlling periodic flow pulsations. The neglect of real 3D effects in the 2D axisymmetric hypothesis tends to alter the amplitude and frequency of the pressure history, and even overestimate the critical aerodome size at which the flow pattern transitions from violent pulsations to mild oscillations. The discrepancy between the 2D axisymmetric and 3D solutions becomes smaller after activation of pulsation control. Through a comparative study, it is observed that in contrast to the case of zero angle of attack, the instantaneous flow field of the non-zero angle of attack case becomes increasingly symmetrical. • Large aerodome facilitates the transition of flow mode from pulsation to oscillation. • Pulsations can be explained by the alternating dominant of two high-pressure gases. • Ignoring the real 3D effects triggers the over-prediction of critical aerodome size. • The use of 2D axisymmetric hypothesis is feasible for the prediction of oscillations. [ABSTRACT FROM AUTHOR]

In this paper, the evolution of a cloud of interacting particles in their orbital motion around the Earth is considered. In numerical simulations in a non-inertial orbital coordinate system, there is a large difference in the orders of the quantities used, which prevents calculations. To prevent this from happening, the equations of motion of particles are written in a special form. The proposed research algorithm allows one to obtain the relative configurations of the clouds of particles with allowance for various perturbations and with different required computational accuracy. • External influences on the space debris area are studied. • Large difference in the orders of used quantities in simulations is discussed. • Evolution of space debris of LEO satellites disposal zone is considered. [ABSTRACT FROM AUTHOR]

This paper investigates the control issue of the noncooperative spacecraft proximity with prescribed performance by exploiting only relative motion information. To accommodate the direct use of relative measurement, the 6-degree-of-freedom (6DOF) relative dynamics is formulated in the body frame of the pursuer. Based on the smooth Gauss error function, a new asymmetric saturation function is developed to satisfy the practical control limitations. To maintain the relative attitude and position within specified constrained ranges, a novel performance function in possession of the preassigned settling time and convergence set of steady state error is proposed. Based on the backstepping framework, the relative attitude and position controllers are surveyed in sequence. In particular, a state constrained virtual control is first developed for each relative dynamics by introducing a nonlinear transformation. Then a neuroadaptive saturated control law is designed by implementing a Nussbaum function based adaptive control coefficient. Stability analysis indicates that the closed-loop 6DOF relative dynamics are asymptotically stable with prescribed performance. Numerical simulations further verify the proposed control strategy. • A control strategy is developed by using only relative motion information. • A novel performance function with accurate and intuitive settling time is proposed. • An asymmetric saturation function is designed to satisfy the control constraints. • A neuroadaptive algorithm is introduced to ensure the high control accuracy. [ABSTRACT FROM AUTHOR]

The Earth orbit is in a serious predicament caused by millions of space debris, and operational spacecraft are facing a risk of collision with those objects. The net capturing method has been considered as one of the most promising method to remove space debris objects. However, a tumbling target has a risk to slide toward the service satellite and collide with it after captured. To avoid this from happening, a control method must be designed for the combined system. This paper first presents the necessity of controlling the combined system right after the net capture. In another respect, the existing configurations of the combined system always include an additional mechanism to de-tumble the target which will dramatically complicate the design. Instead of using an additional mechanism, a novel control strategy is proposed to de-tumble the target only via the tether connecting to the service satellite and the target. Dynamics of the system is analyzed and flexibility of the tether is considered by employing the mass–spring–damper model. The control method includes three control modes to actively control the pose of the service satellite which significantly impacts the tether tension so that the target can be maintained a safe distance to the service satellite. Numerical simulations demonstrate that the proposed control strategy can be utilized to effectively de-tumble the target within a short period of time and mitigate the risk of collision with the service satellite. • Post-capture dynamic model of a chaser–tether–target combined system is formulated. • A novel control strategy is proposed to de-tumble a target only via tethers. • Three cases demonstrate successful detumbling of a non-cooperative target. [ABSTRACT FROM AUTHOR]

A small-scale experimental burner with an improved rotary valve has been designed to measure solid-propellant pressure-coupled response for mean pressures on the order of 2.2–2.9 MPa and oscillating frequencies of 23–161 Hz, within 9% amplitudes. This paper proposes an indirect method for measuring the instantaneous valve area and phase delay angle based on a circular grating system and a corresponding mathematical model, and the validity of which is confirmed by several cold-gas experiments and numerical simulations. The effects of frequency on oscillating pressure, exhaust mass flow rate and pressure-coupled response are studied by carrying out several hot-gas experiments and numerical simulations. The results show that the mean pressure is not affected by the oscillating frequency and maintains at 2.9 MPa. The exhaust mass flow rate of the rotary valve is determined by the instantaneous valve area and independent of frequency. The peak-to-peak pressure is inversely proportional to the oscillating frequency, reducing from 0.539 MPa at 23 Hz to 0.053 MPa at 161 Hz. The magnitude of pressure-coupled response appears to be constant (0.8) at 23–161 Hz, with pressure oscillations lagging instantaneous valve area by 0.504π-0.531π。The study is expected to benefit research on similar rotary valves and can serve as a basis for measuring the solid-propellant pressure combustion response. • Design of a small-scale experimental burner for measuring solid-propellant pressure-coupled response. • Three-dimensional internal transient flow field analysis of an improved rotary valve. • Effects of oscillating frequency on oscillating pressure, exhaust mass flow rate and pressure-coupled response. [ABSTRACT FROM AUTHOR]

This paper investigates the finite-time attitude tracking control problem for rigid spacecraft exposed to the external disturbance and unknown inertial parameters. Initially, the basic controller is developed under the assumption that the inertial parameters and the upper bound of external disturbances are available to the designer. However, this assumption is inapplicable in practical applications. Considering this limitation, adaptive laws and Minimum-learning- parameter (MLP) algorithm are employed in the second controller to estimate the unknown system dynamics. Finite-time stability for tracking errors is achievable under the proposed methods. Though theoretical analysis and numerical simulations, effectiveness of the proposed controllers has been validated. • Two finite-time attitude tracking controllers are proposed. • The proposed controller is inertia-free for implementation. • MLP Algorithm has been employed to reduce the computational burden. [ABSTRACT FROM AUTHOR]

The paper describes an analytical and experimental framework which investigates performance of electrically-driven ground vehicles powered by battery packs, with the objective of providing simple indications for preliminary sizing of rovers designed for mission on the surface of extraterrestrial bodies. For this class of missions, power is at a premium, and energy-efficient locomotion is clearly a critical issue. An analytical model for the estimate of cruise distance is derived and mobility performance is thus analyzed as a function of cruise speed, relevant vehicle parameters, soil mechanical properties, and atmospheric data. Results are then validated by numerical simulations and an experimental campaign for an Earth-based rover. • Performance of a battery-powered extraterrestrial rover is discussed. • A novel battery-discharge model is used to derive a closed-form expression for cruise range. • Best range speed is determined for different vehicle/soil relevant parameters. • The approach is validated by numerical simulation and an experimental campaign. • The procedure has general validity and guidelines are provided for preliminary design of EVs. [ABSTRACT FROM AUTHOR]

Before an active debris removal process is carried out, the attitude stabilization of the debris needs to be executed to ensure security. Fortunately, for the tethered system, a full despun is not always necessary; a stabilization of the motion of the tether attachment point is usually sufficient. This paper discusses the way to stabilize the motion of the tether attachment point of a captured noncooperative space debris without any auxiliary mechanisms. The debris possesses unknown inertia information and states after capture, while the tether tension is the only input to manipulate its attitude. The direction and value of the tension is regulated based on the continuous dissipation of the rotational energy of the debris. Moreover, the range of the tether direction is restricted to a small conical region, in order to reduce the tether libration motion and the orbit maneuver of the main satellite. The motion of the tether attachment point is desired to be observed instead of measuring the attitude motion of the debris in the feedback loop, which might be easier to attain in practical applications. Numerical simulations are carried out to validate the effectiveness of the control method and to check the performance of the controller with different parameters.] • The motion of the tether attachment point is stabilized by the tether tension. • No inertia information concerning the debris or the measure of its attitude motion is needed. • The range of the tension direction and convergence rate can be adjusted by choosing proper control parameters. • Different equilibrium states are discussed by theoretical analyses. [ABSTRACT FROM AUTHOR]

Ground-based optical tracking systems used in space surveillance provide no range information, and an observed arc is too short compared to the orbital period to provide enough dynamical information, especially for objects in low Earth orbit. Thus, multiple arcs of tracking data should be accumulated and associated in case of long-range orbit prediction. In this paper, a strategy to associate multiple orbit solutions and obtain accurate orbit solution to predict long-term trajectory using an optical space surveillance system is proposed and analyzed. The strategy consists of three steps: unscented batch estimation, chi-square testing, and element fitting. Each step is validated by numerical simulations, and the practical applicability is verified by applying actual tracking data from the Optical Wide-field Patrol Network (OWL-Net). In both the simulations and the practical cases, the associations are well determined, and the post-fit orbit solution provides a predicted trajectory error of 20 km for a week. The proposed strategy is a standalone process requiring no additional observations or orbit database, and it supports successive tracking by optical equipment with a restricted field of view. • A strategy to confirm the correlation of multiple orbit solutions is proposed. • The strategy consists of estimation, correlation testing, and element-fitting. • Each step is validated by numerical simulations. • The proposed process is applied to actual tracking data. • The proposed strategy supports successive tracking by optical system. [ABSTRACT FROM AUTHOR]

This paper presents the design and characterization of a miniature docking mechanism for nanosatellites. Potential applications are several, including servicing of orbital vehicles (e.g. refuelling, components replacement, deorbiting or reboosting) and assembly of large structures (e.g. telescopes, antennas). The mechanism responds to the constant demand of enabling technologies from the booming small satellites market. The developed system has a traditional probe–drogue configuration; it is equipped with a sensor to detect contact and a single servo-actuator to lock the connection. The simple, though effective, design fills a relevant gap in the field of nanosatellite technologies. Numerical simulations have been conducted to evaluate the dynamics of docking procedures and to estimate loads exchanged at contact. Experimental results validate the simulations and prove the high tolerance to angular and linear (lateral) misalignment. • Miniature docking mechanism based on probe–drogue configuration. • Compatibility with the CubeSat standard and extreme simplicity. • Possibility to equip prepared, passive orbital objects. • Numerical simulations of docking dynamics. • Extensive laboratory validation of full docking procedure. [ABSTRACT FROM AUTHOR]

Mission analysis using aerogravity-assist (AGA) maneuver at Titan is presented in this paper. Considering Saturn arrival, Titan encounter, and different post-AGA orbits, the analysis uses vector diagrams to illustrate the relations among all conditions as well as design constraints. The analysis focuses on Enceladus mission concept and also present a summary result for Cassini-type Saturn system missions. Turn angles using AGA maneuver at Titan are quantified via numerical simulation, showing that 25-deg AGA turn angles are achievable with a Titan inbound excess velocity (V ∞ , IN ) of 10–20 km/s. Results show that for an Enceladus mission, V ∞ , IN of 10 km/s is the minimum velocity required for AGA maneuver to achieve an outbound transfer orbit directly to Enceladus; the corresponding peak g-load is less than 6 g, peak heat rate less than 200 W/cm2, and total heat load of around 25 kJ/cm2. For a Cassini-type Saturn mission, a minimum V ∞ , IN of 5 km/s can be used for Titan AGA. The results showed are for V ∞ , IN of 6 km/s; and the corresponding peak g-load, peak heat rate, and total heat load are 1.5 g, 20 W/cm2, and 5 kJ/cm2 respectively. • Design space for Saturn system missions using Titan aerogravity-assist is fully explored. • A novel method combining graphic approach using vector diagrams and numerical simulation is presented. • Readiness of entry vehicles and TPS technologies for using Titan aerogravity-assist maneuver is evaluated. [ABSTRACT FROM AUTHOR]

A parameter identification method using interval analysis algorithm for structural dynamics is presented in this paper. The proposed uncertain identification method is investigated by using central difference method and ARMA system. With the help of the fixed memory least square method and matrix inverse lemma, a set-membership identification technology is applied to obtain the best estimation of the identified parameters in a tight and accurate region. To overcome the lack of insufficient statistical description of the uncertain parameters, this paper treats uncertainties as non-probabilistic intervals. As long as we know the bounds of uncertainties, this algorithm can obtain not only the center estimations of parameters, but also the bounds of errors. To improve the efficiency of the proposed method, a time-saving algorithm is presented by recursive formula. At last, to verify the accuracy of the proposed method, two numerical examples are applied and evaluated by three identification criteria respectively. [ABSTRACT FROM AUTHOR]

This paper develops a new approach to solve the reachable domain of a spacecraft with a single maximum available impulse. First, the distance in a chosen direction, started from a given position on the initial orbit, is formulated. Then, its extreme value is solved to obtain the maximum reachable distance in this direction. The envelop of the reachable domain in three-dimensional space is determined by solving the maximum reachable distance in all directions. Four scenarios are analyzed, including three typical scenarios (either the maneuver position or impulse direction is fixed, or both are arbitrary) and a new extended scenario (the maneuver position is restricted to an interval and the impulse direction is arbitrary). Moreover, the symmetry and the boundedness of the reachable domain are discussed in detail. The former is helpful to reduce the numerical computation, while the latter decides the maximum eccentricity of the initial orbit for a maximum available impulse. The numerical simulations verify the effectiveness of the proposed method for solving the reachable domain in all four scenarios. Especially, the reachable domain with a highly elliptical initial orbit can be determined successfully, which remains unsolved in the existing papers. [ABSTRACT FROM AUTHOR]

This paper proposes an adaptive control scheme for hypersonic vehicle with external disturbances, parametric uncertainties, actuator failures, input saturation and state constraints. Firstly, the longitudinal model of generic hypersonic vehicle is divided into velocity subsystem and altitude subsystem. Secondly, based on the tangent barrier Lyapunov function and auxiliary system, two adaptive fault tolerant controllers are designed for velocity subsystem and altitude subsystem, respectively, which can solve the problem of input saturation. By using the controllers, the flight states containing the angle of attack, flight path angle, pitch angle and pitch angle rate can be guaranteed in prospective ranges. Finally, the Lyapunov stability theory and numerical simulations are adopted to verify the effectiveness of the presented controllers. • The prescribed tracking performances are guaranteed. • State constraints of the system are guaranteed. • The input saturation and external disturbance are solved. [ABSTRACT FROM AUTHOR]

The electron cyclotron resonance (ECR) neutralizer is a component of the 10-cm electron cyclotron resonance ion thruster (ECRIT) and its performance has a great effect on the overall performance of the thruster. To clarify the mechanisms of electron extraction of the ECR neutralizer, a particle-in-cell/Monte Carlo collision (PIC/MCC) solver was developed to simulate the plasma in the ECR neutralizer. This paper equivalently simplified the three-dimensional model of the ECR neutralizer into a two-dimensional axisymmetric one according to experimental results and the circumferential electric drift of charged particles. Hence, a two-dimensional three-velocity component full-particle calculation model was used in the solver. Through analyzing ECR heating mechanism, the validity and accuracy of the equivalent ring antenna were discussed. The distribution of plasma parameters and extracted electron current were obtained by tracking the motion of electrons and singly charged xenon ions, simulating the elasticity, excitation, ionization and Coulomb collisions. The results showed that the density of neutral particles can be adjusted to coincide simulation current with experiment. It was also found that due to the magnetic field gradient, plasma will migrate to the magnetic well, where the ion was in dominant position to form a potential well and hinder the extraction of electron. • Based on experiment, an equivalently 2D axisymmetric full PIC/MCC model was derived. • The model can simulate the plasma property of neutralizer and save computation time. • The simulation reveals the electron extracting mechanism of the ECR neutralizer. [ABSTRACT FROM AUTHOR]

In this paper, a new hybrid method based on invariant manifold and chaos control is presented to design Earth-Moon low-energy transfer trajectory in the planar circular restricted three-body problem. In the proposed method, one segment trajectory emanates from the period orbit around the Earth to Lyapunov orbit associated with the Lagrange points L 1 , which is designed by chaos control method to reduce the transfer time wandering in the chaotic region. Another segment starts from the Lyapunov orbit and reaches the Moon orbit, which is designed by the invariant manifold method. The two segments are patched with varying Jacobi constant by utilizing a Poincaré section, which is determined by the pericenter relative to the Earth. Several candidate low-energy transfer trajectories can be obtained by the proposed method, among which a suitable trajectory could be selected via the trade-off between fuel and flight time. Extensive numerical simulations are implemented to demonstrate the effectiveness of the proposed method. • Development of a hybrid method by combining invariant manifold and chaos control. • Low-energy transfer can be designed via the trade-off between fuel and flight time. • Comparison of the proposed method with invariant manifold and chaos control method. [ABSTRACT FROM AUTHOR]

For the visibility optimization problem of satellite constellations, this paper presents an efficient hybrid optimization method taking multiple constraints into account. To reduce the time needed for the optimization, the satellite constellation is assumed to have a specific configuration. The optimization includes two steps. The first step searches for optima over a large scale, while the second step is designed to obtain the accurate optimal solution based on the result of the first-step optimization. Based on the analysis of the coverage performance function, a linear adaptive population method is proposed for the first step. When compared with the traditional optimization method, calculation time can be efficiently reduced while ensuring an accurate optimization. Several numerical simulations are conducted for validation. It proves to be a practical choice for engineering applications. • Different types of satellite load sensors and different coverage performance indices are considered. • Constraints of double- or more-level coverage are discussed. • The trend of the number of local maximum points of one-level coverage to the ground point are analyzed. • A novel hybrid two-step optimization method is proposed. [ABSTRACT FROM AUTHOR]

This paper focuses on the powered-descent guidance (PDG) problem involving uncertain-but-bounded parameters. An efficient numerical algorithm based on the high approximation accuracy of the Chebyshev series expansion is presented to solve the nonlinear optimal control problem with interval uncertain parameters. First, an infinite number of uncertain optimal control problems is translated into a series of deterministic optimal control problems by exploiting the Chebyshev interval inclusion. Subsequently, the deterministic optimal control problem is solved by a convex optimization method based on the lossless convexification of the PDG problem. Afterwards, these techniques are used to generate the enclosure of the PDG trajectory with uncertainty and the upper and lower bounds of the minimum fuel consumption during the pinpoint landing mission. Finally, the effectiveness and superiority of the proposed approach are validated by illustrative numerical simulations. • The proposed model accounts for uncertain mission parameters. • Only the upper and lower bounds of the uncertainty parameters are needed. • The enclosures of the trajectory and fuel consumption are generated. [ABSTRACT FROM AUTHOR]

The homotopy methods have long served as powerful tools in solving nonlinear optimal control problems, particularly for which the solutions are highly sensitive to the unknown initial conditions. The principle of homotopy methods is that a homotopic parameter is embedded into the formulations of the optimal control problems, and the original problem can be solved by tracing the optimal solutions of the embedded problems. The existing homotopy methods typically introduce the homotopic parameter into the time variable, the necessary conditions, the performance index or the rightside of the differential equations. In this paper, a new fractional homotopy method is presented, the homotopic parameter of which is embedded into the derivative of the differential equations. By using the proposed method, the optimal solution of the target homotopy problem can be found by solving a series of fractional two-point-boundary-value-problems. Numerical demonstrations in a nonlinear optimal control problem and a three-dimensional minimum-time low-thrust orbital transfer problem are presented to illustrate the applications of the method. • A new fractional homotopy method to solve optimal control is proposed. • The homotopic parameter is embedded into the derivative. • A continuous zero curve can be constructed. • Effectiveness of the proposed method is demonstrated by numerical simulations on 2 examples. [ABSTRACT FROM AUTHOR]

This paper seeks to provide a probabilistic control framework, named Probabilistic Multiple-Horizon Multiple-Model Predictive Control, for the soft and precise landing on an asteroid. The modified version of the Predictive Path Planning method is also introduced to generate a safe and smooth reference trajectory for landing. The Probabilistic Multiple-Horizon Multiple-Model Predictive Control is carried out in two phases; the offline phase to calculate the difference between models, and the online phase to use a probabilistic approach to track the reference trajectory. The difference of the dynamical models is utilized in the online control method to compensate the accuracy of the approximate models in the Multiple-Horizon Multiple-Model Predictive Control method. Furthermore, the presented control framework has the model-updating characteristic that helps to modify the prediction according to past information. The numerical simulation results demonstrate the effectiveness of the proposed guidance and control approach in the reduction of the fuel consumption as well as the landing error in the presence of different sources of uncertainties and disturbances. • A heuristic path planning is developed to design a smooth and safe trajectory. • The guidance law guarantees the safety, accuracy, and vertical landing condition. • A robust model predictive control is designed to reduce the uncertainties' effects. • The control method benefits from low computational time, efficiency, and accuracy. [ABSTRACT FROM AUTHOR]

This paper aims to quantify the responsiveness of two on-orbit servicing infrastructures providing services to multiple serviceable platforms in coplanar medium Earth orbit (MEO) and geostationary orbit (GEO). The platforms to be serviced are assumed to be made of elementary units (EUs). EUs are small standardized structural units capable of aggregating with each other and gathering the key functions of a typical satellite within the size of a 6U cubesat. The first servicing infrastructure considered in this paper, called “ Without Depot” (WoD), includes a launch vehicle and a robotic servicer. The second servicing infrastructure, called “ With Depot” (WD) includes a launch vehicle, a robotic servicer and an orbital depot of EUs. The responsiveness of these infrastructures is quantified using a queueing theory-based stochastic simulation on Simulink taking into account the launch and platform random failures. The metrics used to quantify the responsiveness are the service completion rate and the average waiting time before an EU is replaced over a 10-year period of operation. With respect to those metrics, it is demonstrated that WD is more responsive than WoD. However, WD is also shown to be likely to cost more than WoD. Finally, the sensitivity of the responsiveness of WD to the capacity of the orbital depot is analyzed to find the minimum storage capacity for which WD is the most responsive. This last result could be of interest for space designers dealing with the sizing of on-orbit servicing infrastructures. We believe that the concept introduced in this paper will be a critical milestone in the design of a responsive integrated space infrastructure dedicated to the development and prosperity of a new MEO/GEO economy. [ABSTRACT FROM AUTHOR]

Electrodynamic tethers represent one of the possible means to de-orbit defunct satellites from Low-Earth-Orbit at end of life. However, tethers survivability to orbital debris impacts is still debated because of the large area they expose to the space environment. Recently, increasing consideration has been given to thin-tape tether geometries, whose response to space debris threat is believed to be better than that of round wires. This paper describes a new Ballistic Limit Equation (BLE) applicable for thin tapes, to go beyond previous investigations referring at most to the impact resistance of round-wires (furthermore neglecting the damage dependence from the impact velocity and angle). In this paper, a new approach for BLE derivation is presented, which combines experimental results (in total 24 impact tests) and numerical simulations (in total 112 runs). The resulting BLE is non-monotonic with respect to the impact angle and presents a minimum at certain values of the impact obliquity, depending from the debris size and speed. In other words, the minimum particle diameter which is just able to cut a tape at a given velocity decreases with increasing impact obliquity up to a certain angle above which the damage is reduced due to early debris fragmentation triggered by shock waves propagating into the material. Notably, it has been observed that there is a minimum value of debris velocity v* below which no critical damage is possible and, furthermore, there is a minimum velocity-dependent value d* of debris diameter below which no critical damage is possible. This feature of BLE is extremely important, since it sets a minimum particle diameter for risk assessment and thus excludes a large part of the flux from risk computations. In conclusion, the newly-developed BLE confirms that thin tapes are significantly more resistant than round wires of equivalent cross-section; this is due to the intrinsic ballistic response of tapes, not only to their reduced cross-section at high impact obliquity. [ABSTRACT FROM AUTHOR]

This paper proposes a new mission concept devoted to the identification and tracking of space debris through observations made by multiple spacecraft. Specifically, a formation of spacecraft has been designed taking into account the characteristics and requirements of the utilized optical sensors as well as the constraints imposed by sun illumination and visibility conditions. The debris observations are then shared among the team of spacecraft, and processed onboard of a “hosting leader” to estimate the debris motion by means of Kalman filtering techniques. The primary contribution of this paper resides on the application of a distributed coordination architecture, which provides an autonomous and robust ability to dynamically form spacecraft teams once the target has been detected, and to dynamically build a processing network for the orbit determination of space debris. The team performance, in terms of accuracy, readiness and number of the detected objects, is discussed through numerical simulations. [ABSTRACT FROM AUTHOR]

This paper focuses on the complex dynamics and control of deploying a payload using a long flexible tether from a spacecraft. Two different system models are involved, one for detailed dynamics study and the other for control design. Both the flexibility and extensibility of tether and the attitude motion of spacecraft have been incorporated into the system dynamics. In addition, the orbital dynamics of the tethered system are modeled using a set of modified equinoctial elements and considering the J 2 perturbation forces. A two-stage mission design has been proposed to divide the deployment into two parts. The first is an ejection separation stage for the payload to be ejected away from the spacecraft to a predetermined position, and the second is a deployment control stage for active tether deployment by regulating tether tension. The attitude motion control of the spacecraft is performed during deployment based on an optimal fuel control method. A closed-loop feedback controller is developed to achieve the deployment by tracking an open-loop optimal trajectory of tether libration states and control input. Finally, numerical simulations have been performed to demonstrate the feasibility of conceptual mission design and effectiveness of the synthetic control scheme. • A two-stage mission design has been proposed. • Spacecraft attitude, tether flexibility and extensibility are accounted for. • Attitude stabilization and tether deployment are simultaneously achieved. [ABSTRACT FROM AUTHOR]

As a preliminary of an active debris removal process, the attitude motion control of the debris need to be carried out to ensure security of following operations. The tethered space system is one of the most attractive methods for active debris removal missions. However, the tether tension becomes the only input to manipulate the attitude motion of the debris, causing a highly underactuated problem. Fortunately, the fully attitude stabilization of the debris is not a prerequisite for a secure deorbit operation. As long as the nutation of the debris is not so violent that makes the tether winds, the towing process can be applied continually, which is a significant advantage of the tether method. In this paper, a switched tension control law is presented to damp the nutation of the debris. Only two constant tension values are needed in the control method. Under the assumption that the tether direction is fixed in the inertia frame, the bounded stability of the control algorithm is proved based on the Barbalat's Lemma, and an asymptotic stability can be achieved for an asymmetrical debris symmetrically captured. Numerical simulations with tether libration considered are carried out to show the capability of the control method. • The nutation of the space debris is damped by the tether tension only. • No inertia information of the debris or measure of its attitude motion is needed. • Only two constant tension values are needed and the switching criterion is simple. • An asymptotic stability can be achieved for an asymmetrical debris symmetrically captured. • The ability that only use the tether tension to achieve complete spin counteraction is explored. [ABSTRACT FROM AUTHOR]

Abstract Finding relative orbits for satellite formations flying around the Earth, which are long-term bounded under various perturbations, has been a vibrant field of study. However, much less attention has been given to detecting such orbits for circumlunar formation flying missions. As opposed to low-Earth orbits, in circumlunar missions the third-body effect is large, and the magnitude of the C 22 sectorial harmonic has the same order as the J 2 zonal harmonic. This renders the analysis of bounded relative orbits more challenging. In this paper, we detect a new family of long-term bounded lunar relative orbits, which can be used for lunar formation flying missions. The main idea is to find an analytical solution for the mean relative distance among the formation satellites, and use it to derive formation geometries that are resilient to the gravitational and third-body perturbations. Two invariant mean-distance conditions and one bounded mean-distance conditions are derived. Numerical simulations indicate that the newly-derived invariance conditions yield long-term bounded relative motion. Highlights • A new family of long-term bounded circumlunar relative orbits is found. • An analytical solution for the mean relative distance among the formation satellites is obtained. • Two invariant mean-distance conditions and one bounded mean-distance conditions are derived. • Numerical simulations indicate that the satellite relative motion stays long-term bounded. [ABSTRACT FROM AUTHOR]

Abstract Solar radiation torque has a great impact on Chinese new generation geostationary meteorological satellite FengYun-4 since its single winged solar array configuration. Daily accumulation in angular momentum stored in reaction wheels may increase about 30Nms, and should be unloaded via bipropellant thrusters during a daily 15-min housekeeping period. How to prevent zero-crossing and saturation of wheels, how to reduce the operation speed range of wheels, and how to ensure the service continuity if any one of wheels fails are the problems to be faced. In addition, both the uncertainty of thruster torque and the variability of solar radiation during all the life span may go against with the automaticity of angular momentum management. In this paper, a null-space-based momentum management method is presented, upon which the angular momentum is managed by null motion control while service and dumped during the housekeeping period. In-flight and numerical simulation results demonstrate the reliability and validity of this strategy. Highlights • The angular momentum management strategy for Fengyun-4 satellite is proposed. • The dumping target is set in three-dimensional space and null space respectively. • Each momentum wheel's rotation speed range is optimized. • Quick momentum self-adjusting is realized when the wheel configuration changes. • Auto-unloading logic is specified for the uncertainty of thrust and solar radiation. [ABSTRACT FROM AUTHOR]

Abstract This paper investigates the multi-objective attitude tracking problem of a flexible spacecraft in the presence of disturbances, parameter uncertainties and imprecise collocation of sensors and actuators. An integrated robust H∞ controller, including an output feedback component and a feedforward component, is proposed, and its gains are calculated by solving Linear Matrix Inequalities. The output feedback component stabilizes the integrated control system while the feedforward component can drive the attitude motion to track the desired angles. The system robustness against disturbances, parameter uncertainties and imprecise collocation is addressed by the H∞ approach and convex optimization. Numerical simulations are finally provided to assess the performance of the proposed controller. Highlights • The attitude tracking control of flexible spacecraft is a multi-objective problem. • An integrated robust H∞ controller is proposed to perform attitude tracking. • The proposed controller is robust to imprecise collocation of sensors and actuators. • The system robustness is addressed by the H∞ approach and convex optimization. [ABSTRACT FROM AUTHOR]

Abstract This paper studies the relative position tracking and attitude synchronization problem of spacecraft rendezvous with the requirement of collision avoidance. To achieve the implementation of the rendezvous procedure, the docking port of the chaser is required to direct towards the counterpart of the target, while the relative distance between the two spacecraft should be larger than the radius of the danger zone during close proximity phase. In order to address the concerned problem, a novel sliding mode control strategy based on artificial potential function is developed, and more specifically, the sliding manifold of the close-loop system is chosen along the negative gradient of the artificial potential function. Within the Lyapunov framework, the proposed control laws are proved to guarantee the convergence of relative position and attitude errors while avoiding any accidental collision between the two spacecraft, even in the presence of external disturbance. Numerical simulations are carried out to demonstrate the effectiveness of the designed control laws. Highlights • Nonlinear relative pose control for spacecraft proximity operation with the requirement of collision avoidance is presented. • State feedback control laws are developed by employing artificial potential field and sliding mode technique. • A 6DOF simulation example is carried out to demonstrate the excellent performance of the proposed control laws. [ABSTRACT FROM AUTHOR]

Abstract This paper deals with applications of fault detection and diagnosis algorithms based on the Kalman filter and fault factor method for an open-cycle liquid propellant rocket engine in a steady state. In order to develop the algorithms, we designed the Kalman filter with a linearized model of a liquid propellant rocket engine and energy balance equations with fault factors. To confirm the fault detection and diagnosis algorithms, we developed a mathematical model of an open-cycle liquid propellant rocket engine and artificially injected various faults, such as decreasing turbine efficiency and a sensor fault. We then executed the fault detection and diagnosis algorithms and analyzed the results of each algorithm. These processes are numerically demonstrated for the open-cycle liquid propellant rocket engine at full thrust level by using the simulated measurement data of the liquid propellant rocket engine mathematical model. Highlights • A mathematical model of an open-cycle liquid propellant rocket engine (OCLPRE) is developed. • Kalman filter and fault factor methods are used for fault detection-diagnosis of the OCLPRE. • The pros and cons of each fault detection-diagnosis algorithms are analyzed. • The performances of the algorithms are demonstrated by numerical simulations. [ABSTRACT FROM AUTHOR]