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There is growing momentum, as articulated through the visions of space agencies, to establish permanent settlements on planetary bodies such as the Moon and Mars. One key hazard that structures will be subjected to will be hypervelocity impacts of micrometeoroids and orbital debris. Shielding against such impacts is essential to ensure the safety of structures and their human inhabitants. There has, however, been limited research on such shields and no field applications to date. In order to inform the composition and design of protective shields for planetary structures, recourse can be made to protective shields utilised in spacecraft and space stations that have benefitted from years of research and development. This paper reviews the existing shielding practices for spacecraft and space stations as a starting point for the shielding of planetary structures. Shield configurations are firstly introduced. Both experimental tests and numerical simulations required to evaluate the performance of the shields are then discussed. Analytical prediction models for shields are reviewed, and the predictions of these models are compared with test results of corresponding shields from the literature. Finally, the effect of parameters such as projectile diameter, impact velocity, impact angle, and shield geometry on the overall shield performance is analysed. This paper ultimately provides technical knowledge about the shielding systems for spacecraft and space stations that can be used as a point of reference for scientists and engineers to develop shields for structures on planetary bodies. • Review of shielding systems on spacecraft and ISS for MMOD protection. • Analysis of theoretical models for predicting protective shield performance. • Recommendations provided for shielding systems of Lunar and Martian structures. [ABSTRACT FROM AUTHOR]

With an expected global revenue exceeding 1 trillion USD by 2040, the space industry is one of the world's fastest-growing sectors. Given the booming investment in the space industry and the anticipated space tourism era, it is crucial to assess the impact of already operative launch assets as well as to adopt design-to-sustainability strategies for the under-development and future launchers. This paper discloses an integrated methodology to estimate nitrogen oxides emissions of a hydrogen-fuelled air-breathing concept. The proposed strategy combines multi-fidelity propulsive system modelling with 0D chemical-kinetic air/hydrogen combustion numerical simulations to define a high-fidelity emissive database representative of various on-ground and in-flight operative conditions. The propulsive and emissive databases are then used to suggest an analytical formulation able to predict nitrogen oxides emission indices in any in-flight conditions, perfectly fitting the conceptual design needs. Throughout the paper, the Skylon spaceplane and its Synergetic Air Breathing Rocket Engine are used as case study. The methodology disclosed allows proving the high competitiveness of this air-breathing space launchers with respect to famous past and current competitors, as the Space Shuttle and the Falcon 9. [Display omitted] • Integrated design methodologies can support NOx estimations in conceptual design. • 0D chemical-kinetic simulations with Z24 scheme provides reliable E I N O x estimation. • Air-breathing launchers prove to be environmentally friendlier than conventional ones. • Upgraded P 3 T 3 method can support E I N O x estimation for hydrogen-fuelled launchers. [ABSTRACT FROM AUTHOR]

Autonomous spacecraft relative navigation via monocular images became a hot topic in the past few years and, recently, received a further push thanks to the constantly growing field of artificial neural networks and the publication of several spaceborne image datasets. Despite the proliferation of spacecraft relative-state initialization algorithms developed, most architectures adopt computationally expensive solutions relying on convolutional neural networks (CNNs) that provide accurate output at the cost of a high computational burden that seems unfeasible for current spaceborne hardware. The paper addresses this issue by proposing a novel pose initialization algorithm based on lightweight CNNs. Inspired by previous state-of-the-art algorithms, the developed architecture leverages a fast and accurate target detection CNN followed by a line segment detection CNN capable of running with low inference time on mobile devices. The line segments and their junctions are grouped into complex geometrical groups, reducing the solution search space, and subsequently, they are adopted to extract the final pose estimate. As a main outcome, the analyses demonstrate that the lightweight architecture developed scores high accuracy in the pose estimation task, with a mean estimation error of less than 10 cm in translation and 2.5°in rotation. The baseline algorithm scores a mean SLAB error of 0.04552 with a standard deviation of 0.22972 in the test dataset. Detailed analyses demonstrate that the uncertainties on the overall pose score are driven mainly by errors in the relative attitude, which gives the highest contribution to the pose error metric adopted. The analyses on the error distributions point out that the uncertainties on the estimated relative position are higher in the camera boresight axis direction. Concerning the relative attitude, the algorithm proposed has higher uncertainties in estimating directions of the target x and y axes due to ambiguities related to the target geometry. Notably, the target detection CNN trained in this work outperforms the previous top scores in the benchmark dataset. The performances of the proposed algorithm have been investigated further by analyzing the effects on the accuracy due to the relative distance and the presence of background in the images. Lastly, the paper delves into the possibility of adopting a sub-portion of the 2D-to-3D match matrix made by the most complex perceptual groups identified that positively affects the overall run-time, pointing out the performances in terms of accuracy of the estimates and providing a comparison of both the baseline and the reduced match matrix versions against state-of-the-art algorithms concerning relative position and attitude errors and solution availability, highlighting the high accuracy and solution availability of the proposed architectures. • New relative pose initialization algorithm based on the detection of line segments. • Detailed literature review and description of available spaceborne image datasets. • Improvement of state of the art in spacecraft detection using lightweight CNN. • Assessment of accuracy of the proposed architecture and error sources analyses. • Proven feasibility of the approach via comparison with other architectures. [ABSTRACT FROM AUTHOR]

Local linearization is often used to enable analytical methods in spacecraft dynamics, navigation, and control. However, the linear approximation becomes inaccurate at large distances from the linearization point and/or when the underlying dynamical system is strongly nonlinear. As a result, linearization-based methods like the finite time Lyapunov exponent (FTLE) can underestimate stretching in phase space. Alternatively, this paper presents a semianalytical method to quantify nonlinear stretching: the eigenpairs of higher-order tensors are used to bound the maximum stretching of a state deviation over time. The tenors in question are derived from the Taylor series expansion of a nonlinear solution flow about a reference trajectory, i.e., the state transition tensor (STT) expansion. It is well known that the eigenvalues of the Cauchy–Green tensor (CGT) bound the stretching of state deviations over time for a linear system, and the largest eigenvalue can be related to the FTLE. This paper will present higher-order, nonlinear analogs to the CGT and the FTLE to more accurately bound nonlinear stretching for applications in guidance, navigation, and control. • Linearization is often used to study the time evolution of state deviations. • Linear methods are insufficient for strongly nonlinear systems or large deviations. • Tensor eigenpairs can be used to bound the nonlinear growth of state deviations. [ABSTRACT FROM AUTHOR]

This paper investigates the orbital pursuit strategy for spacecraft swarms, where both pursuers and evaders possess limited delta-v maneuvering capabilities. Given that swarms can consist of numerous members, ranging from dozens to thousands, the swarm-to-swarm problem has significant coupling effects that cannot be solved by merely superimposing one-to-one or few-to-few problems. Otherwise, one would cause a dimension explosion in decision-making space. To avoid this issue, the swarm orbital pursuit problem is divided into two coupled problems: swarm target allocation and orbital maneuver. First, considering the relative orbit state between pursuers and evaders and estimated orbital maneuver costs of pursuit, the swarm target allocation problem is formulated into three coupled sub-problems, i.e., target allocation order, target allocation for sub-swarms, and target allocation for individuals. These sub-problems are solved by the presented double-layer contract net protocol algorithm to maximize task completion efficiency. Then, the orbital maneuver problem is decomposed into two coupled problems to maximize pursuit benefits, i.e., long-range rendezvous, and close-distance game maneuvers. For the long-range rendezvous, the optimal Lambert algorithm is developed to efficiently obtain the discrete optimal pursuit time and corresponding orbital transfer maneuver. To address close-distance games more proficiently, a hybrid orbital maneuver algorithm, combining the proximal policy optimization and the Lambert algorithm is proposed. Finally, comparison simulations verified the effectiveness and superiority of the proposed method. • Satellite swarm pursuit problem is transformed to the coupled problem with swarm target allocation and orbital maneuvers. • Swarm target allocation is formulated as a coupled problem with target order, allocations for subgroups and individuals. • An improved DLCNP algorithm is proposed to pursue a reasonable target allocation solution for higher task efficiency. • Optimal Lambert algorithm is developed to obtain the optimal maneuver strategyfor long-range rendezvous. • A hybrid algorithm combining PPO and the Lambert algorithm is presented to get closed-loop control for close-distance games. [ABSTRACT FROM AUTHOR]

The increasing space debris population in critical orbits due to spacecraft failure dictates the need for action. On-Orbit Servicing (OOS) has been proposed as a method for mitigating this trend by repairing existing space assets. The development of large servicing spacecraft has been given significant attention. In this paper, we are approaching the problem with a new paradigm for a servicing spacecraft by proposing a CubeSat class servicer equipped with a one-degree-of-freedom (DoF) robotic arm. Considering the OOS context, we choose the Orbital Replacement Unit (ORU) installation on a target spacecraft as a challenging benchmark task for the proposed servicer configuration. To carry out this task, we formulate and compare performance of four controllers to achieve coordinated control of both the CubeSat and the robotic arm for the installation task. We start with a basic PD controller and progress to the more advanced impedance controller, Model Predictive Control (MPC) and Model Predictive Impedance Control (MPIC) designs. The controllers are evaluated and compared across several relevant metrics for the ORU installation and their merits for practical implementation are discussed. • Comparative analysis of model-based controllers against traditional control methods. • Monte Carlo simulations used to test robustness of the controllers. • Novel, mechanically simple CubeSat servicer for simulated ORU installation. • MPC, PD controllers show similar success rates, MPC requires higher control efforts. [ABSTRACT FROM AUTHOR]

Spacecraft positioning plays a crucial role in scientific space mission design, as well as in the further scheduling of scientific observations for such a mission. We examine the application of global navigation satellite systems (GNSS) to solve the orbit determination problem in a pure space very long baseline interferometer (VLBI) project. A network of GPS, GLONASS, Galileo and BeiDou navigation satellites may be a more efficient solution to determining the position and velocity of space radio telescopes and space interferometer baselines. For such a project, it is necessary to take into account the special conditions of GNSS observations, as well as the high accuracy of determining not only position but also velocity. In this paper, we estimate visibility of navigation satellite systems for Low and Medium orbits of a pure space-VLBI system and simulate GNSS code and phase measurements. Orbit determination was performed on smoothed code measurements. Phase measurements simulated with a step of 1 s and combined in Doppler measurements yield average position and velocity errors of 1.01 m and 8.8 mm/s in Medium Earth orbit. The results showed that GNSS observations are sufficient to solve the problems of orbit determination of a space-VLBI interferometer both in Low-Earth and Medium Earth orbits. • Simulated GNSS observations for medium and low Earth orbits using special noise model. • Calculated the visibility of navigation satellites for low and medium Earth orbits. • Obtained navigation solution for space radio interferometers in low and medium orbits. • GNSS observations provide sufficient accuracy of orbit determination. [ABSTRACT FROM AUTHOR]

Mitigating solar eclipses presents a formidable challenge in the design of spacecraft orbits, particularly in the case of a distant retrograde orbit (DRO), which is susceptible to significant solar eclipse threats. This paper undertakes a comprehensive analysis of the occurrence characteristics and patterns of both Moon and Earth eclipses on a DRO. The proposed approach involves a two-stage design procedure, encompassing initial orbit selection and orbit extension. The initial orbit selection method primarily focuses on mitigating Earth eclipses by employing the out-of-plane motion inherent in a DRO. The study meticulously explores the influence of three pivotal elements — amplitude, period, and initial phase — on the efficacy of Earth eclipse avoidance. Concurrently, when faced with unavoidable solar eclipses of the initial orbit, a series of tiny maneuvers are systematically applied to extend the baseline orbit. The parameter design space is discretized, and a tree search method is employed to identify viable maneuver sequences. The research successfully achieves the long-term stabilization of the baseline DRO, ensuring that solar eclipses do not exceed 2 h in 10 years. The proposed method's robustness and applicability are substantiated through validation under the ephemeris model. • The occurrence characteristics and patterns of eclipses on DRO are analyzed. • The two-stage design approach contains initial state selection and orbit extension. • The out-of-plane motion of quasi-DRO is leveraged to mitigate Earth eclipses. • Maneuvers are applied to extend the DRO baseline, using the tree search method. • The robustness and applicability of the method validated under the ephemeris model. [ABSTRACT FROM AUTHOR]

This paper concentrates on robust attitude control issues for spacecraft formation in gravitational-wave detection missions. Notably, with the large-scale communication delays brought by the million kilometers-scale inter-satellite distance, the exponential convergence of the closed-loop delayed system is firstly achieved via a new delay-dependent nominal attitude coordinated control scheme. Then, a finite-time disturbance observer is deployed to reconstruct the unknown lumped disturbance from internal uncertainties and the external environment. Meanwhile, an integral sliding manifold is embedded in the delay-dependent attitude controller for superior systems' robustness. Finally, two simulation examples illustrate the feasibility and effectiveness of the proposed attitude control strategy for in-orbit gravitational-wave detection spacecraft systems. • Spacecraft formation attitude control in gravitational-wave detection is studied. • Large-scale communication delays (LCD) among formation members are firstly modeled. • An attitude coordinated controller is put forward to handle the LCD. • An integral sliding manifold with disturbance observer is embedded for superior robustness. [ABSTRACT FROM AUTHOR]

A long-duration microgravity environment has numerous detrimental effects on the human physiology. The most obvious solution for this problem related to long duration space exploration missions is to remedy the lack of gravity. This could be done using short arm human centrifuges but they do not seem to be sufficiently effective, perhaps because of the short duration exposure of this countermeasure and/or the huge body gravity gradient. New views have to be investigated, such as to see if a (very) long-arm rotating system generating a continuous 1 g or partial gravity field might resolve this issue. Besides the expected benefits regarding astronauts' microgravity pathologies, additionally the spacecraft itself, its on-board (sub-)systems and procedures might benefit from a rotating configuration. In this paper we address very briefly the medical issues, but the work is mainly focused on the advantages regarding engineering, operations, life support, safety and budget of having a constantly rotating spacecraft first in Low Earth Orbit and later for long duration missions to Mars. A large rotating spacecraft is feasible and affordable to build, operate and maintain. It has advantages for governmental and commercial use but also in light of the expected increase in space tourism. It will also save crew time and billions of dollars now being spent to counteract the effects of microgravity. A continuous rotating spacecraft providing Artificial Gravity will: • Improve crew health and wellbeing. • Improve mission safety and success. • Reduce costs significantly. • Simplify numerous operations in flight. • Provides the use of 1 g verified systems. [ABSTRACT FROM AUTHOR]

Until recently space travel was considered a privilege of governmental astronauts, with the exception perhaps of a few billionaire space tourists, who had been accommodated on the International Space Station on some rare occasions. Nonetheless, with major space pioneers announcing plans for space tourism activities and after the successful launch in orbit of space tourism vehicles from the Virgin Galactic and Blue Origin companies, academic interest about human spaceflight for purposes of leisure and recreation has ignited. Taking into account that space tourism is in fact an unregulated activity, this paper will firstly examine whether space tourism activities fall under the scope of space law or under the scope of international air law, depending on the planned trajectory of the space tourism vehicle. Furthermore, the authors will attempt to draw conclusions concerning the legal implications of space tourism activities through a comparative analysis with the 1974 Athens Convention on the Carriage of Passengers and the Carriage of their Luggage by Sea. This international convention provides a great example on the regulation of touristic ventures, as it refers to commercial touristic activities realized within a res communis omnium , i. e., the high seas, similarly with touristic activities in space. Moreover, said Convention provides for a specialized regime for touristic activities and is of particular interest, since it has achieved stability and a desirable balance of interests between private tourism operators and passengers. In this connection, the authors will evaluate the provisions of the 1974 Athens Convention in comparison with the space treaties applicable to space tourism activities. Furthermore, suggestions will be made as to whether it would be appropriate and feasible to introduce in the space law realm certain rules of law similar to those of the 1974 Athens Convention, in order to facilitate growth in the space tourism sector, without compromising the safety of human spaceflight for space passengers. • The 1974 Athens Convention in comparison with the space treaties applicable to space tourism activities. • Legal analogy between space law and maritime law. • Legal regime governing space tourism activities. [ABSTRACT FROM AUTHOR]

Flexible vibration has a great impact on flexible spacecraft attitude control, which will reduce the pointing accuracy of the spacecraft platform and bring severe challenges to the design of a high-precision attitude control system for flexible spacecraft. Aiming at the problem of attitude and vibration coupling control of flexible spacecraft, an anti-unwinding attitude coupling controller using the positive position feedback (PPF) control is proposed in this paper. The model compensation control term is constructed based on the prior information to improve the tracking performance. The proposed attitude coupling control method can effectively suppress the flexible vibration while achieving more accurate attitude control. The obtained closed-loop system under the proposed control scheme is asymptotically stable, which is proved by the Lyapunov stability theory. The numerical simulation illustrates that the proposed method can effectively improve the attitude control accuracy of the flexible spacecraft and is capable of large-angle attitude maneuver control because of its prominent performance in attitude tracking and its unwinding-free performance. • Coupling control for flexible spacecraft. • Active vibration control using positive position feedback. • Model compensation control term is constructed to improve the dynamic performance. • The obtained closed-loop system achieves the asymptotic stability and anti-unwinding performance. [ABSTRACT FROM AUTHOR]

This paper addresses the fault-tolerant attitude tracking control problem for spacecraft with limited communication capability and input saturation. The design of a hysteresis quantizer aims to alleviate communication burden between the controller module and the actuator module. Then, the attitude tracking problem for spacecraft with input quantization, input saturation, actuator faults and external disturbances is transformed into an attitude control problem with uncertain input coefficients and bounded disturbances. Thereafter a dynamic loop gain function-based approach and a hyperbolic tangent function term with a time-varying boundary layer are introduced to address the uncertainties of input coefficients and the bounded disturbances, respectively. To facilitate the boundness analysis of the signals in the closed-loop system, a pertinent lemma about the dynamic loop gain function is proved. Finally, numerical simulations are employed to validate the effectiveness of the proposed scheme. • The spacecraft attitude tracking problem with disturbances and uncertainties. • The quantized control law alleviates communication burdens of spacecraft. • A dynamic loop gain function method is used to address the impact of uncertain. • A hyperbolic tangent term is devised to counteract bounded disturbances. [ABSTRACT FROM AUTHOR]

This paper introduces a flexible framework to enable high-fidelity simulations of the fuel slosh dynamics for verification of the Guidance, Navigation and Control (GNC) algorithms. The sloshing phenomenon must be tackled during the control system design phase and the GNC software run and validated within appropriate simulation frameworks. Equivalent mechanical models can well approximate sloshing under specific assumptions, allowing the derivation of the multibody vehicle's equations of motion; however, while suitable for analysis and control synthesis purposes, this approach presents several limitations for implementation as a simulation model. In this work, the multibody plant embedding slosh dynamics is modeled by means of the DLR's Vertical Landing Vehicles Library (VLVLib) written using the object-oriented Modelica modeling language. The advantages of using Modelica are explored throughout the paper. Given that control synthesis, analyses and simulation campaigns are usually performed in the Matlab/Simulink environment, a core point of the proposed solution is to achieve a good synergy within the latter and the Modelica plant model with the least readaptation burden for the developer. The strategies to achieve a full integration and verification tool chain are illustrated. The potential of this approach is demonstrated via two different test cases: a lunar landing scenario of the ALINA spacecraft developed by PTS, and CALLISTO reusable rocket demonstrator. • A flexible framework to enable high-fidelity simulations of the fuel slosh dynamics for verification of the GNC algorithms is introduced. • The spacecraft model embedding the slosh dynamics is implemented exploiting Modelica language. • A GNC testing framework is obtained with the integration of the causal Simulink environment with the compiled Modelica model. • Two test cases are presented: the ALINA spacecraft lunar landing scenario and CALLISTO reusable Vertical-Takeoff–Vertical-Landing demonstrative mission. [ABSTRACT FROM AUTHOR]

Attitude control methods of highly flexible spacecraft have seen increased interest over the last decades thanks to the technological development of flexible solar panels and deployables, which improves the capabilities of satellites. However, a high-fidelity model of the flexible dynamics is hard to obtain on-ground testing because not all modes can be observed, complicating the controller design. This paper proposes a method to develop a high-fidelity model of a spacecraft with a flexible appendage subject to large deformations by modeling it as a finite series of rigid links connected by torsional springs and dampers. To overcome the uncertainties or unknowns in the flexible dynamics, an onboard estimation through an adaptive controller is performed for these unknowns while the spacecraft is maneuvered. The controller uses integral concurrent learning (ICL), an adaptive scheme that records inputs and outputs provided by sensors mounted on the flexible body. The novelty of this investigation is the development of self-adapting control gains for the learning matrix obtained from ICL. This approach achieves the objective of tracking a desired trajectory while accurately learning the unknown physical parameters of the flexible appendage by only using the recorded measurements of the appendage displacements and the spacecraft attitude and velocities. It was observed that for a finer discretization of the flexible appendage and therefore a higher fidelity model of the flexible dynamics, the estimation algorithm is able to observe all the frequencies necessaries to learn the unknown mechanical properties of the flexible body. • Modeling a flexible deployable as finite chain of rigid links connected by unknown torsional joints to consider large deformations and reduce onboard computational effort. • Accurately estimated unknown flexible dynamics using integral concurrent learning (ICL). • Introduced a self-tuning algorithm for eigenvalue placement of the learning matrix to guarantee the convergence of any number of estimated unknowns to their true value. [ABSTRACT FROM AUTHOR]

Space debris near the synchronous orbit poses significant collision risks to spacecraft in this area. Therefore, it is necessary to analyze the orbital dynamics of space debris in order to predict its orbital evolution and prevent it from colliding with the spacecraft. Due to the combined effect of the dissipative forces and Earth's 1:1 tesseral resonance, the space debris with high area-to-mass ratio near the synchronous orbit exhibits two special dynamical behaviors: resonance capture and resonance jump. By using a simplified model, this paper analyzes in detail the orbital evolution of the debris of the resonance capture and resonance jump, and explores the conditions for their occurrence. It is found that the dynamical behavior of space debris around the resonance capture or resonance jump is highly sensitive to its initial conditions. In addition, the long-term direct numerical integrations with the full dynamical model for the orbital evolution of the debris are performed. The difference between the phase space parameters of the resonance capture and resonance jump for space debris is discussed, and the pendulum-like oscillation of the resonant angle is also studied. • Resonance capture and resonance jump for space debris around the synchronous orbit. • Initial conditions of resonance capture and resonance jump of space debris are found. • Resonance capture and resonance jump of debris are sensitive to initial conditions. • Resonant angle of near-synchronous debris exhibits pendulum-like oscillation. [ABSTRACT FROM AUTHOR]

Coilable booms made of thin-ply composites provide mass and volume efficiency for deploying large structures in space. The recovered shape of the booms after being stowed in the coiled configuration is critical to the deployed stiffness and shape precision of the spacecraft, but is heavily influenced by the viscoelasticity of the polymer composites. A numerical study of coiling, stowing, uncoiling and shape recovery of a viscoelastic thin-ply composite coilable boom is presented in this paper. The formulation of two micromechanics-based viscoelastic shell models and their numerical implementation in structural simulations are presented. The shape recovery of a composite boom coiled for two years, replicating the typical timeline for deployable structures used in solar sail missions, has been simulated. The demonstrated modeling and simulation capability is useful for designing deployable spacecraft structures. • Ply homogenization computes viscoelastic stiffness faster than laminate homogenization. • Ply homogenization has comparable accuracy to laminate homogenization. • Composite deployable booms for solar sails have minimal distortions after stowage. [ABSTRACT FROM AUTHOR]

NASA's risk classification system dates back to an era when every new NASA space mission was a one-of-a-kind build, and the only way to obtain reliability was as a by-product through a combination of reliability analyses, extensive and stringent quality requirements, and extensive testing. Originally, there were very limited commercial capabilities to develop systems to work reliably in space, so NASA considered its own homegrown approach the only recipe for success. This approach involved very detailed and prescriptive piece-part controls and no reliance on (and to some extent a rejection of) any type of commercial practices. Often risk was considered to be the lowest when NASA had the maximum amount of control and prescription, and the highest when commercial practices were largely employed, and these principles drove risk classification in the agency. Over time, however, commercial capabilities grew, and many products became standardized and commercialized, while the agency maintained its tried-and-true approach, paying little attention to the evolution of the commercial sector. In fact, the commercial sector was developing systems that have direct, proven reliability, established over time, while NASA still maintained the approach to ignore the reality of the commercialized aspects of standard products, label them as high risk, and attempt to change them to align with the agency's piece-part control practices. A table of mission classification vs lifetime for missions launched after 2000 indicates no correlation between lifetime and classification, with the few exceptions involving missions that have very limited objectives and no valid purpose to continue after they were met. This paper steps through some of the key historical elements in risk classification and NASA's overall approach to assurance, and presents some elements being brought forward to modernize the approach and take advantage of the growing capability in the commercial sector. • NASA's approach for assurance and mission risk classification dates to the original premise of one-of-a-kind builds for spacecraft and instruments. • TThis paper identifies the clash between piece part controls and established components, and provides a new approach [ABSTRACT FROM AUTHOR]

50,000+ satellites are scheduled for launch by 2030, bringing both massive opportunities and equally massive challenges. Key among these challenges is the ability to sustainably operate Low Earth Orbit (LEO) spacecraft in an increasingly congested environment where collision avoidance manoeuvres are routine and fuel is limited. This paper explores the application of ionospheric drag as an alternative, propellantless collision avoidance mechanism for satellites with altitudes ranging between 350 and 500 km. This is achieved through the use of an in-house propagation suite developed to support UNSW Canberra Space's in-orbit Ionospheric Aerodynamic Experiment (IEX) on the M2 spacecraft (NORAD ID: 47967 & 45727). The relative impact of neutral (thermospheric) to charged (ionospheric) drag on along-track separation for a range of spherical bodies with surface potentials ranging between 0 V to -100 V with respect to a quasi-neutral freestream plasma was investigated. This provides a first-order indication of the potential effectiveness of ionospheric drag for collision avoidance based on prior numerical and ground-based investigations into ionospheric drag. The Debris Risk Assessment and Mitigation Analysis software from the European Space Agency was also harnessed to discern the relative performance of both the charged and flared attitude satellites against a common control satellite. The ESA stipulates that a Collision Probability Level value greater than the accepted value of 1 E − 04 between two satellites requires a risk avoidance manoeuvre, with a conservative approach requiring manoeuvres for values greater than 1 E − 05 . For all test regimes at 350 k m altitude, at least a 60% reduction in time to achieve the NASA approved risk reduction factor of 31.6 at the time of closest approach for a pre-manoeuvre Collision Probability Level of 2. 8 E − 05 was achieved when using the flared neutral method in comparison to the feathered charged method. Based on these results, this paper finds that ionospheric drag may represent a useful alternative propellantless collision avoidance mechanism for satellites within the altitude range of 350 to 500 km. Future work is needed to refine ionospheric drag models and consider the absolute effectiveness of using ionospheric drag at higher altitudes and in tandem applications to the flared arrangement. • Ionospheric drag can successfully manoeuvre satellites to avoid collisions. • Ionospheric drag is inferior to industry standard avoidance manoeuvres. • High Low Earth Orbit altitudes show ionospheric drag becomes more competitive. • The adopted python approach provides flexibility for future satellite simulations. [ABSTRACT FROM AUTHOR]

Model-based systems engineering (MBSE) is an emerging methodology not only in the field of spacecraft systems engineering. The key objective is to store the whole system design in one single source of truth system model. The continuously maintained and customizable open-source MBSE tool Virtual Satellite developed by German Aerospace Center (DLR) offers a framework for storing such a system model. Since the open-source 3D computer aided design (CAD) software FreeCAD has a dedicated Virtual Satellite workbench for importing the 3D model file generated by Virtual Satellite, it is possible to export the 3D model from Virtual Satellite, apply changes to it in FreeCAD, and afterwards reimport it in Virtual Satellite. This paper presents a workflow implementation to conduct automated steady-state and transient thermal analyses using the Virtual Satellite system model as a basis, utilizing the aforementioned features. Part of this effort was to extend Virtual Satellite by thermal modeling capabilities, such as material parameters, surface characteristics, thermal interfaces, boundary conditions and further relevant aspects. With custom Java applications, so-called "Apps", all aspects of the system model relevant for a thermal analysis are exported from Virtual Satellite and written to accordingly generated text files. Moreover, the before described 3D model export is used to export the model to FreeCAD. The FreeCAD Python console, in combination with the finite element method (FEM) workbench, is used to create a script that executes meshing, application of the boundary and initial conditions, as well as identification of contacts between parts and application of the specified contact conductance values. The obtained files form a complete thermal model and can be fed directly into the open-source FEM software CalculiX for executing the thermal analysis. Then, the result can be processed by an additional Virtual Satellite App. To account for the space thermal environment an additional feature was added to the FreeCAD script. The feature uses orbit information obtained from external mission analysis tools to calculate the approximate combined heat load from Sun, Earth infrared, and albedo radiation at each time point in the specified simulation time interval. This new workflow allows to execute quick thermal analyses at an early design stage and therefore opens new opportunities in the evaluation of different system designs, as early as in a concurrent engineering study. Configuration parameters are provided that also allow more sophisticated analyses in later development phases, such as custom mesh definition and local mesh refinement. • Thermal design process of spacecraft relies on engineer's experience in early stages. • Potential help for the thermal design process is thermal simulation. • Thermal simulations require significant manual setup effort. • Created workflow automates simulation setup and result feedback. • Data for simulation setup maintained in MBSE system model. [ABSTRACT FROM AUTHOR]

Slosh dynamics in a liquid propellant tank during a spacecraft maneuver can have an effect on the system stability and potentially lead to critical problems for the mission operation. In particular, the fuel sloshing problem has drawn significant attention in aerospace applications where a large portion of the gross weight is comprised of liquid fuel. Recently, many active control methods have been introduced as alternatives to passive approaches. However, the majority of these methods basically require the full-state condition. To guarantee the availability of the full-state condition, this paper proposes a cascade nonlinear state observer based on the sliding mode theory for estimating the unmeasurable slosh variables in a planar model. Additionally, for large amplitude estimation, a complete nonlinear mathematical model with a pendulum analogy is formulated without a linearization process. The stability analysis is theoretically proven by Lyapunov method. The effectiveness is demonstrated through simulations, and the performance comparison with a different observer method is also presented in the results. [Display omitted] • Observer design of cascade structure for slosh-variable estimation. • Complete nonlinear equation with pendulum analogy for slosh dynamics. • Full-state availability and large amplitude estimation are achieved. • The stability and small bound convergence are theoretically proven. • The effectiveness and performance comparison by numerical simulations. [ABSTRACT FROM AUTHOR]

This paper examines options for orbit configurations for a space interferometer. In contrast to previously presented concepts for space very long baseline interferometry, we propose a combination of regular and retrograde near-Earth circular orbits in order to achieve a faster filling of (u , v) coverage. With the rapid relative motion of the telescopes, it will be possible to quickly obtain high quality images of supermassive black holes. As a result of such an approach, it will be possible for the first time to conduct high quality studies of the supermassive black hole close surroundings in dynamics. • A new configuration for the space interferometer is proposed. • This configuration uses near-Earth regular and retrograde orbits. • Calculated orbits are stable more than 10 years. • The supermassive black hole can be imaged within 20 h. • Synthetic simulations show observations of black holes are possible in dynamics. [ABSTRACT FROM AUTHOR]

Estimating the pose of an uncooperative spacecraft is an important computer vision problem for enabling the deployment of automatic vision-based systems in orbit, with applications ranging from on-orbit servicing to space debris removal. Following the general trend in computer vision, more and more works have been focusing on leveraging Deep Learning (DL) methods to address this problem. However and despite promising research-stage results, major challenges preventing the use of such methods in real-life missions still stand in the way. In particular, the deployment of such computation-intensive algorithms is still under-investigated, while the performance drop when training on synthetic and testing on real images remains to mitigate. The primary goal of this survey is to describe the current DL-based methods for spacecraft pose estimation in a comprehensive manner. The secondary goal is to help define the limitations towards the effective deployment of DL-based spacecraft pose estimation solutions for reliable autonomous vision-based applications. To this end, the survey first summarises the existing algorithms according to two approaches: hybrid modular pipelines and direct end-to-end regression methods. A comparison of algorithms is presented not only in terms of pose accuracy but also with a focus on network architectures and models' sizes keeping potential deployment in mind. Then, current monocular spacecraft pose estimation datasets used to train and test these methods are discussed. The data generation methods: simulators and testbeds, the domain gap and the performance drop between synthetically generated and lab/space collected images and the potential solutions are also discussed. Finally, the paper presents open research questions and future directions in the field, drawing parallels with other computer vision applications. • Deep Learning is used extensively for monocular spacecraft pose estimation. • Deployability is an important limitation of the current DL-based solutions. • Lack of space-borne images during training leads to the domain gap problem. [ABSTRACT FROM AUTHOR]

Spacecraft in low-Earth orbit are exposed to environmental threats which can lead to material degradation and component failures. The presence of atomic oxygen and collisions from orbital debris have detrimental effects on the structures, thus affecting their performance. Cranfield University has developed a family of drag augmentation systems (DAS), for end-of-life de-orbit of satellites, addressing the space debris challenge and ensuring that satellites operate responsibly and sustainably. De-orbit devices are stowed on-orbit for the duration of the mission lifetime and, once deployed, the devices must withstand the harsh low Earth environment until re-entry; a process which can take several years. The DAS' deployable aluminised Kapton sails are particularly susceptible to undercutting by atomic oxygen. In preparation for commercialising the DAS, Cranfield University are investigating the degradation process of the drag sail materials, with the end goal of qualifying the materials for the specific application of drag sails in low Earth orbit (LEO). This paper will outline the proposed research and the expected benefits from the projects. This paper will conclude in a summation of the different on-going research projects at Cranfield University related to commercialising the DAS family. This research will benefit the wider space community by expanding the understanding of the effects of long-term exposure on certain materials, as well as improving the validity of future low Earth atmospheric models. • Drag augmentation systems for end-of-life de-orbit of satellites. • Low-cost, simple solutions for future space debris mitigation. • Long-term degradation of aluminised Kapton in low Earth orbit. • Testing campaigns characterising material, investigating atomic oxygen erosion. • Assessment of the drag sail environmental durability, performance and scalability. [ABSTRACT FROM AUTHOR]

Motivated by the fact that the mass of the Moon is about one percent of the Earth mass we consider in this paper Keplerian orbits of satellites (or spacecrafts) in the Earth–Moon system. These orbits represent approximate solutions of the basic equations of motion. As a result a satellite in one of these orbits will require some thrust to stay in orbit. In spite of this issue these orbits might be of interest when satellites orbits have to satisfy some practical constraints. In this paper we explore two types of such orbits. The first is around the Earth–Moon center of mass. The second consists of orbits whose center is along the line from Earth to the Moon but at some distance from Earth–Moon center of mass. As examples we consider briefly orbits for an Earth observation station and an Earth–Moon cycler (Casoliva et al., 2008). (In the following we use the term "Earth–Moon shuttle" to mean "Earth–Moon cycler"). Throughout this paper we use symbiosis between analytical and numerical methods to derive the desired results. • We consider Keplerian type orbit in the Earth–Moon System. • Thrust and Energy requirements are addressed. • Circular Orbits have minimum energy requirements [ABSTRACT FROM AUTHOR]

Magnetic control has been used successfully for momentum management and attitude control when pointing precision requirements are loose. Typically, such systems are constructed using configurations of magnetic coils oriented in a way to provide command authority by varying the magnetic moment about all three body axes. Although magnetic control actuators and related systems are typically robust and not as prone to failure when compared to reaction wheels or thrusters, it is of interest to see if coarse 3-axis attitude pointing can be achieved with the loss of control about one of the 3 axes of magnetic moment actuation. These types of cases could occur during the normal life of the spacecraft or could also occur intentionally on very small satellites where space is limited, so any control modes that could prolong the life of a damaged satellite or intentionally underactuated satellite would be useful for some missions. This paper demonstrates the feasibility of 3-axis pointing of spacecraft using only two axes of magnetic moment actuation using linear time varying analysis to provide sufficient conditions for controllability. The paper then develops an LQR-based controller for these two torquer cases. Finally, the paper concludes with simulations that indicate it is feasible to control with only two axes of magnetic moment, although with the expected increased time to achieve targeted attitudes and reduced disturbance rejection. • Identifies sufficient conditions for controllability in systems with only two magnetic coils. • Implementation of a LQR controller using two magnetic coils. • Testing of controller with disturbance torques relevant to Low Earth Orbit. • Monte-Carlo analysis to test robustness of this approach to parameter variations. [ABSTRACT FROM AUTHOR]

The vast majority of Earth-orbiting satellites carry an electrical power subsystem (EPS) which main components are solar panels and secondary batteries. During eclipse periods, satellites are powered only by rechargeable batteries which have a large energy density but a limited power density. This fact limits the power capabilities of small satellites during eclipse periods. Due to the large power density of Supercapacitors (SCs), ground and on-orbit tests have been conducted to verify their applicability on satellite EPSs. At the moment, no studies are underway on the issue of sizing SCs for eclipse operations. Hybrid configuration could reduce the mass and volume of EPS or maintain those reference values but increasing peak power capabilities. This paper deals with this issue. On the one hand, novel analytical expressions of a variable capacitance SCs are derived, including time–voltage dependence in constant power applications. Also an equivalent formulation for a SCs bank (SCsB) is derived. On the other hand, a SCsB sizing procedure is presented, considering the energy and power requirements of a particular space mission as an input. Finally, a simple mission is analysed, the results showing an improvement on the hybrid EPS design with respect to the traditional, being the mass reduction of 36%. • Implementation of a simple but accurate supercapacitor (SC) model for spacecraft applications. • Innovative analytical expressions of the performance of a SC in constant power applications. • Novel determination of the electrical parameters of a SCs bank, treated as an equivalent SC model. • Estimation of key variables in the operation of hybrid storage systems for space applications. • Preliminary design of a SCs bank devoted to supply high power during eclipse times. [ABSTRACT FROM AUTHOR]

Spacecraft modularity is a powerful design methodology for shortening satellite development cycles and reducing costs. The exploitation of standardized and reusable modules, universal interfaces and protocols, generic design, and a set of defined conventions, allows for rapidly integrating a flexible and cost-effective platform with ease of component modifications. Anyhow, the implementation of modularity poses several issues at any level of mission and spacecraft design, including subsystem design. This paper deals with the Attitude Determination and Control Subsystem (ADCS), with a particular accent on making the estimation software independent of any specific requirements dictated by the application in question. In this respect, a generic and reusable Murrell's Multiplicative Extended Kalman Filter (MMEKF) that can operate on different satellites, with diverse estimation requirements, working with changeable integrated ADCS boxes and sensor suites, but without changes either in the equations or in the interface with the rest of the system, has been proposed. The algorithm has been designed to maximize flexibility and ease of applicability to different missions without a priori-knowledge of the actual configuration. The execution of the algorithm is based on a plug-and-play logic in which middleware software enables the automatic discovery of the available sensors and the transmission of required data to the estimation algorithms. An adaptable and modular 3D simulator has been built to test and validate the performance of the generic MMEKF over multiple test cases with assigned configurations. The results of the simulations confirm the capability of the algorithm to meet estimation requirements for each simulated case, without ad-hoc , case-by-case, optimization, or customization. • A generic and reusable estimation algorithm is proposed to estimate the attitude in a modular plug-and-play satellite. • The algorithm is designed by adapting Murrell's Kalman Filter to be fully parametric and mission abstracted. • The method involves a plug-and-play execution based on centralized middleware software. • Relevant requirements and constraints to reuse the same estimation algorithm in heterogenous missions are pointed out. • Performance in the estimation accuracy is verified over multiple cases with assigned configurations. [ABSTRACT FROM AUTHOR]

This paper deals with the problem of constructing a flight scheme to Venus, in which a spacecraft flying to the planet after a gravity assist maneuver and transition to a resonant orbit in order to re-encounter with Venus, makes a passage of a minor celestial body. The 117 candidate asteroids from the NASA JPL catalogue, whose diameter exceeds 1 km, were selected. The flight trajectories which meet the criteria of impulse-free both flyby Venus and asteroid, and the subsequent landing on the surface of Venus were found within the interval of launch dates from 2029 to 2050. The trajectory of the spacecraft flight from the Earth to Venus including flyby of Venus and asteroids with a subsequent landing on the surface of Venus was analyzed. • Found 53 possible scenarios for a flight to Venus with free-impulse flyby of an asteroid for the launch dates from 2029 to 2050. • In the framework of the developed schemes it is possible to perform the free-impulse encounter with comet 2P/Encke for launch in 2032 and 2044. • The impulse-free flyby of Venus and the asteroid occurs when the spacecraft transfers to resonant 1:1 orbit after Venus flyby. • The link between attainable landing points and conditions of the free impulse flyby of an asteroid was established. [ABSTRACT FROM AUTHOR]

Collisions with micrometeoroids and space debris can perforate spacecraft shields and hulls; in-situ repair or substitution of the damaged elements is currently considered a complex task, in particular the case of crewed modules with pressurized elements. In this context, cold-welding can overcome the current technological difficulties allowing fast and safe repairing procedures. To date, it has been shown that similar metallic materials can fuse or weld at ambient temperatures provided that there are sufficiently high contact forces. In the space environment this fusion is aided by the fact that the joint surfaces do not re-oxide after wear and, as a consequence, atomic diffusion of the metal occurs at lower contact forces. It has also been demonstrated that, even under terrestrial conditions, the action of a low fretting load can nearly double this adhesion force. In this paper, the suggested utilization of cold-welding phenomenon for spacecraft hull repair is described, as well as a verification road map to demonstrate its feasibility. It is proposed to develop an experimental test rig to apply custom repair patches of different materials to pre-damaged metallic structures and monitor the performance the adhered joint in low orbit and during re-entry. The test samples will include simple plates, Whipple shields, and sandwich panels that have been subjected to hypervelocity impacts. In the first phase, the cold-welding phenomena will be investigated in ground laboratories and validated in a vacuum chamber; afterwards, it is proposed to repeat the experiment in microgravity conditions, during a sounding rocket flight. Results will include an assessment of the contact forces necessary to apply the repair patch and an evaluation of the structural properties of the refurbished systems. • Cold welding for spacecraft repair following a hypervelocity impact by space debris. • In-situ repair from inside the spacecraft is preferable. • Test apparatus to apply custom repair patches. • Experimental setup integration as a sounding rocket payload. • Hypervelocity impacts tests are designed to have realistic samples. [ABSTRACT FROM AUTHOR]

Attitude control is an essential operation for spacecraft to perform space missions, and control strategies play a decisive role. In this paper, a second-order quaternion system of rigid spacecraft attitude dynamics is derived based on the fully actuated system approach. The quaternion system has two constraints that lead it to be a fully actuated system. Two closed-loop control strategies are proposed to achieve attitude tracking and stabilization for the fully actuated system. The stabilization strategy can control the scalar quaternion to reach the desired value and be used to deal with the unwinding phenomenon. Numerical simulations demonstrate the second-order quaternion system's validity and the control strategies' effectiveness. • A second-order quaternion system is derived from an intuitive perspective. • The control laws are much easier to design than the first-order system. • Arbitrary three quaternions can be employed as the controlled variables. • The unwinding phenomenon is handled in a simple way. [ABSTRACT FROM AUTHOR]

For impulse transfers, the solution of two-impulse rendezvous is generally obtained by solving the Lambert orbit problem, and the search for the exploration window of the target celestial body is a two-dimensional optimization based on velocity increment calculation. In order to assess the accessibility and exploration window of large-number asteroids, this paper approximates the transfer orbit from the Earth to the target asteroid as a co-planar two-impulse transfer orbit and gives an effective method. Based on the theory of cotangential transfer orbit, this paper proposes a method to express the transfer orbit with one single variable (the aphelion argument of the transfer orbit, θ T), which can also determine the start time of the transfer. After that, the paper combines the two-impulse orbit transfer theory (without considering the rendezvous) with the calculation method of the exploration window (considering the rendezvous), proposing the concept of the rendezvous difference angle ε R. ε R is defined as the mean anomaly angle difference between the spacecraft and the target object when the spacecraft enters the target orbit. The rendezvous difference angle ε R has the same minimum point as the velocity increment in exploration window search, while it can be directly calculated from the transfer orbit and the transfer time. When cotangential transfer orbits are used, ε R can be expressed as a function of θ T. Taking ε R as the optimal target variable, the exploration window search can be simplified to a one-dimensional problem. This method is suitable for the large-scale accessibility calculation for asteroids with low inclinations, so as to quickly determine the accessible target group. • A new method for launch window calculation in asteroid exploration missions is proposed. • The optimal and sub-optimal transfer orbit from the Earth to an asteroid is analyzed. • Analyse of two exampled asteroids was carried out. [ABSTRACT FROM AUTHOR]

In 2021, the DART spacecraft will be launched by NASA to intercept the binary system Didymos and impact the moonlet (Dimorphos), to test the effectiveness of kinetic impactors for the deviation of hazardous asteroids' trajectories. The impact is expected to generate a cloud of particles, whose study could provide valuable data about asteroid's properties and impact models. Observations from ground to characterize ejecta and plume evolution after impact would be possible, but the resulting quality would be significantly lower than a short-range imaging. In this framework, ASI (Italian Space Agency) and NASA started a collaboration to embark on the US DART vehicle the LICIA (Light Italian CubeSat for Imaging Asteroid) 6U CubeSat as a piggyback payload. LICIA is devoted to grasp science data by imaging the ejecta plume after the impact, and observe Dimorphos surface while flying by the binary system. The Italian consortium involved in the mission sees ARGOTEC for the platform development, and INAF, Politecnico di Milano and Università di Bologna to cover the mission science, trajectory design and orbit de-termination tasks, respectively. The paper focuses on the mission analysis and maneuvers design, strongly driven by the science return, under the tight constraints enforced by the platform. In particular, it describes the preliminary procedure followed to derive trajectory constraints and to select the best option for the flyby. Despite the mission is still under development, and subjected to frequent updates of several parameters, the final performances derived in this paper maintain their validity as they represent the desired target values for the mission. • The LICIACube mission will serve as scientific support for NASA's DART kinetic impactor. • LICIACube will be the first CubeSat to visit an asteroid system. • Hardware limitations and the complexity of the mission narrow down the trajectory options. • Accurate optimization is adopted to identify best options for scientific return maximization. • Robustness of the selected trajectory, in presence of uncertainties and contingencies, is derived. [ABSTRACT FROM AUTHOR]

The robust non-fragile attitude stabilization controller for flexible spacecraft with model parameters uncertainty is difficult to be synthesized with negative imaginary system which is widely applicable for the vibration control of flexible structures. To address this issue, this paper proposes a negative imaginary output feedback H ∞ synthesis strategy, in which the model parameters uncertainty, controller gain perturbations, unknown inertia and actuator faults are considered simultaneously. This strategy can guarantee that the closed-loop attitude control system has asymptotical stability with H ∞ performance and approximates to negative imaginary. The iterative algorithm under linear matrix inequality constraints gives the conditions for the existence of such controller. In addition, external disturbance and input saturation are also taken into account. Compared with the general non-fragile H ∞ controller, the numerical simulations demonstrate the superiority of the proposed controller in attitude stabilization and vibration suppression of flexible spacecraft. • The spacecraft attitude and vibration are stabilized with controller perturbations. • The NI controller can well suppress the elastic vibration of flexible appendages. • The two-stage framework is of effective controller performance synthesis. [ABSTRACT FROM AUTHOR]

Utilizing the autonomous rendezvous and docking among multiple spacecraft to construct a large space structure is a prospective on-orbit autonomous assembly mission. This paper presents a planar air-bearing experiment on the autonomous assembly of two chaser spacecraft simulators and one spinning target spacecraft simulator. The mission process includes four distinct phases. While the first two phases use the real-time assembly and collision avoidance controllers proposed, the assembly control laws with analytical expressions play its role in the remaining phases. Both neighboring spacecraft simulators and external object serve as the obstacles during the design process. The hardware-in-the-loop experimental system consists of three spacecraft simulators with electromagnetic docking mechanism installed, a granite testbed, a motion tracking system and a wireless transmission system. The experimental results show that the proposed multistage control strategy can successfully realize the autonomous collision-free assembly of three spacecraft simulators in the spinning scenario. • The autonomous assembly of multiple spacecraft is experimentally validated on the air-bearing testbed. • The more challenging assembly of a spinning structure is innovatively selected. • The wide-range approaching operation, especially the collision avoidance problem with the external objects, is considered. • A novel APF-based attitude alignment controller and a new Lyapunov function is proposed. [ABSTRACT FROM AUTHOR]

A Robotic Active Debris Removal (ADR) mission requires close-proximity operations for capturing potentially tumbling, uncooperative targets and disposing of them. These operations rely on accurate relative navigation between the chaser and the target to safely perform the mission. In the case of a known piece of debris that has not been designed to be serviced or disposed of, the knowledge of its shape factor allows relative pose determination. In the case of unknown space debris, however, no relative pose determination is possible as no reference is available beforehand. In this paper, we derive an onboard software architecture for estimating the shape and dynamical model of an unknown, uncooperative space debris, and perform tracking, using features detected in the image and the associated depth. Firstly, an Error-State Kalman filter (ESKF) is used to estimate the chaser trajectory and attitude as well as the target shape in an arbitrarily fixed target frame. The next step in the architecture is to apply a Gaussian Processes (GPs) algorithm to model the rotational dynamics of the debris, allowing an accurate attitude propagation independently of its tumbling mode. Finally, a second ESKF is instantiated using existing knowledge to track the position and attitude of the target. This work assumes the pre-existence of a feature detection, tracking and matching algorithm, and fusing data from a depth sensor allowing the retrieval of pixel coordinates and depth for each detected feature. The proposed architecture has been tested and validated in a computer-simulated environment with parameters representative of a real-world ADR mission. Results from over 100 simulations show a mean position and attitude error of 0.3 m and 1.5°, respectively. • Dealing with an unknown space debris requires building knowledge on the go. • A multi-stage architecture is presented for onboard estimation of such objects. • A machine learning algorithm is used to predict the tumbling of the debris without inertia. • Simulations shows a position and attitude accuracy under 1 m and 2° respectively. [ABSTRACT FROM AUTHOR]

Highly porous aluminum alloys and honeycomb structures may be utilized as single-use shock absorbers in landing gear and other elements of spacecraft thus ensuring safety of the landing process. This paper presents the results of studies that make it possible to estimate mechanical properties and shock-absorbing characteristics of highly porous aluminum alloys and honeycomb aluminum alloy structures, the initial average densities of said alloys and structures ranging from 330 to 2250 kg/m³. We list analogs that correspond with different stages found on stress-strain curves of highly porous aluminum alloys and honeycomb structures. The paper also provides the foundations of calculating the specific and effective energies of shock absorption. • The deformation diagrams of cellular aluminum are close to the idealized one. • Specific and effective impact absorption energy calculated by the deformation diagram. • Some honeycomb structures have a deformation diagram with hardening. • The presence of hardening worsens the shock-absorbing characteristics. • The deformation diagram depends on the porosity, matrix material and pore structure. [ABSTRACT FROM AUTHOR]

To progress towards as quick as possible feasible pre-design of space missions, spacecraft early design today is organized gathering together experts from different spacecraft subsystems that share information (and results). This Concurrent Design (CD) approach needs accurate but quick solutions from all subsystems involved in each phase of that mission early design. At IDR/UPM Institute researchers have been working in simplified models for the spacecraft power distribution subsystem. In the present paper three models, for solar panels, the battery and DC-DC converters are described and tested with experimental data. The models were designed to be coupled within the simulation of a spacecraft mission. The model regarding the solar panels is focused on accurately estimate the power production from the panels, whereas the model of the batteries is based on the discharged energy rate as the main/control parameter instead of the discharged/charged ampere·hour rate. Finally, the solar-panels/battery coupled design is described in this paper, the UPMSat-2 mission being used as case study. • The testing results related DC-DC spacecraft converters are used to model these parts. • Spacecraft solar panels are modeled with simplified approaches based on information from the manufacturer. • A simple model for Li-ion battery performance is proposed. • The power system of the UPMSat-2 spacecraft mission is simulated. • DET and MPPT are compared, DET being more efficient for the UPMSat-2 mission. [ABSTRACT FROM AUTHOR]

In 2017, Technische Universität Berlin (TU Berlin) has launched the first space-rated fluid-dynamic actuator (FDA) on the TechnoSat mission. This type of actuator is unique with respect to its dynamic properties, as it offers a very high torque but limited angular momentum storage capabilities in comparison to reaction wheels (RWs). In combination with the tetrahedron RW assembly featured on the spacecraft, a unique combination of attitude control actuators has been formed. The spacecraft axis parallel to the FDA is over-actuated, featuring control actuators of contrasting nature which offers novel options in terms of spacecraft operations. However, with the camera axis on TechnoSat being aligned parallel to the actuator's axis of rotation, there is limited operational use to this setup. In July 2019, TU Berlin launched Berlin Experimental and Educational Satellite 9 (BEESAT-9) – the first CubeSat featuring a set of three orthogonal RWs and a single picosatellite fluid-dynamic actuator (pFDA) for attitude control. The camera axis of BEESAT-9 is perpendicular to the actuator axis, and therefore allows the demonstration of the proposed novel modes of operation. Among the proposed modes is an artificial increase of the swath and the acquisition of stereo images of multiple, successive targets on the ground. Crucial to these operation scenarios is the control allocation between the RW assembly and the single pFDA. This paper first introduces the spacecraft and describes the proposed novel modes of operation that are enabled by the actuator combination realized on BEESAT-9, followed by a review of the state of the art of control allocation for over-actuated spacecraft. The main part of the paper comprises a description of the implementation of the dynamics models of the attitude control actuators and spacecraft, followed by a description of the implemented high-level control laws and control allocation methods. Finally, simulation results for the artificial swath increase and the single spacecraft stereo imaging mode are appended and discussed. • Picosatellite fluid-dynamic actuators enable novel modes of satellite operation. • Control allocation enables combination of novel actuator and reaction wheels. • First CubeSat with fluid-dynamic actuator and reaction wheels launched in 2019. • Attitude maneuvers in the orbit plane enable stereo imaging from single spacecraft. • Attitude maneuvers out off the orbit plane enable artificial swath increase. [ABSTRACT FROM AUTHOR]

In deep space exploration, the application of low-thrust propulsion technology and multiple gravity-assist technology will greatly improve the scope of spacecraft exploration but will significantly increase the difficulty of trajectory design. This study proposes a fast initial design method of low-thrust trajectories, which reasonably combines low-thrust propulsion and multiple gravity-assist technology. In this paper, the Bezier shape-based method is combined with the gravity-assist algorithm to obtain a fast algorithm for calculating the low-thrust multiple gravity-assist transfer trajectories. For the design of the Jupiter rendezvous trajectories, the feasibility of the proposed method is proved by comparing several calculation examples with and without gravity assistance. In order to prove the effectiveness of the proposed method, it is compared with the finite Fourier series (FFS) shape-based method. The simulation results show that the proposed method can use shorter computation time to obtain transfer trajectories with smaller velocity increments than the FFS shape-based method. In order to prove the superiority of the proposed method, the results obtained by the proposed method are used as initial values for further optimization of the Gauss pseudospectral method (GPM). The simulation results show that the difference between the performance indexes of the Bezier method and the GPM is only about 2%, while the Bezier method only uses approximately 2% of the computation time of the GPM. The method described in this paper is significant for obtaining a feasible initial solution quickly for a large number of design cases. • A fast algorithm for calculating the low-thrust multiple gravity-assist transfer trajectories is proposed. • Low-thrust Earth–Jupiter transfer using multiple gravity-assist scheme is studied. • Four sets of trajectories are designed, and the influence of launch window on transfer fuel consumption is analyzed. • The feasibility of using the results obtained from the fast trajectory calculation method for the direct optimization algorithm is evaluated. [ABSTRACT FROM AUTHOR]

Space debris detection and characterisation is an important and pressing field, as the amount of space debris is growing due to the increase in spacecraft launches and increasing interests for space uses. Space agencies like NASA and ESA keep track of over 5,000 spacecraft and 20,000 spacecraft-related debris objects currently orbiting the Earth. An active spacecraft may undergo trajectory changes through an impulsive maneuver, or an inactive spacecraft may change its trajectory because of uncontrolled explosions due to residual fuel without warning. These pose risks to operational satellites. In addition, measurement data from tracking and detection consists of uncertainties, causing maneuver detection and characterisation methods to become computationally expensive. Thus, the goal of this work was to develop a quick, reliable, computationally inexpensive method for spacecraft maneuver detection and characterisation to be used by satellite operators and assist in collision avoidance measures. This paper presents such a technique, which is applied to simulated measurement data of an orbiting spacecraft taken from one ground site. With these objectives, a method using a modified Extended Kalman Filter (EKF) with covariance inflation from previous studies was explored and simplified. A new parameter definition is presented for maneuver detection and characterisation, named the Kaderali-parameter, or K , with focus on transverse apogee and perigee maneuvers. A relationship was then found between the parameter and the magnitude of the impulsive artificial maneuver. The method was tested in several different scenarios, varying Δ v magnitude and scale, Δ v direction, maneuver orbital location, orbit geometry, noise covariance, detection parameter threshold, state covariance inflation threshold, and maneuver duration. In general, it was found that as the Δ v magnitude increases, the magnitude of the detection parameter also increases. This relationship was found to be a nonlinear logarithmic relationship through best fit curve testing. Possible extensions of this work in the future include: testing multi-maneuver scenarios, multi-sensor data collection and fusion, break-up or berthing events, rendezvous or mating events, and matlab EKF protocol implementation. • Unknown spacecraft maneuvers pose collision risks, leading to a Kessler Syndrome. • A computationally inexpensive one-parameter Kalman Filtering solution is utilised. • New parameter, K , is created to clearly detect and characterise impulsive maneuvers. • An analytical relationship between K and the maneuvering speed is established. • The relationship describes a nonlinear direct relationship to be used by operators. [ABSTRACT FROM AUTHOR]

Recent studies demonstrate the possibility of navigating in proximity of uncooperative space resident objects by using only monocular images. Despite the results achieved, the development and testing of new algorithms are strongly constrained by the availability of spaceborne image datasets. To overcome this, a new algorithm embedded in a tool to generate synthetic high-fidelity spaceborne image datasets is presented here. The architecture developed can be tailored to a wide range of scenarios and it is based on an open-source ray-tracing software. All assumptions and simplifications adopted are discussed in detail for the different models considered, including a trade-off between accuracy and rendering time. The new method described is subsequently adapted to a baseline scenario where the optical properties of a reference spacecraft model are tuned. Both qualitative and quantitative validations are detailed and successfully carried out for the baseline case, demonstrating the high photo-realism achievable with the proposed method. As a consequence of this main outcome, the paper details the generation of labeled spaceborne image datasets publicly available and reports the analyses that confirm the high level of representativeness, making them suitable for training and testing image-based navigation algorithms. As another outcome, the most comprehensive multi-purpose labeled dataset of validated spaceborne synthetic images currently publicly available is presented. • New methodology to generate spaceborne synthetic images via open-source ray-tracer. • Detailed analysis of atmosphere and topology modeling for celestial bodies. • Definition of a qualitative and quantitative validation step via image comparison. • Successful validation of the proposed approach and generation of validated datasets. • Detail of the largest multi-purpose labeled spaceborne image dataset available. [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]

The paper analyzes numerically the development of the freely expanding flame before the transition to detonation in three-dimensional space. The risks related to the transition to detonation in unconfined space arise at launch places and near-Earth space during rocket engine accidents. Clear and adequate models for such scenarios are in demand for the elaboration of safety measures. Here it is shown that the process of transition to detonation involves multidimensional effects related to the generation of compression waves and their amplification when interacting with the developing flame. The necessary conditions for deflagration-to-detonation transition are flame acceleration up to near-sonic speed and generation of transversal compression waves (propagating along the flame front). So, while the subsonic stage of flame acceleration can be successfully described by the quasi one-dimensional model supplemented by a known self-similar law of the flame area growth (folding factor), the model for the pre-detonation stage should take into account considered multidimensional effects. The presented results imply that there is no unambiguous relation between the transition to detonation and the folding factor. • Three-dimensional flame propagation in unconfined volume is studied. • Local structure of the flow is analyzed. • The folding factor is obtained for free flame directly before transition to detonation. • No unambiguous relation between the transition to detonation and the folding factor is found. [ABSTRACT FROM AUTHOR]

Orbit determination (OD) for spacecraft with unknown maneuver is always a challenging task. This paper proposes a novel framework for efficiently solving continuously maneuvering spacecraft OD problems by merging the Long Short-Term Memory (LSTM) neural network and the filter algorithms. A polynomial-representation fitted the unknown continuous maneuver is first proposed. Rather than directly output the estimated maneuver, the LSTM is trained to detect the unknown maneuver and then estimate the coefficients of the polynomial-representation. Then a fusion is designed to combine the prediction of the LSTM and the estimation of the filter for a more accurate estimation. The proposed LSTM-based framework is successfully applied to solve a Low-Earth-orbit OD problem and a Middle-Earth-orbit OD problem. The continuously maneuvering target is well tracked and the unknown maneuver is accurately estimated. Numerical simulations show that the LSTM trained based on one training dataset can also be applied to other scenarios that share some common features with the training dataset. • A framework for solving maneuvering spacecraft OD problems is proposed. • A polynomial-representation is used to fit the unknown maneuver. • A fusion strategy is designed to combine the predictions of the LSTM and estimations of the filter. [ABSTRACT FROM AUTHOR]

Robotic on-orbit servicing presents the opportunity for repairing and refueling satellites and deorbiting space objects. Performing such tasks is challenging because they must be done autonomously and accurately with high robustness. Ground testing, such as air-bearing-based spacecraft simulators, is widely used to develop and validate the design and control techniques of space robotics required by modern spacecraft applications. This paper presents the development of planar air-bearing microgravity simulators with a 3-DOF robotic manipulator for autonomous spacecraft tracking, rendezvous, and capture control of a free-floating target. The system includes a pseudo-galactic star tracking system for pose determination and navigation of spacecraft simulators. A second optical tracking module was also developed to determine the pose of a target spacecraft relative to the chaser spacecraft. Path planning and PD trajectory tracking algorithms were developed for spacecraft tracking and rendezvous. The experiments successfully demonstrated that the testbed could be used to study autonomous spacecraft rendezvous and capture a tumbling spacecraft by a robotic manipulator in a space-like testing environment. • Developed air-bearing spacecraft simulators for spacecraft rendezvous control. • Developed air-bearing lifted robotic manipulator for capturing a spinning target. • Tested autonomous path planning and autonomous robotic capture by the system. [ABSTRACT FROM AUTHOR]

The propulsive characteristics of an Electric Solar Wind Sail are usually evaluated using a simplified model in which all the sail tethers are coplanar and form a sort of rigid disk. However, the three-dimensional arrangement of the tethers is fundamental information in the study of the spacecraft performance, and must be accounted for in refined mission analyses. In this paper, a Finite Element approach is chosen to estimate the deflected shape of the tethers, thus allowing important information on the structural response of the sail to be obtained. A parametric code is developed to perform a static analysis of an Electric Solar Wind Sail, whose requirements are given in terms of payload mass and spacecraft characteristic acceleration. In particular, the tether structural response is investigated using three different beam models, which are compared in terms of accuracy and computational efficiency. The analysis is specialized to the noteworthy case of a Sun-facing sail that is placed at a distance of one astronomical unit from the Sun. The numerical results, which concern a set of possible sail configurations, are compared with those taken from analytical models. • The paper analyzes the E-sail deformed shape using a finite element approach. • A parametric procedure is proposed to perform the E-sail static analysis. • The tether structural response is investigated using three different beam models. • The numerical results are compared with those taken from analytical models. [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]

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]

Electronic components can be seriously damaged during the launch of spacecraft due to the intense shock loads, which are generated by the launch operations such as the separations of the launcher stages and the release of the spacecraft. A shock verification program should be carried out in the development phase of space instruments to guarantee that the parts most susceptible to shock damage, including electronic components, can withstand the severe shock environment. A new shock verification methodology is described in this paper, where the novelty is the utilization of the numerical simulations to verify the capability of space instruments to withstand the specified shock loads. The proposed procedure has been applied to the units of the Energetic Particle Detector (EPD) payload of the Solar Orbiter mission. The selected approach is not the usual method in the space systems for shock verification, where the preferred method is by testing. Therefore, this work represents one of the few documented cases where numerical simulations have achieved the demonstration of the shock capability of space structures without the need to perform a complete shock test campaign. The application of the numerical analyses for the shock verification of one of the EPD units is explained in detail, where the calculated responses such as maximum expected stresses, accelerations and relative displacements are compared to the allowable limits to demonstrate that this instrument and its sensitive electronic components can adequately support the intense shock loads during the launch phase. • Proposal of a new methodology to verify the shock effects by numerical analysis. • Description of the shock verification for the EPD payload of the Solar Orbiter. • The novelty is that numerical FEM analyses achieved the final shock verification. • The detailed description of each task and the results are presented in the paper. • The focus is on the shock verification of the electronic components of these units. [ABSTRACT FROM AUTHOR]