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Based on the increasing need for higher data rates in CubeSat missions primarily for Earth observation and communication missions mainly focused on Direct-to-Earth (DTE) applications, DLR started the development of a highly compact laser communication payload for CubeSats, called CubeLCT. In parallel TESAT has used its experience for preparing and developing this technology for volume production. At the same time, to ensure the compatibility with the ground stations, DLR started and led the CCSDS working group for Optical On-Off-Keying (O3K) to further extend the commercial use of CubeLCT terminals. This cooperation between industry and research center has been very successful. The PIXL-1 mission was accomplished in a very agile new space approach, developing and demonstrating the technology in record time. Moreover, in order to enable the potential customers to deploy a complete end-to-end system, TESAT in cooperation with German Space Operation Center (GSOC) has specified the interface requirements for developing the first demonstrator for DTE optical system mission planning, mainly focused on the link planning challenges. This concept can be applied to current missions, and allows identifying the needs also for further applications. In addition to DTE links, the increasing demand for higher data rates and low latency communication in upcoming constellations of CubeSats also drives the need to extend the CubeLCT terminal with Inter-Satellite-Link-functionality. In continuation to the first development, IKN, RSC3 and TESAT, are currently developing a CubeLCT for intra-plane communication capable of transferring 100 Mbps over 1800 km. The upcoming challenges in operation of inter-satellite-links on CubeSats combined with the possibilities of enabling DTE are also covered in this paper. • Highly compact (1/3 U) laser communication payload for CubeSats. • Mainly focused on Direct-to-Earth (DTE) applications. • Design compatible with CCSDS working group for Optical On-Off-Keying (O3K). • Agile DTE link planning including local weather information. • Technology ready for industrial volume production. [ABSTRACT FROM AUTHOR]

OrigamiSats, a new concept in solar sailing, are origami spacecraft with reflective panels that, when flat, operate as a conventional solar sail. Shape reconfiguration, i.e. "folding" of the origami design, allows the OrigamiSat to change operational modes, performing different functions as per mission requirements. For example, a flat OrigamiSat could be reconfigured into the shape of a parabolic reflector, before returning to the flat configuration when required to operate again as a solar sail, providing propellant-free propulsion. The attitude dynamics and shape reconfiguration of OrigamiSats are known to be highly coupled, thus presenting a challenge from a control perspective. This paper investigates the problem of integrating attitude and shape control of a Miura-fold pattern OrigamiSat through the use of variable reflectivity, allowing differences in solar radiation pressure to be used to enact shape reconfiguration and attitude manoeuvres. A closed-loop feedback controller is presented which combines and balances the attitude and shape control requirements, and gain-scheduling is implemented to address some specific features of the system dynamics. Numerical simulations of the multibody dynamics of the system are used to test the proposed controller and simulations of some example manoeuvres are performed which demonstrate the system's performance. • OrigamiSats are a new concept in solar sailing. • Solar radiation pressure can be used to "fold" an origami solar sail. • An integrated attitude and shape control law is developed. • Gain-scheduling is used to address some nonlinearities of the system. • Attitude and shape control is demonstrated through numerical simulations. [ABSTRACT FROM AUTHOR]

Space domain systems must go through different types of ground testing. For system-level black-box functional testing, ground support equipment is built ad hoc, customized for each different mission. Building any of this specific equipment involves considerable engineering effort that can be diluted using techniques that allow reusing configurations of the test setup environment without recodification. These techniques can be merged into a hierarchical checking process based on independent filters to compare the received outputs with the expected ones. The hierarchical checking allows the separate definition of several layers or levels of validation, from the lowest protocol packet levels to the logical abstractions at the highest application level. This paper introduces a solution based on Model-Driven Engineering for Low-Code Ground Support Equipment. It uses the abovementioned mechanisms to reduce the effort to develop and customize ground support systems for different missions. This solution is integrated within an environment called MASSIVA, which allows the automatic configuration and definition of system-level test procedures. This environment has been used in the software verification and validation process of the Instrument Control Unit of the Energetic Particle Detector on board Solar Orbiter. • Low-Code solution for Ground Support Equipment development. • Model-Driven Engineering support for black-box system-level test definition. • Automatic checking process based on hierarchical definition of independent filters. [ABSTRACT FROM AUTHOR]

Tether-nets are one of the most promising approaches for active removal of space debris. This paper investigated the collision dynamics of flexible nets with different configurations, the focus is on the selection of an optimal configuration with respect to collision dynamics. Firstly, the design method of net configuration is proposed based on undirected graph theory, and five flexible nets with different configurations are designed, including quadrilateral, pentagon, octagon, dodecagon and circle. Secondly, the collision dynamics model is presented and its effectiveness is verified via numerical simulations and ground experiments. Finally, collision dynamics of the five flexible nets are simulated based on the verified model and the differences of these nets are compared and analyzed. Comparison results show that the flexible net with octagon configuration has better collision dynamics than others according to the maximum deformation, stress, and energy dissipation during the collision process, which provided an optimal configuration of flexible nets for space debris removal. • A design method of net configuration based on undirected graph is proposed. • The collision dynamics model of flexible nets is proposed and verified. • The octagonal net has the best dynamic characteristics. [ABSTRACT FROM AUTHOR]

Knowledge of the thermospheric density is essential for calculating the drag in low Earth orbit satellites. Existing models struggle to predict density accurately. In this paper, we propose thermospheric density prediction using a deep evidential model-based framework that incorporates empirical models, accelerometer-inferred density from the CHAMP satellite, and geomagnetic and solar indices. The framework is investigated on both quiet and storm conditions. Our results demonstrate that the proposed model can predict the thermospheric density with high accuracy and reliable uncertainty in both quiet and storm times. The predicted results from the evidential model are advantageous over the Gaussian Processes (GPs) model in our previous studies. Furthermore, the proposed model can also provide insightful aleatoric and epistemic uncertainties. • Predict thermospheric density with a deep evidential model-based framework. • The method is investigated on both quiet and storm periods. • The density predictions are accurate with reliable uncertainty in all conditions. • The method can provide insightful aleatoric and epistemic uncertainties. [ABSTRACT FROM AUTHOR]

In this paper, Particle Swarm Optimization with energy-to-fuel continuation is proposed for initializing the co-state variables for low-thrust minimum-fuel trajectory optimization problems in the circular restricted three-body problem. Particle Swarm Optimization performs a search of the solution space by minimizing the weighted sum of squares of the two-point boundary-value problem final boundary condition residuals for the minimum-energy problem. Next, an energy-to-fuel homotopy is employed to transition the minimum-energy trajectory to a minimum-fuel trajectory, starting from the generated guess. The proposed methodology is applied to two low-thrust transfer problems in the Earth–Moon system: a transfer from a geostationary transfer orbit to an L1 halo orbit, and a transfer from an L2 halo orbit to an L1 halo orbit. The resulting minimum-fuel trajectories are validated with the literature. It is demonstrated that the methodology can successfully generate guesses for the initial co-state variables which converge to a solution for both scenarios. A strategically chosen particle swarm size is shown to improve the efficiency of the methodology. The proposed approach is of simple implementation, can be easily extended to other trajectory optimization problems, facilitates the discovery of multiple candidate trajectories, and does not require a user-provided guess, all of which are advantageous features for the preliminary phase of mission design. • Particle Swarm Optimization is proposed to guess the Lagrange co-states variables. • PSO initialization is paired with homotopy continuation for trajectory optimization. • The proposed algorithm facilitates discovery of multiple unique optimal trajectories. • Two minimum-fuel trajectory optimization problems are solved. • The effect of the particle swarm size on successful convergence is analyzed. [ABSTRACT FROM AUTHOR]

This paper proposes a robust finite element Kalman filter for the state estimation of flexible space tether systems, which are critical for space exploration and satellite missions. Due to the limited access to placing sensors along the flexible tether to measure them directly, the measurement of the state of these systems is challenging. To address this, a fully expandable, scalable, and robust approach is developed that fuses position/velocity sensors at tether ends, distributed fiber optic strain measurement along the tether, and a position/velocity estimation by a finite element-based Kalman filter to estimate the state of the tether system. The proposed estimator is mathematically proven to be observable, stable, and robust. The effectiveness of the estimator is demonstrated through numerical simulations of a space tether system in orbit, and the method can be extended to any flexible body system. The results of the simulations show that the proposed approach can accurately estimate the state of the tether system, even in the presence of disturbances and noise. • Expandable, scalable FE-Kalman state estimator for flexible space tether systems. • Data fusion with tether end position/velocity, distributed fiber optic strain measure, and estimated internal position/velocity. • Mathematically proven observability, stability, and robustness to noise of the estimator. [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]

Precise regulation of oxidizer mass flow rate is a crucial technology for hybrid rocket motors, and a flow control valve serves as a core component of an oxidizer delivery system. In this paper, we designed a variable area cavitating venturi (VACV) as a flow control valve to control oxidizer mass flow rate. Three types of experiments were carried out to investigate VACA-based flow feedback control and its application in hybrid rocket motors. Specifically, we measured flow characteristics at different pintle strokes and conducted five cold tests and one ground hot test based on flow feedback control. Test results are as follows: first, the VACA flow characteristic curve shows a wide regulation range with a flow regulation ratio of 21.8 and performs good linearity of 96.76%. Second, the mass flow rate curve measured with flow feedback control tracks the target curve. Flow feedback control effectively improves control accuracy by maintaining the steady-state error within 1.2%. Third, hot test results demonstrate that the VACA with flow feedback control can effectively maintain the accurate and consistent oxidizer mass flow rate during the firing stage of the hybrid rocket motor. The findings of this research indicate that VACA with flow feedback control can be used as an effective flow regulation component to maintain a constant and accurate oxidizer mass flow rate during the working stage of a hybrid rocket motor. • A firing test of a hybrid rocket motor based on feedback control was conducted. • VACV shows good linearity with a wide regulation range in flow characteristics curve. • Feedback control with VACV improves accuracy, keeping steady-state error within 1%. • Feedback control maintains an accurate and consistent oxidizer mass flow rate. [ABSTRACT FROM AUTHOR]

Experimental study is performed to investigate the dynamic response characteristics of the flame when global equivalence ratio increases suddenly during the condition transition in a cavity-based scramjet combustor. High-enthalpy supersonic flow enters the combustor on conditions that Mach number, stagnation temperature, and total pressure are 2.52, 1486 K, and 1.6 MPa, respectively. In this paper, more attention is paid to the characteristics of flame evolution after sudden increase of fuel injection pressure and combustion characteristic when the injection pressure gets stabilized again. Three flame modes after condition transition have been observed and analyzed according to the transient chemiluminescence image captured by high-speed photography. There are two kinds of stable flame modes after transition mission at a low injection pressure difference, which are mainly reflected in the distribution of flame and intensity of visible light. Whereas, once the injection pressure difference exceeds a critical value, intense periodic flame flashback phenomenon begins to appear in the combustor. It can clearly capture the back and forth motion of the flame in the combustor, along with the local blowout and reignition phenomenon. Furthermore, the dominant frequency of flame flashback decreases and contrast of combustion intensity becomes more obvious as the injection pressure difference increases. • The highlights of this article are listed in the following: • The dynamic response characteristics of flame when equivalence ratio increases suddenly during condition transition in scramjet are investigated experimentally. • Three flame modes after condition transition have been observed and analyzed. • The effect of injection pressure difference on combustion characteristics is explored. [ABSTRACT FROM AUTHOR]

The near-Earth asteroid 1998 KY26 is a rendezvous target of the Hayabusa2 extended mission, with a diameter as small as 20 – 40 m and a spin period as short as 10. 7 min. This research aims to develop a plan for Hayabusa2 operations around this particular asteroid constrained by the remaining spacecraft resources. The main part of this paper begins with the characterization of the orbital and surface dynamics about 1998 KY26. The analysis demonstrates that the orbital motion of the spacecraft in the microgravity environment is dominated by solar radiation pressure, while the centrifugal force due to the high spin rate plays a prominent role at the surface. The dynamical environment significantly differs from that of the original target Ryugu, for which Hayabusa2 was specifically designed, posing technical challenges in close-proximity operations. In particular, this article highlights the feasibility of two primary operations: hovering for station-keeping and controlled descent for close-up observations. The discussion is further extended to more advanced operations, such as those using the projectile and target marker aboard the spacecraft. Our research pushes the envelope to explore one of the smallest members of the Solar System. • Hayabusa2 ♯ target asteroid 1998 KY26 is a fast rotator with a diameter of ∼ 30 m. • Low gravity and rapid rotation form distinctive orbital and surface environments. • Station-keeping can be achieved by hovering 1 – 3 km above the asteroid. • Controlled descent to lower altitudes shows minimal guidance error and fuel cost. • Advanced options include operations using the projectile and the target marker. [ABSTRACT FROM AUTHOR]

The problem of relative drift elimination between the satellites in the swarm is considered in the paper. The proposed decentralized control takes into account a communication constraint such as limited size of communication area. Only the satellites within the communication area can be identified by relative motion determination system. The control aim is to eliminate the mean relative drift between all the satellites inside the communication area. The purpose of the work is to study the performance of the proposed decentralized control algorithm. It is shown that the system matrix of differential equations for the vector of relative drifts is related to the Laplacian matrix of the communication graph. In the case of the connected swarm, all but one eigenvalues of the system are negative, and the remaining one is equal to zero. It means that all the relative drifts converge to the same value under the proposed control. The speed of convergence is defined by the minimum absolute eigenvalue that depends on the graph topology. The initial drift and the convergence speed make it possible to estimate the communication distance that provides the connectivity of the graph. Considering normally distributed errors of the initial velocity after the launch, it is possible to estimate the distance between any two satellites after the convergence. It allows us to estimate the communication distance that ensures the relative drift elimination between all the satellites in the swarm. The obtained estimations are validated using Monte Carlo simulations. In numerical simulations the swarm of 3U CubeSats in low-Earth orbit is considered. The decentralized control is implemented by differential aerodynamic drag via the change of cross-sectional area using onboard reaction wheels. • Problem of relative drift elimination between satellites in the swarm is considered. • The proposed decentralized control takes into account a limited communication area. • Estimation of the communication area size that guarantees convergence is derived. • Analytical estimation of communication distance is obtained using graph theory. • Performance of swarm control by differential drag is studied numerically. [ABSTRACT FROM AUTHOR]

In this study, the effect of a velocity-pulsating stoichiometric hydrogen-oxygen mixture entering an expanding combustion chamber on the detonation wave is investigated. The hybrid sixth-order weighted essentially non-oscillatory centered difference scheme was used to solve the Navier–Stokes equations with a one-step two species chemistry model. The results verify that the induction zone is extended under the influence of pulsating flow. The discontinuous structure is formed in the flow field. A lot of unburned jets between the wavefront and the reaction zone were generated, which resulted in insufficient energy release. Coupled with the weakening of the wave strength by the Prandtl–Meyer (P-M) expansion fan, the detonation wave show different degrees of attenuation. The decay of the detonation wave shows different patterns at different amplitudes and periods. When the velocity amplitude is 0.3 V C J , the decay rate of detonation wave increases with the increase of the disturbance period. This attenuation law is different from the amplitude of 0.1 V C J and 0.2 V C J . This is because the discontinuous structure area formed by high-amplitude disturbance is large. Therefore, the longer its existence time is, the more serious the detonation wave attenuation is. • This paper is an innovative study on the propagation of detonation waves under pulsating velocity inflow. • Unburned jet generated by the expansion channel weakens the intensity of the detonation wave. • The propagation law of detonation wave varies with the period of velocity disturbance under different amplitudes. • Under the high-amplitude disturbance, the discontinuous structure impacts the propagation of detonation wave. [ABSTRACT FROM AUTHOR]

This paper presents the methodologies to design the transfer trajectories from Earth nearby to heliocentric orbits. For multiple spacecraft with the same initial condition but aiming to different phases in the heliocentric orbit, several lunar gravity assists are used to divert them into different interplanetary orbits. Firstly, based on the patched-conic approximation, a complete design process for deep-space transfer orbits was proposed for multiple spacecraft performing lunar gravity assist maneuvers. In this design process, it was assumed that each spacecraft had the same initial conditions. The periodic matching method was introduced to design the interplanetary transfer orbit, satisfying the initial and terminal constraints. An iterative method was proposed to search lunar gravity assist parameters that satisfied the hyperbolic excess velocity condition for the escape trajectories. Furthermore, a semi-analytical correction method was proposed to further approximate the precise transfer trajectory. This correction method was found to have a reduced time cost and a better error convergence compared with conventional correction methods. • Design of deep-space transfer trajectories for multiple spacecraft with identical initial conditions. • Periodic matching analysis of heliocentric orbit design. • Iterative method for searching lunar gravity assist parameters for spacecraft. • High performance recorded for the proposed semi-analytical correction method. [ABSTRACT FROM AUTHOR]

The electrohydrodynamic (EHD) actuator system, which produces thrust by ionizing a gas and accelerating it, is a novel propulsion system for near space vehicles. In this paper, an analytical model for EHD actuators, called EHD-AM, has been presented to estimate the thrust produced by the actuators in a wire-to-cylinder configuration. Experimental data from six previous studies have been used to validate the model and the mean absolute percentage error (MAPE) between each configuration and EHD-AM has been calculated under different environmental conditions. Based on the evaluations, the MAPE between EHD-AM and the experimental results for different pressures and wire-to-airfoil structures is 7.18% and 6.16%, respectively. The MAPE between EHD-AM and the experimental data measured by particle image velocimetry (PIV) method and with conventional experimental setups is 10.59% and 6.76%, respectively. Based on the EHD-AM, at a constant voltage, the relation between thrust and pressure has a local minimum and a local maximum point in the range of 0.05–1 atm. From 0.05 atm to 0.5 atm, as the pressure increases, the thrust decreases and then increases. As the pressure increases from 0.5 to 1 atm, the thrust decreases. In addition, a higher voltage can result in a higher pressure at which thrust is maximized. Furthermore, the generated thrust is monotonically increasing as the voltage amplitude increases. By implementation of EHD-AM thrust scale in a simple numerical model, the thrust generated by the actuator has been computed and simulated numerically. Initially, the numerical model was verified using experimental results from a previous study. The MAPE between the numerical model and the EHD-AM is 24.08%. Finally, when the interelectrode gap, the collector radius and the emitter radius are relatively large, the thrust of the actuator is predicted by EHD-AM and compared with the numerical results. EHD-AM can be used in the design and optimization of EHD actuators, and also for controller design and related analyses in a control system. • In EHD actuator, the relation between the thrust and the pressure is determined. • The relationship between corona current and voltage has been considered based on the space-charge-limited current relations. • In high and low mobility of ions, the thrust is obtained as a function of voltage. • Analytical model for corona discharge-based EHD plasma actuator is obtained by incorporating environmental conditions. [ABSTRACT FROM AUTHOR]

This paper explores the operational performance parameter range for inertial confinement fusion propulsion systems under the assumption of best case and worst case mission scenarios. This includes the spacecraft thrust, jet power and specific power. We show the derivation of the key driving equations and then simplify these with approximations for ease of domain space analysis. This includes a consideration of minimum and maximum values for jet efficiency, momentum weighting factor, nozzle divergence angle, fuel burn-up fraction, capsule mass and detonation pulse frequency. It is estimated that the range of parameter values may be of order 0.001 N–100 MN thrust, 1000 W–1,000 TW jet power, 0.01 W/kg–10 GW/kg specific power. Although this range includes values that are likely outside the realistic design space of applicability and to show this data from published designs is also examined. The analysis is considered for reaction isotopes of low mass hydrogen and helium nuclei only and does not take into account the possibility of propellant expellant augmentation, vehicle staging, design optimality or enhanced fuel reactions. • Attempt to examine inertial fusion propulsion theory from a broad perspective. • Derivation of driving equations for various inertial fusion propulsion parameters. • Analysis of min-max range inertial fusion propulsion performance space missions. • Analysis of inertial fusion propulsion performance data from concept designs. • Map of fusion propulsion parameters to derive the requirements for missions. [ABSTRACT FROM AUTHOR]

It is common practice in industry to deliver a Craig–Bampton reduced version of the spacecraft's finite element model to the launch authorities to perform the dynamic coupled load analyses. During the dynamic coupled load analyses, the finite element model of the spacecraft is coupled with the finite element model of the launcher to predict the responses caused by the vibration environment produced during the launch phases (e.g., lift-off, engine start-up, etc.). Before being Craig–Bampton-reduced, the finite element model needs to be correlated and validated against physical test results to guarantee its capability to reproduce the actual behaviour of the physical hardware, which is essential to perform a meaningful dynamic coupled load analyses. Given the complexity and size of the finite element model, the entire process can be very demanding. In this paper, an alternative to the mathematical models usually used in the dynamic coupled load analyses is investigated. The main idea is to synthesize a Craig–Bampton equivalent model directly from the response data of a vibration test experiment. The system identification problem is solved using a particle swarm optimization algorithm, tailored specifically for this type of application, the hybrid PSO and Local Search (hPSO-LS). A mutation operator from genetic algorithm is used to improve the solution and the local search method to refine the final result. The proposed method is numerically tested on a 5-degrees of freedom lumped mass system, with full stiffness and damping matrices. A sensitivity analysis of the search space size is also carried out to investigate its influence on the final results. The hPSO-LS is tested in both noise-free and noisy scenarios. • Hybrid variant of PSO with refined research is proposed to solve system identification problem. • Reconstruction of full stiffness and damping matrices of a lumped mass system from accelerations only. • Sensitivity analysis on the size of the search space is carried out. [ABSTRACT FROM AUTHOR]

The purpose of this paper is to develop an effective method for onboard absolute orbital elements estimation of the non-cooperative target, which may beneficial to collision avoidance. The proposed estimator consists of two parts: the target acceleration observer and the orbital elements filter, where the observer generates partial input for the filter. For the estimator, the data source comes from the orbit determination information of the obstacle-avoidance-satellite and the relative measurement information of the non-cooperative target. These data are fed into an observer to estimate the maneuvering acceleration of the target. Then, the square root cubature adaptive H ∞ information filter (SCAHIF) for target-orbital-elements estimation under the complicated noise condition is designed, which is driven by the observer outputs and the target data source. The performance of the overall "observer-enhanced SCAHIF" (OSCAHIF) is evaluated through three scenarios where the target is under zero maneuvering thrust, constant maneuvering thrust, and time-varying maneuvering thrust. Test results indicate that the proposed OSCAHIF provides a feasible solution to real-time absolute orbital elements estimation of the non-cooperative target with the presence of time-variant mixed Gaussian white noise. • OSCAHIF is proposed for orbit element and thrust estimation of noncooperative target. • An observer is developed for maneuver acceleration estimation. • The adaptive H-infinity method is introduced to improve the robustness. • The observer, square root cubature IF, and adaptive H-infinity method forms OSCAHIF. [ABSTRACT FROM AUTHOR]

In this paper, an efficient algorithm for correlating the transient behaviour of thermal lumped models to reference data is reported. The algorithm is based on an iterative non-linear least-squares minimization process that allows for fast convergence. The correlation process requires the calculation of the so-called Moore–Penrose pseudoinverse of the Jacobian matrix. However, obtaining the transient Jacobian matrix can be computationally expensive for large models. A new method for calculating the transient Jacobian matrix of lumped thermal models, called Jacobian Propagation, is also presented. For large models, the Jacobian Propagation method, when combined with an implicit solver, has a negligible computational cost compared to the solving process. The correlation algorithm and the Jacobian Propagation method have been successfully tested with two thermal models, obtaining better performance than previous approaches. • Efficient transient correlation algorithm for lumped parameters thermal models. • Use of Jacobian matrix formulation for high-convergence rate. • A new method for calculating the transient Jacobian matrix. [ABSTRACT FROM AUTHOR]

NASA's Deep Space Network (DSN) serves as a critical element in the exploration of deep space, providing telecommunication services such as tracking, telemetry, command, and delivery of science data to its customers around the globe. In addition to supporting robotic spacecraft for scientific missions, the DSN also supports human spaceflight (HSF) endeavors, including NASA's flagship Artemis missions to the Moon. This paper describes DSN lessons learned during preparations for the first Artemis mission, Artemis I. Topics covered include DSN performance and reliability, network loading, interface management, international collaboration and the role of international partners. Lessons learned will be applied to future missions, including Artemis II and beyond. • NASA is managing projected growth in antenna demand and improving supportability to customers. • NASA is working on mission support enhancements to enable a sustained human presence on the Moon. • DSN antenna scheduling processes and downtime planning are designed to meet mission needs. • DSN support for human spaceflight necessitates years of coordination across many entities. • DSN coordination with international partners provides support for human spaceflight endeavors. [ABSTRACT FROM AUTHOR]

Although humans are destined to live in a world of chaos, there is also hidden beauty within it. Analogically, this implies that there can be a creative way of living through the exploration of the unknown. Considering humans are part of the dynamism of the whole universe that consists of matter and energy, what would it be like if they could feel being embraced through the universe? Having continuous dialogues with the organizer throughout the planning phase, an artistic performance with the sounds of the universe is presented at the cultural event Connected Ink 2021 in Tokyo. This performance is entitled Life, the Universe and Everything in homage to The Hitchhiker's Guide to the Galaxy by Douglas Adams. In this paper, the background, concepts, physics, sound design, and the process and result of the project are discussed in collaboration with artists, sound designers, and experimental physicists. Regarding the sound design, it had been developed, in part, based on what the authors have shown through the preliminary sound the installation: Particle Post – Letters from the Universe in 2019, through which the signals from cosmic muons that are captured by a scintillation detector are poetically designed. Transforming imperceivable elements of the universe into an appreciable form in a poetic manner is one of the crucial essences of the authors' research and practice. With this, we believe that we could approach the wonders of the universe aesthetically. Hence, imperceivable elements such as cosmic muons and the parameters of solar wind (density, speed, and temperature in particular) captured from the satellite DSCOVR orbiting the L1 Lagrange point, are brought into the audio system in almost real time. Five iron fuurin , Japanese wind bells, ring triggered by cosmic muons that constantly arrive on Earth, while the stylized sounds of solar wind surround the large space at the venue. Consequently, these two different kinds of 'wind' sounds create harmony within the space, hoping to create a similar harmony through the bodies and minds of the audience as well. This project also signifies the value of genuine dialogue and communication with the purpose of deepening an interdisciplinary collaboration and opening up further creation. • We explore this question: how can humans share the wonders of the universe by being embraced through the sounds? • The signals from cosmic muons and the parameters of solar wind are brought into the audio system almost in real time. • This collaborative project signifies the value of genuine dialogue and communication in the process of creation. [ABSTRACT FROM AUTHOR]

This paper deals with the attitude motion of VLEO CubeSats during and after the deployment of flexible stabilizing panels driven by torsion springs. Unlike in higher orbits, in VLEO, the attitude motion of the satellite is primarily influenced by the interaction with the atmosphere. Therefore, in order to study the evolution of the satellite attitude during the panels deployment, we considered not only the torques from the torsion springs and the elastic oscillations of the flexible panels, but also the restoring and damping aerodynamic torques, which depend on the satellite shape and thus change substantially during the panels deployment process. Numerical simulations show that if the satellite has intermediate stable equilibrium positions, the combined effect of all above-mentioned torques may result in qualitatively different regimes of oscillations in the angle of attack after the deployment. These oscillations have low-frequency harmonics, related to the aerodynamic torque, and a high-frequency one, induced by the elastic oscillations of the panels after the collisions in the system caused by panels latching. The amplitude of the high-frequency oscillations depends on the post-collision state of the system, which is found analytically using the Appell algorithm. The post-collision state, in its turn, depends on the inevitable inaccuracies of the deployment process, causing the panels to start their rotation at slightly different moments in time. As a result of the study, recommendations for the torsion spring-driven panels deployment in VLEO are formulated. Their focus is to maintain the attitude stability of the satellite and reduce the vibrations of its body in presence of disturbances, such as elastic oscillations of the panels themselves or uneven panels deployment. • CubeSats attitude motion during and after flexible panels deployment is considered. • Collisions in the system due to panels latching are treated using Appell algorithm. • CubeSat with aerodynamic instability may have several attitude oscillation regimes. • Recommendations for the torsion spring-driven panels deployment are formulated. [ABSTRACT FROM AUTHOR]

New industrial dynamics are disrupting the space sector. New stakeholders bringing in capital, technologies, and knowledge from other industries are developing next-generation space infrastructures and services. Both commercial and institutional space projects have to be valuable for a wider set of end-users, asking not only for economic returns but also social and environmental benefits. Space organizations urge understanding and fostering value in the New Space Economy ecosystem. Indeed, end-users are still struggling to enact the expected value of satellite data and solutions for their business. This paper aims to investigate the expected value and the level of adoption (enacted value) of satellite data and satellite-based solutions in the New Space Economy ecosystem from end-users' perspective. We interviewed 21 managers from end-user organizations in different sectors (i.e., Insurance & Finance, Energy & Utility, Transportation & Logistics). Value is deeply discussed in general management literature, and we identify Value Theory as the theoretical lens with the most explanatory power for the phenomenon under examination. From the end-users' perspective, we frame the expected value and the enacted value in taking strategic and tactical decisions regarding their activities, services and products, laying the foundations for further studies of value mechanisms in the New Space ecosystem. Our research set a theoretical and conceptual foundation on value in the New Space ecosystem. It also delineates the blurred boundaries of the New Space ecosystem, the main stakeholders involved, and their perception of value. Insights and implications for strategic and innovation management are also provided. Practitioners may exploit our research results and leverage the end-users-oriented framework to develop next-generation space projects in the New Space ecosystem. We also discuss the limitations and further developments of this research. • We study the value perception of satellite data from the end-users' perspective. • We show how and why end-users adopt satellite data in tactical and strategic choice. • Satellite data are mostly adopted in tactical decisions rather than strategic ones. • End-users face high transaction costs and a lack of literacy about satellite data. • Data providers should focus on providing competencies (not only data) to end-users. [ABSTRACT FROM AUTHOR]

The Martian Moons eXploration (MMX) mission by the Japan Aerospace Exploration Agency, JAXA, will explore the Martian moons Phobos and Deimos. In addition to investigating the moons remotely and returning samples from Phobos, MMX will also deliver a small (about 25 kg) Rover to the surface of Phobos. The payload of the Rover consists of four scientific instruments: RAX, a Raman spectrometer to measure signatures of the mineralogical composition of the surface material, NavCam, a stereo pair of cameras to investigate the terrain and allow for navigation, miniRAD a radiometer to measure the surface brightness temperature of both, regolith and rocks, and two WheelCams which will observe the wheel-surface interface, and thus investigate the properties and dynamics of the regolith. The cameras, will serve both, technological and scientific needs. The rover will be deployed during the rehearsal phase of the landing operations for the main spacecraft and will thus land in proximity to the first sampling area. During its operational lifetime on Phobos (about 100 days), the in-situ measurements performed with the rover payload will provide scientific context for the returned samples, provide ground truth, study the surface heterogeneity and obtain information on the physical properties of undisturbed surface material. The MMX launch is planned for September 2024 and the Rover delivery is scheduled for 2027. The Rover is a contribution by the Centre National d'Etudes Spatiales (CNES) and the German erospace Center (DLR) with additional contributions from INTA (Spain) and JAXA. • In this paper we give an overview of the rover, which will be part of the MMX mission to Phobos. • The MMX rover will explore the surface of Phobos with cameras, a Raman spectrometer and a radiometer. • Scientific objectives and the mission outline are discussed. [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]

Experimental verification of numerical models is a key aspect in all scientific and technological disciplines, yet space systems can be particularly hard to test on ground: in fact, a terrestrial mock-up replicating the orbital dynamics should be in a microgravity environment. One of the most common approaches to (at least partially) realize this condition, consists in free-floating systems, in which the friction between an experimental platform and the working surface is almost completely removed. The vertical axis is still subject to gravity, and only 3° of freedom (two horizontal translations and one rotation about the vertical axis) are available to replicate the orbital behavior. Our research group had been developing since 2012 a free-floating platform named PINOCCHIO (Platform Integrating Navigation and Orbital Control Capabilities Hosting Intelligent Onboard). The need to investigate with higher accuracy the attitude control problem for flexible satellites called for the design, realization and characterization of a new platform with enhanced performance. This paper describes the main components of the system, including the pressure air system, the navigation and communication system, the Guidance, Navigation and Control (GNC) architecture and relevant hardware. The attention is specifically on the characterization of the subsystems' performance. The physical (mass, moment of inertia), control (thrusters' torque and force) and Inertial Measurement Unit (IMU) characteristics are experimentally measured with the associated uncertainty level. The accuracy of the external metrology system in static and dynamic conditions is also characterized, obtaining a 100 μm-level resolution for external measurement of rigid and elastic displacements; this information is fundamental for performing tests of fine pointing control algorithms for flexible and agile satellites. The proposed characterization methodology and associated results could be of interest for university and industry laboratories intended to develop new free-floating platforms. • The characterization procedure of a free-floating platforms is reported. • A stepwise procedure requires to first characterize the supplied thrust accurately. • Inertia properties can be next determined. • Accurate optical navigation is presented and characterized. • Performance of the platform for flexible satellite attitude test is introduced. [ABSTRACT FROM AUTHOR]

The paper describes a spreadsheet-based analysis of the motion of a climber ascending an Earth space elevator tether. The tether is represented by elements of varying lengths, each of mean cross-sectional area based on a taper ratio equation from earlier studies. The tether tension force is calculated in each element based on gravity and centrifugal forces plus the tension force in the element below. A climber is defined by mass, drive power and maximum speed: no consideration is given to design details, the analysis assumes a tractive force simply defined by the traction power. These details derive the mean climber speed on each tether segment and hence the time to ascend each segment. Spreadsheet logic then allocates multiple climber masses on elements at variable travel time spacings, for example 24 h spacings for climbers despatched from the Earth Port once per day. The effective weight of each climber yields an additional tether tension force, giving the total tension in the tether at any altitude. Spreadsheet enhancements include an algorithm for daylight duration at varying altitudes, variable with the time of year: this permits an option for climber spacing to be calculated for solar-powered climbing. The impact of input parameters (climber mass, power and maximum speed, departure intervals, continuous or daylight-only climbing, etc.) yield outputs such as maximum tether tension and climb time. The value of this technique becomes apparent when inputs are adjusted to yield similar tether tensions, representing a scenario with a maximum tether stress limit. It is possible to find, for example, how the maximum climber gross mass varies with maximum speed or drive power. Examples of findings include (1) the benefits of higher power and maximum speed are complex, and highly dependent on the climber power/weight ratio (2) 24-h climbing might allow 20% more payload (for any given tether strength and climber design) compared with daylight-only climbing (3) two smaller climbers launched daily might enable 15% more payload to be raised compared with a single daily launch. Such deeper understanding of climber dynamics highlights the complexity of climber design optimisation: key parameters are the net climber power/mass ratio and the maximum climber speed. • Method described to maximise payload • Smaller climbers with 24-hr power are best [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]

The paper introduces RODiO (Radar for Earth Observation by synthetic aperture DIstributed on a cluster of cubesats equipped with high-technology micro-propellers for new Operative services), an innovative mission concept selected by the Italian Space Agency (ASI) for a phase A study in the framework of Alcor program. The mission relies on a cluster of 4 passive CubeSats flying in formation with ASI PLATiNO-1 (PLT-1) Synthetic Aperture Radar (SAR) mission. The mission is aimed at demonstrating from space the concept of Distributed Synthetic Aperture Radar (DSAR) through the application of multi-platform image synthesis, i.e. the generation of an improved performance image through the coherent combination, so in both amplitude and phase, of a stack of low performance bistatic raw data collected by formation-flying cubesat receivers. RODiO system parameters are discussed and preliminary performance evaluated. Results confirm that multi-static beamforming is able to both improve the Signal-to-Noise Ratio (SNR) and the level of ambiguities in the generated DSAR images, making the relevant values compliant with high-quality SAR imaging standards. • The brand-new CubeSat based concept of RODiO mission is presented exploiting the principle of Distributed SAR (DSAR). • Basic DSAR characteristics like imaging features and applications are described. • Useful tools for preliminary design and performance evaluation of a DSAR mission are introduced. • Application to RODiO mission case with simulated images is presented, demonstrating mission capabilities and features. [ABSTRACT FROM AUTHOR]

In terms of cataloged debris produced, the anti-satellite test carried out by Russia, in November 2021, at an altitude of about 480 km, leading to the destruction of the old satellite Cosmos 1408, was the second worst to date and represented the third worst fragmentation in orbit. It generated more than 1/4 of the cataloged debris produced over 55 years by all such tests and almost twice as many as were produced by all previous Soviet tests. After placing this event in its historical context, this paper analyzes in detail how the evolution of the Cosmos 1408 debris cloud affected the environment below 600 km in the first seven months, focusing on the two operational space stations and the Starlink large constellation of satellites. During the first six months following the test, the Cosmos 1408 cloud of fragments nearly doubled the average flux of cataloged objects on the International Space Station and increased by about 3/4 that on China's Tiangong. In the same period, the Starlink large constellation saw an average increase in the flux of cataloged objects of about 20%. Some orbital planes, the "counter-rotating" ones with respect to the Cosmos 1408 debris cloud, were more affected than others, and the affected planes gradually changed over time, due to the differential precession of cloud and constellation nodes. However, being the Starlink constellation 70 km higher up, the flux of Cosmos 1408 cataloged debris steadily decreased over the period analyzed, due to the cloud orbital decay, reducing to just over a quarter of its extrapolated initial value after seven months. • A review of ASAT tests as a source of orbital debris was presented. • The impact of the Cosmos 1408 destruction was put into context. • The cataloged debris flux increase on space stations was computed and discussed. • The debris flux increase and evolution on Starlink orbit planes was analyzed. • A moratorium on this type of ASAT tests was recommended, including low LEO. [ABSTRACT FROM AUTHOR]

MIRS is part of the French contribution to the Martian Moons eXploration (MMX) mission that will be launched in 2024 by the Japan Aerospace Exploration Agency (JAXA). It is a near-infrared imaging spectrometer developed by the French Laboratory of Space Studies and Astrophysics Instrumentation (LESIA), of the Observatoire de Paris – PSL , with close cooperation and financial support from CNES. One of the major mission goals is to understand the origin of Phobos and Deimos, providing important insights into planetary formation and the transfer of matter in the region connecting the inner and outer solar system. During the four years of the mission, the MIRS control centre will provide precise sequencing of MIRS activities and guidance of its line of sight, with two main drivers: optimize scientific return and comply with the operability constraints of the instrument, the satellite and the system. In this perspective, observation strategies are designed, modelled and tested iteratively throughout all preparation phases of the mission. Discussions on concrete simulation results and illustrations improve mutual understanding, raise unexpressed needs and constraints and promote feedback on the efficiency of the proposed strategies. They also favour a global optimization of the entire system with the refinement of the trajectories and of the scheduling considering the seasons of the Martian system. This paper describes the whole process of observation strategies development in cooperation with the MMX and MIRS system and scientific teams, as well as the current status and results. • Observation strategies elaboration process from design to operations is presented. • It involves a continuous and close cooperation with the Scientific and System teams. • It defines precisely How and When to perform observations. • Optimization algorithms are implemented to solve scheduling combinatorial problems. • Continuous validation and performances assessment allow to refine the design. [ABSTRACT FROM AUTHOR]

This paper presents an experimental activity aimed at assessing performance of techniques for inertia and attitude parameters estimation of an uncooperative but known space target. The adopted experimental set-up includes a scaled-down 3D printed satellite mock-up, a spherical air bearing and a low-cost solid-state LIDAR. The experimental facility also comprises a motion capture system to obtain a benchmark of the pose (position and attitude) parameters and an ad-hoc designed passive balancing system to keep the centre of mass as close as possible to the centre of rotation. The LIDAR-based 3D point clouds, collected while the target rotates on the spherical air-bearing to reproduce the rotational dynamics of an almost-free rigid body, are processed to obtain pose and angular velocity estimate. Those data are used to determine the inertia properties of the target by solving a linear system based on the integral form of Euler equations. Performance and robustness of the algorithm for attitude and inertia properties estimation are assessed considering both the inertia benchmark provided by a high-fidelity CAD model of the target and the pose solution obtained from the motion tracking system and its extrinsic calibration with respect to the LIDAR. • The estimation of attitude and inertia parameters of an uncooperative space target is addressed. • A LIDAR-based algorithmic architecture is experimentally validated. • The design and calibration of the experimental set-up is described. • Errors below 10° and 20% are obtained for attitude and inertia parameter, respectively. [ABSTRACT FROM AUTHOR]

Missions to small celestial bodies rely heavily on optical feature tracking for characterization of and relative navigation around the target body. While deep learning has led to great advancements in feature detection and description, training and validating data-driven models for space applications is challenging due to the limited availability of large-scale, annotated datasets. This paper introduces AstroVision, a large-scale dataset comprised of 115,970 densely annotated, real images of 16 different small bodies captured during past and ongoing missions. We leverage AstroVision to develop a set of standardized benchmarks and conduct an exhaustive evaluation of both handcrafted and data-driven feature detection and description methods. Next, we employ AstroVision for end-to-end training of a state-of-the-art, deep feature detection and description network and demonstrate improved performance on multiple benchmarks. The full benchmarking pipeline and the dataset will be made publicly available to facilitate the advancement of computer vision algorithms for space applications. • Missions to small bodies rely heavily on humans for optical feature tracking. • Deep Learning has led to great advancements in feature detection and description. • We present AstroVision, a first-of-a-kind dataset for vision-based tasks near small bodies. • We conduct an extensive evaluation of feature detection and description methods. • We demonstrate increased performance by training a deep feature network on our data. [ABSTRACT FROM AUTHOR]

To continue space operations with the increasing space debris, accurate characterization of fragment fly-out properties from hypervelocity impacts is essential. However, with limited realistic experimentation and the need for data, available static arena test data, collected utilizing a novel stereoscopic imaging technique, is the primary dataset for this paper. This research leverages machine learning methodologies to predict fragmentation characteristics using combined data from this imaging technique and simulations, produced considering dynamic impact conditions. Gaussian mixture models (GMMs), fit via expectation maximization (EM), are used to model fragment track intersections on a defined surface of intersection. After modeling the fragment distributions, k-nearest neighbor (K-NN) regressors are used to predict the desired characteristics. Using Monte Carlo simulations, the K-NN regression is shown to predict the distributions for both the total number of fragments intersecting a given surface, as well as the expected total fragment velocity and mass associated with that surface. This information can then be used to estimate the kinetic energy of the particle to classify the particle and avoid debris collisions. • Machine learning combines high-fidelity simulations with experimental data. • High speed collisions or explosions can be characterized. • Fragment energy and distribution can be predicted. [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]

On-orbit servicing for the failed spacecraft involves two necessary stages: on-orbit observing and approaching. For these two stages, a novel approaching and pointing tracking control scheme under motion constraints and input saturation is proposed in this paper. For the security of approaching, a relative position guidance law for specific surface hovering is proposed. Then, a model predictive controller based on quadratic programming algorithm is designed. For the continuous observation task, the optimal maneuver angular velocity under the motion constraint is planned through model predictive control. In order to reduce the dimensionality of the optimization problem to reduce the computational effort, only the attitude kinematics equation is used as the process model. Thereafter, the adaptive anti-saturation attitude controller is designed to track the optimal maneuver angular velocity. The anti-saturation auxiliary system and the adaptive update law are used to cope with the control torque saturation and the integrated uncertainties, respectively. Finally, the stability of the closed-loop system is demonstrated by Lyapunov method. Simulation results demonstrate that the proposed control scheme accomplishes the approaching and pointing tracking task, and finally service spacecraft maintains hovering above the tumbling target and observing its specific surface. • The problem of approaching and pointing tracking for tumbling target is studied. • A relative pose guidance law for specific surface hovering is proposed. • A novel control scheme under motion constraints and input saturation is proposed. •.The control scheme accomplishes the tracking task with optimal angular velocity. [ABSTRACT FROM AUTHOR]

The purpose of this paper is to study the birth of space age and the distinguished achievements in the space-age painting aesthetics inspired by space exploration. It has been 66 years ever since the first successfully launch of the artificial satellite Sputnik 1 on October 4, 1957, generally recognized as the first year of space age. Exemplified by the best and most well-known "NASA Art Program", many major space agencies have sponsored the art program. In Japan, the Japanese Aerospace Exploration Agency (JAXA) supports the art program through its Public Affairs Department. In the European Space Agency (ESA), one of its well-known residence artists is Aoife van Linden Tol. For 66 years of the space age, there are many outstanding painters with many distinguished achievements and heritage of aesthetic paintings created. In particular, Robert Rauschenberg's work " Buffalo II " painted in 1964 has become a classic masterpiece that realized the price of $ 88 million at Christie's auction in 2019. • It has been 65 years since the launch of Sputnik 1 in 1957 generally recognized as the first year of space age. • After the well-known "NASA Art Program", many major space agencies have sponsored the art program. • In Japan, the Japanese Aerospace Exploration Agency (JAXA) supports the art program through its Public Affairs Department. • In the European Space Agency (ESA), one of its well-known residence artists is Aoife van Linden Tol. • Robert Rauschenberg's work " Buffalo II " painted in 1964 priced $88 million at Christie's auction in 2019. [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]

Under the development trend of large-size antennas, the demand for modular expansion of synthetic aperture radar antennas is increasing, and the connection mechanism is an important part of the modular expansion. This paper designed and studied the connection mechanism between the basic units of the planar deployable antenna mechanism. Firstly, for the basic deployable units of a new type of planar deployable antenna spatial support mechanism proposed previously by the author, the connection mechanism between the basic units was synthesized, and a series of feasible planar deployable mechanism configurations were obtained. Meanwhile, a spatial support mechanism for planar deployable antennas with scissor was optimally constructed by analysis. Through finite element comparative analysis, it was found that the mechanism improves the stiffness and reduces the maximum deformation. Then, scale optimization was performed by using the control variable method. Finally, the prototype was developed, and experimental research was conducted to verify the feasibility of the mechanism. The spatial support mechanism of the planar deployable antenna with scissor proposed in this paper improves the reliability, stiffness, modal, and other related properties of the unfolding process. Therefore, it has broad application prospects in the field of aerospace deployable antennas. • Propose a method to design the connection mechanism between the basic deployable units. • Design a spatial support mechanism for planar deployable antenna without P joints. • Compare the stiffness of the new mechanism and the previous one. • Develop a prototype of the proposed support mechanism and validate the motion principle. [ABSTRACT FROM AUTHOR]

The landing of a probe on the surface of an asteroid can produce a stable operating platform and obtain vital data describing the mechanical response of the asteroid material. The fine-grained regolith area on the asteroid surface is essential in cushioning the probe from impact energy and in avoiding the risk raised by the large protruding rocks during the landing of the probe. Models incorporating macro forces explaining the interaction between the landing terminal of the probe and the regolith on the asteroid surface conclusively reflect the mechanical process and characteristics of the interaction, making them particularly suitable to analyze and control the landing process of the probe. In this paper, we study the macro normal force of the interaction between the footpad of a legged probe and the regolith on the asteroid surface with particular attention paid to the mathematical expression of the force model and the contribution of the most essential soil parameters. The ultimate bearing theory of soil mechanics is extended to discuss the mathematical model of the macro normal force on the asteroid regolith in a micro-gravity environment undergoing different damage modes. Subsequently, a series of discrete element method simulations are implemented to verify the normal force model and to analyze its characteristics under the influence of landing impact. On the basis, the effect modes of several most essential soil parameters of soil porosity, particle size and cohesive strength are revealed through a robust regression analysis of the simulation data. Next, the correlation coefficients of the effects of soil parameters on the macro normal force are researched. The model of macro normal force established in this paper can be used to facilitate an effective performance analysis in the development of the landing-cushioning mechanism of the probe, as well as to conclude a reference for the most essential soil properties parameters of the asteroid regolith via the inverse solution of the interaction data recorded in actual landing missions. • Regolith on the asteroid surface provides benefit for the probe in landing-cushioning. • Model of the macro normal force explaining footpad-regolith interaction is researched. • Slopes of the force variation hold a exponential relationship with the soil porosity. • Particle radius affects the slopes of force variation via the particle surface area. • Cohesive strength has a more complex effect mode on the slopes of force variation. [ABSTRACT FROM AUTHOR]

With commercial participants now undertaking an ever-expanding role within the development and supply-chains of China's space industry post Document No. 60, the resulting legal challenges relating to this emerging industry status-quo has garnered much academic discussion. Specifically, while China's space industry has grown at a rapid pace within the past few decades, China's legislative framework has remained largely silent over matters relating to space. Instead, China continues to remain as the only major space-faring nation without a national space law and regime. This substantive legislative gap over China's space related activities has negatively impacted the industry's current deregulation process. While it has been the academic community's primary solution to address the industry's current substantive legislative gap through coordinated rulemaking activities, this paper aims to advance the current academic discussion by taking an alternative approach and arguing that the fundamental legislative challenge and solution to China's emerging commercial space industry, is not substantive, but rather procedural by nature. Importantly, by reconceptualizing the industry's substantive gap from one that is determined by the number of missing legal instruments, to one determined by the rate of legislative activity i.e., the pace of law-making activities vis-à-vis commercial expansion, this paper argues that only through the implementation of legal procedural rulemaking mechanisms will China's space law regime be able to sufficiently address the current substantive legislative gap at a pace matching commercial expansion. • The deregulation of China's space industry. • Regulatory structure and legal challenges of China's space industry. • A procedural argument to resolve China's substantive space law gap. [ABSTRACT FROM AUTHOR]

On November 26th, 2018, the InSight spacecraft successfully landed on Mars after a 6-month journey. After a long deployment and commissioning phase, the SEIS (Seismic Experiment for Interior Structure) instrument was ready to monitor seismic events on the Elysium Planitia plain on the surface of Mars, coupled with the APSS (Auxiliary Payload Sensor Suite) weather station equipped with a magnetometer, wind sensors, and a pressure sensor. The InSight mission goal is to characterize the deep interior structure of Mars, including the thickness and structure of the crust, the composition and structure of the mantle, and the size of the core. Its nominal duration of two years (2019–2020) has yielded unprecedented results with the detection of the first Martian seismic events ever recorded, and the in-depth characterization of its atmosphere with the best weather station ever deployed on Mars. InSight has collected an outstanding amount of high-quality measurements that the scientific community will spend many years analyzing. The extended mission has started and covers the years 2021 and 2022. This paper will describe the operations of the SEIS experiment on Mars since landing, as well as the challenges of operating this instrument. Energy becomes increasingly limited for payloads on Mars due to a significant amount of dust accumulated on the solar panels and the many dust storms in the Martian atmosphere. A new activity was decided for the extended mission in 2021 which consisted in burying the seismometer cable (or tether) with Martian regolith collected locally using the robotic arm, in order to reduce the seismic noise from that subsystem. Preparation activities, testing, results, associated challenges and lessons learned will be presented. Moreover, the paper will address the challenges faced in carrying out operations with COVID-related constraints, as finding oneself operating a seismometer on Mars from home can be challenging. Finally, management of periods of solar conjunctions, during which communication between Earth and Mars is unavailable, will be addressed. • Deployment operations of a seismometer on Mars. • Detection of the first Mars seismic events. • Mars surface operations when energy is restricted. • Space operations in the context of COVID-19. • Unexpected tether burial and solar array cleaning operations on Mars. [ABSTRACT FROM AUTHOR]

This paper investigates the attitude controller design for the reusable launch vehicle (RLV) in its re-entry phase under the influence of model uncertainties and external disturbances. The navigation of RLV along a predefined safe trajectory is a challenging task since the RLV system dynamics has a complex nonlinear behaviour with uncertain environmental conditions. Therefore, this paper proposes a time-delay estimation technique with fractional order integral sliding mode control abbreviated as FOISMC-TDC, for efficient tracking of a pre-designed reference trajectory that ensures safe landing of the RLV. The time delay estimation method in the proposed research employs the input–output information of the immediate past instant to estimate any uncertainties and un-modelled dynamics. The estimated value is then fed to the composite control for disturbance compensation without primarily relying on the switching control of FOISMC approach. This in turn, eliminates the problem of high-frequency chattering in the control input. In addition, by using FOISMC philosophy finite time convergence of system states is achieved with desired error response. Furthermore, Lyapunov and homogeneity theories have been utilized to establish practical finite-time stability of the overall closed loop system. Numerical analysis with state-of-the-art techniques have also been presented to demonstrate effectiveness of the proposed methodology. • Fractional order ISMC with time-delay estimation technique for RLV system. • Lyapunov and homogeneity theories ensure practical finite-time stability. • Achieve chattering suppression & exemption from knowing the upper limit of disturbance. • Simulation demonstrates efficient tracking as compared to state-of-the-art methods. [ABSTRACT FROM AUTHOR]

On-orbit Servicing in the form of space-manufacturing, refuelling, assembly and robotic manipulation are deemed to become a fundamental part of the space market in the short-to-middle term. One of the key aspects to reach extensive use of OOS operations is having a standardized, reliable, and cost-effective interface that embeds as much functionality as possible to simplify operations. In this context, SENER Aeroespacial has developed SIROM (Standard Interface for Robotic Manipulation), which provides mechanical coupling, electric interface, thermal and data transfer in a single compact solution. In this paper, the current design for SIROM is presented, with a focus on the electronics subsystem and the integration of the electromechanical design. A functional breakdown of the electronic boards is presented after a review of the current state of the art. For the validation, a TRL6 Test Campaign that was successfully passed in early 2021, including vibration, temperature, vacuum and life tests. Its results are presented here, along with the details of the functional validation on the context of the EROSS project in May/June 2021. To conclude the paper, on-going projects involving the SIROM interface are reviewed, finishing with future lines of development on the SIROM roadmap. • On Orbit Servicing is becoming a fundamental part of the space market. • SENER designed the SIROM interface for this purpose, developing a fully integrated model in 2021. • SIROM has passed a qualification campaign for Technology Readiness Level 6. • SIROM has been validated in the H2020 project EROSS in an On Orbit Servicing scenario. • SIROM is developing to reach TRL 8 as part of European projects (MIRROR, PERIOD). [ABSTRACT FROM AUTHOR]

This paper traces the early ideas of lunar architecture developed by the US and Soviet space programs. Starting with a historical overview of those achievements, highlighting specific engineering, architectural, and design elements; illustrated with images and drawings. Some of them will be described in more detail, showing the design and decision-making process from the first sketch to its implementation. One example of the great engineering and management capabilities is the development and inclusion of the deployable lunar roving vehicle (LRV). Other examples include the project DLB Zvezda by Barmin's Design Bureau of General Engineering that was under development between 1962 and 1974, and a US project Horizon – a study to set up a Lunar Military outpost conducted in 1959. The paper includes personal statements from astronauts and cosmonauts, as well as people involved along design and implementation processes. Based on a historical overview of achievements, this paper highlights particular architectural and design unique approaches that defined design strategies for creating space systems, spacecraft and habitats for decades to follow. • Lunar architecture developed by the US and Soviet space programs. • Specific engineering, architectural and design elements. • Lessons learned from earlier lunar base designs. • Space habitat designs. [ABSTRACT FROM AUTHOR]

Martian Moons eXploration (MMX) is a mission under development in JAXA in cooperation with NASA, CNES, ESA, DLR to be launched in 2024. This paper introduces the result of its preliminary design and the latest status of the MMX program, putting more weight on the novel part of the mission. The goal of MMX is to reveal the origin of the Martian moons and then to make progress in our understanding of planetary system formation and of primordial material transport around the border between the inner- and the outer part of the early solar system. Additionally, the mission is to survey two Martian moons and return samples from Phobos. Add to those MMX's contribution to the planetary science field, on the growing discussion on the International Space Exploration activities, MMX's contribution to future human Mars exploration is also considered as an essential aspect of the program. Following the system definition study results presented in the previous conference, the following items will be reported in this paper. First, as a result of the comprehensive completion of the Phase-B activities, the preliminary design is completed in coordination with the design of the spacecraft system, mission instruments, and operation plans. This paper describes the proximity and surface operations around Phobos in detail. Second, Phase-C activities have started, incorporating engineering models manufacturing and tests. Those of critical technologies for surface exploration are described in detail. Moreover third, the programmatic aspects, including international cooperation frameworks and the program schedule, are presented. • Martian Moons eXploration (MMX) is a JAXA's mission to be launched in 2024. • MMX achieves the world's first sample return from a Martian moon, Phobos, in 2029. • MMX investigates the origin of Martian moons and the planetary formation process. • Sample of Phobos contains the ejecta derived from Mars by meteorite impacts. • MMX acquires data that contributes to future human Mars exploration. [ABSTRACT FROM AUTHOR]

For a few years now, CNES has been developing ASTERIA, an on-board autonomy concept combining autonomous station-keeping and collision risks management for Low Earth Orbit (LEO) satellites. ASTERIA, acronym for Autonomous Station-keeping Technology with Embedded collision RIsk Avoidance system, enables both in-track and cross-track control for different LEO missions. The on-board collision risk management process is fully integrated into the autonomous station-keeping in order to maintain the satellite orbit as best as possible and to minimize mission unavailability resulting from the avoidance maneuvers. Collision risk management requires the best possible knowledge of the future trajectory of the primary satellite, up-to-date information on secondary objects, as well as a calculation process with a large amount of data and the propagation of orbital states and covariance. That is why, so far, collision risk management has always been a ground segment activity. But this strongly limits on-board autonomy by imposing a timing and a knowledge of the station-keeping maneuvers not very suitable with a reactive Autonomous Orbit Control (AOC) system. On the contrary, management by the spacecraft itself offers interesting prospects for responsiveness and increased autonomy, but requires to carry out on-board risk calculation, to identify relevant risk assessments and to implement efficient avoidance solutions. The paper aims to show the completeness of the ASTERIA concept. First, the principles of on-board collision risk management are described. Then, the paper presents the study of the on-board computational load, the accuracy of risk estimation and the performance of risk avoidance strategies. The ability to operate such a system is then demonstrated through the study of the related operational management process. The concept has been tested and validated by in flight experiments on the ESOC OPS-SAT 3-Units CubeSat. Some preliminary results of this experiment are presented. Finally, the issues related to an autonomous system are discussed, particularly those concerning space traffic management. • Autonomous orbit control method including station keeping strategy and collision risk mitigation. • Definition of a ground and on-board filtering process for the space objects catalog. • Conclusive validation through in-flight experiments on the OPS-SAT satellite. • Relevant estimate of the level of collision risk calculated on-board. • On-board determination of avoidance solutions compatible with low-thrusts. [ABSTRACT FROM AUTHOR]

The advances in the optical space surveillance technology and the possibly more deployments of space-based optical sensors enable more and more space objects detectable and make the catalog expansion more realistic. The cataloging of an optically detected new object requires finding three or more tracklets over a few days from the object. In this process, the tracklet-to-tracklet association is a critical step, but associating short-arc optical tracklets remains a difficult task. This paper develops a novel orbit determination approach to the optical tracklet association. The main idea is to determine accurate orbit elements from two tracklets of the same object by exclusively using the Gooding angles-only initial orbit determination (IOD) method. Before applying the Gooding IOD method to two tracklets apart by a few days, the dominant secular effects on the angular observations have to be considered. The significant short-periodic effects may also need consideration if tracklet-to-tracklet IOD is to achieve high accuracy. In this paper, the secular effects from the J 2 and lunisolar gravitational perturbations, as well as the short-periodic effects due to J 2 and J 2 , 2 , are accounted for in the tracklet-to-tracklet IOD. Extensive tests using tracklets from low Earth orbit (LEO), high elliptical orbit (HEO), medium Earth orbit (MEO), geosynchronous Earth orbit (GEO), and Molniya orbit are performed to assess the performance of the developed method. It is shown that the true positive (TP) rate for associating two tracklets from the same object within 48 h is 98.3% for LEO objects and higher than 89.8% for HEO and Molniya objects. The TP rate is remarkably high at around 99.0% for MEO and GEO objects. The analytic approach is also computationally very efficient. The results demonstrate the strong applicability of the developed method to associate optical tracklets of uncatalogued objects of all orbit types. • A purely analytical approach to associating too short optical tracklets is proposed. • An accurate orbit is computed with two tracklets after correcting perturbations. • The performance of the tracklet-to-tracklet association is verified by real data. • The True Positive rate is above 98% in LEO and GEO tracklet-to-tracklet associations. • The developed association approach is computationally very efficient. [ABSTRACT FROM AUTHOR]

The Earth's gravity field and upper atmosphere are very important scientific information and their mechanical models can be obtained by inversion of the satellite precision orbit. It would be ideal to develop a constellation of satellites distributed with different orbits to measure atmospheric drag and gravity. Based on this idea, Tsinghua University developed a microsatellite Q-Sat for measuring the gravity field and the density of the upper atmosphere. Low-cost, high-precision measurements can greatly improve the cost effectiveness of scientific missions. In this paper, a spatial modification method for GNSS receivers based on commercial off-the-shelf was proposed. The base module of the receiver is a ground-based general-purpose device, and this modification method allows for the space-batch modification of industrial-grade devices. The modified receiver was successfully applied to the Q-Sat and a large amount of continuous on-orbit data were obtained. Data analysis results show that the receiver can simultaneously receive GPS-L1\L2 and BD1\BD2 multifrequency navigation satellite signals in the highly dynamic spatial environment. To reduce the precision orbit determination error, the GNSS observations were corrected by antenna center modelling, and the long-term orbit determination accuracy is stabilized at the centimeter level. The commercial off-the-shelf based GNSS receiver modification method and precision orbit determination strategy can provide a reference for low-cost and high-precision microsatellites. • There are four highlights in this paper. • Low-cost, high-precision GNSS receiver modification and testing method based on COTS. • Supporting GPS and BDS precision orbit determination. • Antenna center model modification significantly improving the POD accuracy. • Centimeter-level orbiting accuracy supporting microsatellite science missions. [ABSTRACT FROM AUTHOR]

Space Gravitational Wave (GW) detection task puts forward super-low thrust noise for micro-newton propulsion system. Study shows thrust noise can be reduced by thrust closed loop feedback control structure. The effect is determined by thrust estimation accuracy. Traditionally, thrust estimation is realized by electric parameters and thrust estimation equation. However, this way is faced with high thrust estimation error in the low frequency band. Besides, the displacement sensor information of the drag-free control system can also be used for thrust estimation via disturbance observer. However, this way is only effective for low-frequency thrust due to the measurement noise of displacement sensor. A data fusion method is proposed in this paper to enhance thrust estimation accuracy. In this method, the high frequency thrust noise is estimated by electric parameters and thrust estimation equation. Meantime, the low frequency thrust noise is estimated by displacement sensor information and disturbance observer. In this paper, the cusped field thruster is used as the actuator of drag-free control system to simulate. Simulation results show thrust estimation by data fusion can effectively reduce thrust noise in the middle frequency band and enhance stability of satellites compared with that by electric parameters. • high-precision thrust estimation contributes to low thrust noise by feedback control. • New data fusion thrust estimation way is effective for space gravitational detection. • low-frequency thrust noise is got from displacement sensor by disturbance observer. • high-frequency thrust noise is estimated from electric parameters and equation. [ABSTRACT FROM AUTHOR]