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The launcher recovery method known as 'In-Air Capturing' is an innovative approach where a winged rocket stage is captured mid-air by an aircraft and subsequently towed back to the launch site. To achieve this, a capturing device attached to a tether is released from the towing aircraft. This device autonomously connects the reusable launch vehicle to the aircraft, while the two vehicles are in proximity. The tether's flexible dynamics have a strong influence on the maneuverability of the capturing device. Thus, it is critical to model the dynamics of the tether to get a realistic understanding and to evaluate feasibility of the concept. However, precise modelling and control of highly flexible systems with large deformations is both challenging and computationally intensive. For engineering applications which involve closed-loop simulations, finding a trade-off between accuracy and computational effort is crucial. In this paper, the tether is modelled as a discretized chain of rigid bodies connected by rotational springs at the joints. The tether along with the capture system is integrated in a simulation model and open loop tests are performed to analyse the system characteristics. The tether properties are studied to find a suited configuration for 'In-Air Capturing'. • Multibody modelling approach for long cable like bodies with large deformation. • Capture system dynamics for the unique 'In-Air Capturing' launcher recovery concept. • Dynamic modelling for the complete capturing system including towing aircraft, capturing device and tether. • Assessment of tether properties suited for 'In-Air Capturing' with sensitivity studies. [ABSTRACT FROM AUTHOR]

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

Balloon-based exploratory platforms, or aerobots, are a budding planetary exploration technology for Venus to collect scientific data in-situ while keeping a safe distance from the highly inhospitable surface of Earth's "evil twin". Such a mission is expected to deploy and inflate the aerobot during the system's initial descent through the Venusian atmosphere, rather than launched from the surface. The analysis of the entry, descent and float (EDF) sequence of the aerobot is accordingly a delicate and fundamental part of the aerobot design, and generally where the envelope undergoes the largest loads. The inflation phase in particular plays an essential role, especially for variable-altitude aerobots that can have multiple-envelopes with different inflation requirements. This paper addresses a collection of preliminary considerations about the inflation phases of a Venus aerobot – starting from the end of deployment to completion of reaching an equilibrium altitude – where the nominal mission flight profile would then begin. We include two different inflation sequences, highlighting issues and threats, utilizing a simulation of the inflation dynamics of the aerobot conducted in MATLAB. The paper also presents results from a set of rapid inflation experiments, conducted on a subscale prototype of the aerobot, aimed at an initial assessment of feasibility and response of the materials involved. • Inflation sequence design for a variable-altitude aerobot at cloud-level on Venus. • Sequence modification to reduce peak loads on the aerobot. • Constraint-driven analysis to evaluate the impact of the model's initial parameters. • Laboratory demonstration of the inflation utilizing a subscale aerobot prototype. [ABSTRACT FROM AUTHOR]

Virtual reality (VR) and augmented reality (AR) are rapidly developing technologies that may have important healthcare-related applications, including the possible evaluation of the health and performance of astronauts. Advances in the fields of VR/AR and space exploration have independently progressed; however, over the last decade synergies and overlapping applications have emerged between spaceflight and extended reality. This paper is intended to be a resource to update the medical community on the current and future integration of VR and AR into space exploration, particularly as this integration relates to astronaut health. We highlight how these technologies are used in terrestrial healthcare with specific attention to visual function. Furthermore, we elaborate on emerging VR applications being developed for future spaceflight. The paper discusses the current state of the art, such as the Apple Vision Pro headset, and outlines key challenges like mass and volume constraints, power limitations, and the heightened cognitive and physical demands of space travel. It also addresses the potential of VR/AR technologies to mitigate these issues through efficiency, stress reduction, and exercise motivation. Finally, we emphasize the benefits that these technologies provide to overcome possible barriers to long duration spaceflight and manned exploration missions, including spaceflight associated neuro-ocular syndrome (SANS). • Virtual reality (VR) and augmented reality (AR) are rapidly developing technologies that may have important healthcare-related applications, including the possible evaluation of the health and performance of astronauts. • This paper is intended to be a resource to update the medical community on the current and future integration of VR and AR into space exploration, particularly as this integration relates to astronaut health. • We highlight how these technologies are used in terrestrial healthcare with specific attention to visual function. Furthermore, we elaborate on emerging VR applications being developed for future spaceflight. [ABSTRACT FROM AUTHOR]

Recent startups and university space activities extend to cis -lunar and near-Earth interplanetary field. One of the biggest hurdles in those missions is in the orbit determination based on the range measurements, while the radio transmission and reception capability becomes within those small organizations' facilities. Also, about the navigation in cis -lunar space, conventional GNSS (Global Navigation Satellite System) real time positioning is hardly available and some alternative methods have been sought. The most straight forward approach is to be supported by the space agencies' facilities for the range measurement with the transponders onboard the probes. However, it depends on the facilities availability and also on the relatively expensive, old and analogue transponders. This paper presents a new method deriving from the GNSS based devices. It does not require the transponders but inexpensive radio circuits easily built. And both the probes and the ground stations have only to be equipped with the common and simple inexpensive devices. The method enables the small organizations to perform autonomous operation of the probes in a distance beyond the moon's orbit. The paper first presents the latest test data obtained during the development using the Software Defined Radio (SDR) boards. • Low-Cost Ranging Measurement for Deep Space beyond Cis-Lunar Region. • Real Test Results using Software Defined Radio Boards, Hardware. • Semi-Real-Time and Simultaneous Range Measurement and Clock Synchronization. • Absolute Clock Relay Capability via A Pair of Spacecraft. [ABSTRACT FROM AUTHOR]

Space-based data provide essential information on the past, current, and future state of climate change. Satellite data widely contribute to scientific research conducted by recognized institutions such as the IPCC, which underpin most climate policies in Europe. In addition, half of the Essential Climate Variables (ECVs), which are necessary to monitor climate change, directly come from satellite data. However, the role of space is not always accurately reflected in Europe's climate policies, with only a few direct references to ESA's or EU's space programmes and data. Against this background, this paper aims to provide an in-depth assessment on the actual role played by space-based data in the definition and implementation of European climate policies. Based on the outcomes of a year-long research conducted in collaboration with the European Space Agency, the paper will first analysed the use and references to space-based data in over 500 climate policy documents of the 22 ESA Member States. Subsequently, it will address some potential blocking points to the use of space in the climate policymaking process in the analysed European countries. By doing so, the paper will also look into aspects related to science communication, the science-to-policy divide as well as the climate governance in ESA Member States. Finally, this paper will provide recommendations to increase the use of space-based data in the climate policymaking process. • Climate policies in ESA Member States show that references to Essential Climate Variables are abundant. • The origin of the data or the systems used to generate variables are not referenced despite half of variables being only measurable from space. • All ESA Member States reference space in their climate policies. • 17 ESA Member States have reference ESA in their climate policies. • A blocking point to the integration of space-based data in climate policies pertain to the science to policy divide. [ABSTRACT FROM AUTHOR]

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

How does a country embark on large-scale R&D programs? Among R&D programs, this is an intriguing question for large scale Science and Technology (S&T) projects such as space programs, because of the necessity for large resource commitments, and the difficulty of being accountable to the taxpayers when the outcome is not always known. It can be assumed that engineers, scientists, bureaucrats, and politicians play respective roles. This paper aims to shed light on the specific role of politics, through the examples of policies of and decisions by NAKASONE Yasuhiro. Nakasone was the minister for science and technology (1959–1960, 1972), and later became the prime minister (1982–1987). He played a crucial role in establishing the Science and Technology Agency and the National Space Activities Council in the 1950s, which would be the foundational institutions responsible for space policy for the next half century. As prime minister, he decided to participate in the Space Station program. Within his vision for Japan for the 21st century, he emphasized the need for radical change in economic policy, education, defense and S&T investment. He was one of the few politicians that had vision to promote S&T and recognized the strategic value to the nation. He was able to explain to the nation why Japan should promote the development of a launch vehicle, when the engineers faced development difficulties in doing so. Thus, he was influential both in the policy layer and in promoting specific programs to support his national vision. Politics intersects with S&T projects because of the uncertainty and tradeoffs associated with these efforts. Japan did not address space programs for national security and commercial space for a long time. It is a politician's role to create institutions, to allocate resources and, more importantly, to explain to the taxpayers why policy measures or programs are necessary. R&D investment is a societal activity: it reflects the social, political, and economic situation, as well as being important for diplomacy and international relations. Bearing in mind that it would be misleading to attribute the initiation of largescale S&T projects to the leadership of a single individual, this paper will attempt to show the driving forces behind, and fundamental objective of Japan's space policy, in addition to the historical shift in that policy by examining the influence of this notable politician. • Nakasone played the key role in establishing institutions promoting space development in post-war Japan. • Nakasone made the best use of close relationships with the U.S. in promoting Japan's space activities. • Nakasone made decisions on space policy process, resource allocation with vision, thus integrating the stakeholders. [ABSTRACT FROM AUTHOR]

The Korea Pathfinder Lunar Orbiter (KPLO), officially named as Danuri, is the Republic of Korea's first spacecraft to orbit the Moon. The goal of KPLO is to secure core deep space technology for future space exploration in Korea and contribute to lunar science by gathering data from the lunar orbit for one year. This paper presents the results of KPLO's Lunar Orbit Acquisition (LOA) phase in detail, with a particular focus on the operational results of Flight Dynamics (FD). During the LOA phase, the KPLO FD team experienced many expected and unexpected issues. These issues were mitigated by making real-time updates to the LOA phase Design Reference Mission (DRM). The paper also presents the results of the Orbit Determination (OD) and performance of each Lunar Orbit Insertion (LOI) maneuver conducted during the actual flight operation. The importance of considering operational constraints when designing the DRM is emphasized by presenting lessons learned based on actual flight experiences. Unlike the original KPLO LOA phase DRM, KPLO successfully achieved the final mission orbit using only three LOI burns, instead of the planned five LOI burns and an Orbit Trim Maneuver (OTM). Despite the extensive revisions made to the DRM, all mission requirements were still fulfilled during the final lunar orbit insertion, taking into account the significant operational constraints. • Real flight operation results of the Korea Pathfinder Lunar Orbiter's lunar orbit acquisition phase are presented. • Orbit determination and lunar orbit insertion maneuver performance are analyzed based on actual flight data. • Unexpected issues during the actual flight operation were mitigated through real-time updates on the design reference mission. • Considering real-world operational constraints while designing the design reference mission is very critical. [ABSTRACT FROM AUTHOR]

In order to help control the spread of space debris and junk, deorbit sail devices have been used for deorbiting techniques; however, this approach is still considered risky due to potential failure or malfunction as it need to be exposed to the harsh space environment for a prolonged period of time, particularly in low-Earth orbit. China has made new strides forward in managing space junk, as it has successfully unfolded a 25-square-meter deorbit sail developed by the Shanghai Academy of Spacecraft Technology in-orbit for the payload capsule of a recently launched rocket. Accordingly in this paper, a full life-cycle modelling based on Monte Carlo method is developed to reveal the damage behavior of the 25-square-meter deorbit sail surface under the combined effects of atomic oxygen erosion and deorbit sail defects. The simulation results show that as the atomic oxygen fluence increases and the material defect width widens, the interaction effect of the two factors becomes more and more pronounced, which decisively influences the erosion rate of the sail surface. Morphological characterization of aluminized film PET material, activated silanisation modified material and plasma-polymerization coating material before and after atomic oxygen erosion combined defects is carried out and the atomic oxygen erosion of the deorbit sail surfaces is quantitatively assessed. The assessment results show that compared to the PET material and activation silanisation surface modification material, the plasma polymeric coated material showed much better flexibility, folding resistance and irradiation resistance to prevent geological fold and to shield from the atomic oxygen erosion in space. The atomic oxygen erosion combined defects assessment method studied in this paper provides valuable reference data for the subsequent serialisation of deorbit sail products and other typical lightweight film materials for space applications. • Introduction of Monte Carlo method for material surface damage simulation analysis. • Compare the performance of off-orbit sail surface materials under different protection treatments. • Standardize the process of space flexible film material performance evaluation system. [ABSTRACT FROM AUTHOR]

As part of the Artemis era of space exploration, space agencies will be working together with their industry partners to establish systems and infrastructure that enable sustained Lunar missions and develop capabilities for Mars. Lockheed Martin has a longstanding history of designing, building, and operating systems for deep space applications, from Viking to Orion. Lockheed Martin develops concepts, trades, and early system designs for Moon and Mars exploration to enable future missions. In 2016, Lockheed Martin presented a vision for achieving crewed exploration of Mars. Known as Mars Base Camp (MBC), this design reference mission envisioned a crewed vehicle in Martian orbit from which astronauts could perform excursions to Phobos and Deimos and perform telerobotic exploration of the Martian surface, including sample return. This concept served as an "existence proof" for a novel, practical, and affordable path to enable human exploration of the Martian system. In 2017, an update was published that included the production of propellant from water and a single-stage fully reusable Mars lander concept. Then in 2021, Lockheed Martin introduced a nuclear thermal propulsion (NTP) configuration of Mars Base Camp, a technology that will be greatly enabling for crewed missions to Mars. This paper matures the design of the NTP configuration of Mars Base Camp and explores the Mars mission capabilities that the NTP technology will enable. These investigations include optimized trajectories for faster transit to Mars and an analysis of abort scenarios that are made possible by the high thrust and performance of NTP. The results provide insight into the performance and capabilities that are realized when NTP technology is matured and integrated into a Mars transit vehicle. The paper will also describe initial surface infrastructure capabilities that significantly enhance sortie science operations, including mobility and power. The resulting updated Mars Base Camp serves as an example of what is possible with current technologies and technologies under development. The MBC reference mission builds on the capabilities that will be developed and demonstrated through Artemis missions, enabling a viable, sustainable long-term Mars exploration program, with the first crewed Mars mission possible in the 2030s. • Matures design of nuclear thermal propulsion configuration of crewed Mars mission • Explores Mars mission capabilities and trajectories enabled by nuclear propulsion • Describes initial infrastructure to support sustainable long-term Mars exploration [ABSTRACT FROM AUTHOR]

OMOTENASHI is a CubeSat that was launched by a NASA SLS rocket. Its mission was to demonstrate that a CubeSat could make a semi-hard landing on the Moon. A chemical propulsion system which provides high thrust and can be re-ignited is large and difficult to install on a CubeSat. Therefore, a small cold-gas jet makes orbit corrections with only a small ΔV. Moreover, a rocket motor with a high Isp and structural mass ratio is used for a lunar landing deceleration. The spacecraft consists of an orbiting module, a rocket motor, and a surface probe. This paper analyzes the trajectory design of the Moon landing and a method that increases the landing success rate under the tight constraints of the CubeSat. To reduce fuel and power consumption, the maneuver conditions during the lunar transition to reach the target landing site are evaluated based on several error factors. This study considers the trade-off between vertical and horizontal landings. Because active control is not possible for decelerating a rocket motor for the lunar landing, the optimal altitude at the starting maneuver is determined by considering such errors as attitude and timing. Finally, the paper discusses mission feasibility and the landing success rate. • OMOTENASHI was developed as the world's smallest moon lander that was launched by a NASA SLS rocket. • OMOTENASHI uses Rocket Motor to decelerate Surface Probe for a semi-hard landing. • This paper shows an innovative trajectory design method of the moon landing for a CubeSat as a secondary payload. • For a high landing success rate, the CubeSat must accommodate many constraints in selecting the landing site. • This method can be utilized for future CubeSat missions under tight resource constraints, such as a thrust and a power consumption. [ABSTRACT FROM AUTHOR]

ESA Solar Orbiter was launched from Cape Canaveral in collaboration with NASA on the February 10, 2020 to start its journey around the Sun, with the commissioning phase ending in April 2020. After commissioning, ESA set a tighter pointing requirement goal to further improve the performance in Fine Pointing Mode, in preparation of the upcoming start of Science Phase. The first part of this paper outlines the end-to-end process which allowed - in just six months – to update the Fine Pointing Mode Design, re-tune the controller with refined assumptions taken from flight data and robust methods and execute full validation of the new AOCS Software on the Avionics Test bench at Airbus Defence and Space. The second part of the paper describes the upload in flight of the new Software and the analysis of the on-orbit pointing performance, which showed a significant improvement with respect to the previous one, thus contributing to the success of the Mission's scientific goals. • Solar Orbiter Payload Instruments required a better pointing jitter performance (Relative Pointing Error) after launch • Root cause analysis on flight telemetry to identify driver contributors to pointing error • Fine Pointing Mode controller design and tuning updated with refined inputs based on flight data and with robust control methods • Controller uploaded in flight significantly improved the Spacecraft pointing performance [ABSTRACT FROM AUTHOR]

Safe operation of orbital platforms relies heavily on stability of fluid flows in microgravity, which are strongly affected by capillary forces. The paper considers multiphase fluid seepage through a porous medium. Questions of instability of the interface between liquids are discussed. The effect of capillary instability is shown, which was revealed during the processing of experimental data on the imbibition of a porous structure in microgravity conditions during parabolic flights. A mathematical model is proposed to describe the process of multiple imbibition of a porous medium under conditions of variable gravity. The model takes into account hysteresis, various models of capillary pressure are used during imbibition and drainage. The paper also considers a method for modeling the instability by transforming the constitutive equations. The method allows modeling additional blurring of the imbibition front. The results of modeling based on this method for capillary imbibition of a porous medium under microgravity conditions, as well as for the case when chemical interactions occur between flowing fluids, are considered. • The instability of the phase boundary during seepage is considered. • Flows are studied controlling safe operation of spacecraft fluid supply systems. • A new mathematical model for describing the multiple imbibition is proposed • The seepage with chemical interactions between the phases is considered. [ABSTRACT FROM AUTHOR]

Academic and industry experts build their careers on accurate and precise definitions and meanings of domain-specific terminology, whereas novice practitioners and lay-persons often conflate similar or similar-sounding terms to describe concepts in the same domain of interest. This is common in the space sector when individuals of different academic or practitioner backgrounds discuss "commercial" topics. In English, as in other languages, the word "space" is often followed by words such as "industry," "industry segment," "industrial base," "market," "ecosystem" and others. The practice of casually using words with ambiguous meanings may often be appropriate for informal discussion if it facilitates the communication and exchange of ideas and all parties generally understand the implications of the casual usage. This practice, however, is problematic for detailed and analytical applications requiring precision and order of thought. Grounded in contemporary organization theory and economic practice, this paper provides background discussions of different levels of analysis that characterize embedded hierarchical social systems of organizations as they pertain to space. This paper then recommends a hierarchical taxonomy for defining specific terms that describe organizations and activities within the space domain for use by practitioners, analysts and academics requiring more precise language to communicate their market analysis ideas. This taxonomy places specific terms within each level of analysis discussed and provides definitions of each, with the intention of stimulating critical discussion and potential modifications or improvements. • Using phrases like "space industry" can be vague and problematic for precise analytical tasks. • For accurate space terminology, understanding analysis levels in organizational systems is key. • This paper offers a taxonomy for clearer language in space market analysis communication. • Proper terminology aids in strategic decisions for everyone in the space sector. • Feedback on this proposed taxonomy is encouraged for potential enhancements. [ABSTRACT FROM AUTHOR]

With NASA's Artemis program currently underway, the necessity to have lunar structures that can house astronauts for months at a time is now greater than ever. This paper investigates current deployable lunar habitat concepts and finds a paucity of detailed papers in the literature. The hybrid deployable lunar habitat architecture considered here went through many design iterations to simplify and support its manufacturing. A novel deployment mechanism is developed utilizing the release of pressurized gas for an autonomous deployment sequence. A prototype was fabricated that demonstrated a successful proof of concept, leaving a robust foundation for future research. • delom Hybrid autonomously deployable structures for lunar habitats are studied. • Such structures have solid elements plus an inflatable component that deploys after landing. • Astronauts arrive at the Moon and have the structure waiting for them. • A key design challenge is the deployment mechanism. • A scale model is designed and manufactured to test concepts. [ABSTRACT FROM AUTHOR]

In recent years, thanks to advancements in commercial hardware technology miniaturization, small satellites have become more and more important and effective in the Space ecosystem. In previous papers it has been demonstrated that future smallest observation systems, operating at a lower altitude than traditional systems, have the potential for comparable or better performance, much lower overall mission cost, lower risk, shorter schedules, lower up-front development cost, more sustainable business model, mitigating the problem of orbital debris. Flying in very low Earth orbit requires the addition of a propulsion system capable of providing drag compensation. Thanks to its high specific impulse, electric propulsion is potentially well suited for this task. However, the use of an electric thruster often requires significant power that can be provided only with deployable solar panels. To minimize the drag, the solar panels should be aligned with the satellite orbital velocity. At the same time, to maximize the power input, the solar panel surfaces should point toward the Sun. In the majority of cases these two conditions are not met simultaneously, so an optimal trade-off is needed. This topic is investigated in this paper by the use of an orbital model of a generic satellite with deployable solar panels. The model is able to simulate any circular orbit of the satellite in any day of the year and calculate both the power input and the drag from the solar panels based on their inclination. An equivalent specific impulse is also defined taking into account the thrust that the electric thruster has to provide only to push the area of the solar panels required to power the propulsion system. The results show, as expected, that the sun-synchronous Dawn-Dusk (or Dusk-Dawn) orbit is the preferred one for the electric propulsion as the velocity vector and the Sun vector are almost orthogonal. In the other cases, at very low Earth orbit, even relatively small inclinations of the solar panels respect to the velocity vector have a significant impact on the drag produced, increasing the propellant consumption and the required power. Nevertheless, it is shown also that if the solar panels are aligned with the velocity, the power lost can be kept reasonable. The same analysis could also apply to a satellite in low Earth orbit that does not employ an electric thruster but still needs deployable solar panels because of other sources of high power consumption (e.g. the payload). • Deployable solar panels in Very Low Earth Orbit generally produce drag. • If aligned with velocity, ideally no drag is produced but power available reduces. • Penalty for best case (dawn-dusk) is 1–12%, for worst case (noon-midnight) is 47%. • For other panels positions part of the power absorbed and motor thrust are wasted. • Equivalent specific impulse and net power can be determined. [ABSTRACT FROM AUTHOR]

Existing exoatmospheric guidance laws become ineffective when a thrust loss fault (like one or more engines unexpectedly shut down) occurs in a launch vehicle, which makes the nominal orbit unreachable. This paper presents a new abort orbit determination process to determine an abort orbit which the launch vehicle is able to reach. It involves solving at most two optimization problems and the abort orbit may be circular or elliptical. We will present how to transform the nonconvex optimization problems into difference of convex (DC) programming problems which can be solved by existing methods with guaranteed convergence. The main contribution of this paper lies in proposing a relaxation technique to convexify a nonconvex control constraint and the combination of the introduction of new constraints and a relaxation technique to deal with the strict terminal constraints. Validity of the relaxation technique is ensured by theoretical analysis. A unique feature of the transformation process is that the feasible sets defined by both the control constraint and the strict terminal constraints are all enlarged. This brings a critical benefit that the resulting DC programming problems can be very efficiently solved. Thus, an abort orbit determination algorithm, which is based on DC programming, can be designed to very reliably and efficiently determine an abort orbit, which will be demonstrated by numerical examples. • A new concise and practical abort-orbit determination process is proposed. • Relaxation techniques and the introduction of new constraints are properly combined. • Validity of the relaxation techniques are theoretically proved. • An abort orbit can be determined by solving at most two DC programming problems. [ABSTRACT FROM AUTHOR]

Space Trafic Management (STM) has gained relevance in regional and global strategies when there is more significant congestion in space. Ongoing discussions in the United Nations Committee on the Peaceful Uses of Outer Space (UNCOPUOS) and other fora are reflecting the essence of STM in Europe, in the United States and others. In the US, the 2018 Space Policy Directive-3 on National Space Traffic Management Policy has become a focal eleemnt. More recently, STM was underlined in the European Union – European Space Agency Space Council Orientations of November 2020 as part of the European ambitions to achieve strategic autonomy. In February 2021, the European Commission launched STM as a flagship project under its' Action Plan on Synergies' between the civil, defence, and space industries, aiming to reinforce the European industry's competitiveness. In February 2022, a joint communication on a European Union approach on STM was published. In this regard, STM entails several technological, socio-economic, and regulatory dimensions with significant security implications as reflected in national and regional strategies. This paper focuses on a critical element: the cooperation between the civilian and the military and how organizations effectively address the dual-use aspect concerning STM development under the current legal and regulatory framework. Moreover, it examines how the dual-use element of STM influences space security of space technologies. Successful strategies with tangible outcomes require effective cooperation between the civilian and military actors. Hence, this paper seeks to identify potential legal and regulatory barriers and proposes solutions to accommodate divergent strategic interests and needs in STM future strategies. [ABSTRACT FROM AUTHOR]

The increasing number of man-made objects in near-earth space is becoming a serious problem for future space missions around the Earth. Among the proposed strategies to face this issue, and due to the passive and propellant-less character, electrodynamic tethers appear to be a promising option for spacecraft in low Earth orbits thanks to the limited storage mass and the minimum interface requirements to the host spacecraft. This work presents the roadmap that the Electrodynamic Tether Technology for Passive Consumable-less Deorbit Kit (E.T.PACK) is following to develop a prototype of a deorbit device based on electrodynamic tether technology with Technology Readiness Level 4 by the end of 2022. The paper illustrates the roadmap of the activities carried out at the University of Padova, where software and hardware have been prepared to validate some of the critical elements of the deorbit device. Specifically, the software tools include: (a) the software called "DEPLOY" that allows the computation of a reference trajectory for the deployment of the tether and the completion of sensitivity analysis of the deployment trajectory to key error sources; (b) the software called "FLEXSIM" that predicts the performances of electrodynamic tethers as a function of the system configuration employed; and (c) the software called "FLEX" that includes the dynamical effects of tether flexibility and provides important information on the dynamic stability of the system during deployment and deorbiting phase. The paper describes in detail the three software tools and provides results of a simulation showing how it is possible to deorbit a 24-kg satellite from an initial orbital altitude of 600 km in less than 100 days using a 500-m long tape-like bare tether. The team has also developed laboratory mock-ups and performed experimental activities to: (a) determine the tether mechanical properties; (b) test the functionality of mechanisms used to deploy the tether; (c) test the functionality of the attitude control assembly used during the deployment phase; and (d) validate a passive damper designed for dissipating the longitudinal oscillations of the tether and thus guarantee the stability of the system during both deployment and deorbiting phase. The paper provides a description of both the laboratory setup and the experimental activities performed to validate EDT technologies, including the damping capability of a compact passive-damping mechanism, showing how it can reduce consistently the peak forces up to about 80%. • Optimization of full-length deployment trajectory. • High fidelity simulations of deorbiting with electrodynamic tethers. • Experimental validation of technologies for tape electrodynamic tethers. • Close-proximity deployment testing with frictionless table and trajectory control. • Development and testing of passive in-line damper. [ABSTRACT FROM AUTHOR]

Aotearoa New Zealand's emerging New Space economy provides an opportunity for key actors to focus on space and sustainability issues beyond space debris. The conflict between competing definitions and paradigms of sustainability highlights the importance of diverse values, assumptions, and drivers of change that shape the normative understanding of space sustainability issues. This paper recognises that Indigenous knowledges and practices are in parallel with systems-thinking and transdisciplinary approaches to space and sustainability. The aim of this paper is to describe how current actions can have long term impacts on using and accessing space commercially, scientifically, and culturally. • Space sustainability requires a transdisciplinary approach to explore the definitions and values at various scales. • Space sustainability in New Zealand involves addressing issues beyond space debris. • Space sustainability requires addressing businesses in New Zealand, where sustainability frameworks are poorly implemented. • Space sustainability requires policy interventions beyond compliance with environmental legislation. [ABSTRACT FROM AUTHOR]

This paper aims to describe and assess the dispute settlement mechanisms available to States in the context of space and telecommunications activities. Disputes either between private actors or between private actors and States are beyond the scope of this research. The choice to examine mechanisms and procedures for the settlement of international disputes in these specific areas stems from an elective affinity between these sectors, given the existing interaction between them. At a first level, the paper highlights the importance of mechanisms to prevent the creation of international disputes, both in the field of space law (mainly through Art. IX of the Outer Space Treaty) and in the field of international telecommunications (Arts. 45 para. 3, 48 of the ITU Constitution). Further, the general scheme of peaceful settlement of international disputes, as enshrined in the UN Charter not only applies to the disputes in question but is further specified in the context of space and telecommunications law (Liability Convention, Art. 56 of the ITU Constitution), with a clear orientation towards the use of diplomatic rather than judicial means of resolution. Last but not least, if the settlement of disputes through diplomatic means fails, the preferred judicial mechanism is that of arbitration, as it is clearly demonstrated by the Claims Commission of Art. XIV LIAB, the PCA "Optional Rules for Arbitration" of space disputes, the mechanism of Art. 41 of the ITU Convention as well as the relevant "Optional Protocol". The specific options for dispute settlement show that, in view of the common interest of States in "international goods" of a technical nature, the actors involved are mainly interested in the non-occurrence of disputes rather than in their resolution through judicial means. And when they do so, they choose to resort to flexible resolution mechanisms (arbitration) rather than to ordinary international courts. • There is an elective affinity between telecommunications and space activities. • States are mainly interested in preventing disputes than to resolve those arising. • States opt for diplomatic resolution mechanisms instead of jurisdictional ones. • Ιf diplomatic settlement fails, the preferred judicial mechanism is arbitration. [ABSTRACT FROM AUTHOR]

The Surface Water & Ocean Topography (SWOT) mission is a joint NASA/CNES mission launched on December 16th, 2022 dedicated to oceanography and continental hydrology. Once the scientific mission is over, SWOT will have to perform a controlled re-entry to comply with the French Space Operation Act (FSOA). Indeed, the launch authorization is subject to the respect of the FSOA, fully applicable since 2020. The main requirement applying to the end-of-life of the satellite consists in limiting the casualty area and then the probability of having a victim on ground under the threshold of 10−4 in case of a natural re-entry. With a mass over two tons and a casualty area of about 47 m2, SWOT cannot comply with this requirement. Thus, the platform has been designed with a specific propulsion system to carry out end-of-life maneuvers leading to an atmospheric re-entry and impact in the South Pacific Ocean Uninhabited Area (SPOUA). CNES is in charge of defining the re-entry maneuvers strategy, performing end-of-life operations and evaluating the risk of casualties in case of failure during nominal strategy. This paper addresses these topics. The computation of the ten maneuvers used to decrease the perigee altitude and finally target the SPOUA is performed using the DOORS tool. Before describing the maneuvers strategy, this paper presents an overview of the current functionalities of this tool, which has been already used during the five ATV missions and ASTRA-1K de-orbit twenty years ago. DOORS is mainly used for mission analysis purposes, but for end-of-life operations it can be activated by the operational Flight Dynamics System (FDS). Finally, the casualty risk estimation relies on the use of two other CNES tools, DEBRISK and ELECTRA, whose methods are presented. The list of surviving fragments has been computed using the new DEBRISK-V3 methodology in case of controlled and uncontrolled re-entry. The risk of human casualty is evaluated using ELECTRA and its Monte Carlo simulations. This tool generates numerous trajectories of the satellite and its fragments with various dispersed variables. The impact locations compared to an up-to-date grid of the World population lead to the probability of victims. The complete risk of SWOT controlled re-entry encompasses both the risk related to a failure during the last re-entry maneuver and the risk associated with the natural re-entry in case a failure occurs before this last burn (during the mission or during the first re-entry maneuvers). Both of them are weighted by their probability of occurrence. • SWOT controlled re-entry strategy to fulfill French regulation. • Robustness test of the nominal strategy made of 10 maneuvers. • Latest aerothermodynamic models used to build the list of surviving fragments. • Probability of victims considers contingencies during the whole re-entry operation. • Combination of several CNES softwares: DOORS, PSIMU, DEBRISK, ELECTRA. [ABSTRACT FROM AUTHOR]

The National Aeronautics and Space Administration (NASA) has been involved in the development and maturation of metal additive manufacturing (AM) for space applications since the late 2000's. Several efforts have focused on the understanding of AM processes through material characterization and testing, standards development, component fabrication, and infusion into propulsion development and flight applications. NASA matured commonly used aerospace alloys from various alloy families (Nickel, Copper, Stainless and Steel, Aluminum, and Titanium-based) through detailed AM process and heat treatment characterization, in addition to mechanical and thermophysical testing. While these alloys are actively used in many propulsion applications, there is a need for ongoing AM optimized alloys using integrated computational materials engineering (ICME) and process development for high performance applications. The applications targeted are liquid rocket engines; advanced propulsion systems; and in-space propulsion with high heat fluxes, high pressure, and/or that use propellants that can degrade alloys (e.g., hydrogen). This paper highlights the characterization and physical properties of the more common AM alloys using laser powder bed fusion (L-PBF) and laser powder directed energy deposition (LP-DED) processes. Additionally, this paper discusses some of the ongoing novel alloy development and maturation using AM for use in these harsh environments, such as GRCop-42, GRCop-84, NASA HR-1, GRX-810, and C-103. The results from these processes demonstrated that AM could enable rapid development, and that optimized alloys could be developed using ICME, yielding higher performances. These alloys have undergone modeling, fundamental metallurgical evaluations, heat treatment studies, detailed microstructure characterization, and mechanical testing campaigns. This, combined with direct application-specific component fabrication and hot-fire testing, enabled the increase of the Technology Readiness Level (TRL) through high duty-cycle testing. A background and overview of these novel AM-enabled alloys and AM processing developments including metallurgical and mechanical property studies is presented here. The latest advancement in the parallel component development and hot-fire testing and future developments for these alloys is also discussed. • Additive manufacturing (AM) has enabled development of novel alloys for extreme propulsion environments. • NASA has matured alloys including GRCop-42, GRCop-84, NASA HR-1, GRX-810, C-103. • Characterization, material properties, and hot-fire test application data has been made available. [ABSTRACT FROM AUTHOR]

The purpose of this paper is to study the development of space-age aesthetics culture and arts in seven decades since 1950s. The first known use of "space-age" was in 1946, and was defined as "of, relating to, or befitting the age of space exploration." A more popular definition for the "space age" is "The space age is the time period in human history related to the space race, space exploration, and space technology. Most people say this started when Sputnik 1 was launched in 1957, and continues to today. This has influenced culture such as in movies, television, music, art, and architecture." Culture, in the anthropology, is the patterns of behaviour and thinking that people living in social groups learn, create, and share. Culture distinguishes one human group from others. It also distinguishes humans from other animals. A people's culture includes their beliefs, rules of behaviour, language, rituals, art, technology, styles of dress, ways of producing and cooking food, religion, and political and economic systems. Therefore, the space-age art is one of the major elements of the space-age culture of human beings. On the other hand, the space exploration has significant inspiration and influence on the aesthetics art. In particular, the following aspects shall be studied in this paper: architecture and exhibition craftworks. • The space-age architecture and exhibition craftworks were investigated. • Five space-age architecture shapes investigated are disc, rocket, tower, spherical and combined. • Three space-age installation craftworks studied are sky waves, embroidery, planet models, etc. • Many space-age architectures are destroyed due to many reasons. • Hopefully the young generations can be inspired to carry forward these precious heritages. [ABSTRACT FROM AUTHOR]

Partially reusable launch vehicle design has raised a great interest in order to minimize the costs of space transportation by recovering and refurbishing the first stage. Several configurations such as toss-back architectures are now operational. These concepts require additional rocket propellant in order to carry the boost-back and landing maneuvers, that induce several losses in terms of payload mass performance for the ascent mission. In order to limit these losses, this paper focuses on the design of two types of architectures to carry out the return-to-launch-site mission. The first, named glide-back, uses the main propulsion system of the first stage to perform a boost-back burn and returns to the landing site by a gliding mode using additional lifting surfaces. The second, named, fly-back, performs the return-to-launch-site mission using several air-breathing engines located in the nose of the first stage combined with lifting surfaces. In this paper, the design of a reusability kit, allowing to provide the first stage with both expendable and reusable capabilities, is investigated. This kit is composed of the lifting surfaces, the nose including the air-breathing propulsive system (for fly-back configuration), landing gears and additional avionics. This paper addresses the design of such reusability kits and presents the mission specifications, the design process relying on Multidisciplinary Design Optimization techniques, the vehicle performance and the optimal trajectories. Different analyses (aero-propulsive, optimal control, operational considerations, etc.) are detailed and trade-offs between the studied configurations are assessed. [Display omitted] • Reusable launch vehicles. • Multidisciplinary design and optimization of reusability kit. • Inclusion of aeronautical technologies in launch vehicle design. [ABSTRACT FROM AUTHOR]

The detection and observation of space debris in Low Earth Orbit is generally carried out through the use of ground based radars and telescopes. These instruments allow for a precise reconstruction of the space debris trajectories, and therefore represent a key asset for planning avoidance maneuvers when threats of collisions are predicted. The recent deployment of mega-constellations, with the consequent increase of the number of satellites, imposes new challenges in terms of simultaneous tracking capability and readiness of the current space situational awareness systems. This adds to the current need to track small and dull objects to further mitigate the probability of triggering cascade collisions. However, ground based observations are limited due to their intrinsic sensibility to atmospheric refraction, their diurnal inoperability and their dependence on meteorological hazards. This paper proposes to study the feasibility and the benefits of a potential deployment of a constellation of CubeSats in Low Earth Orbit, to acquire optical observations of space debris with enhanced accuracy, as part of the ORCA mission: Orbit Refinement for Collision Avoidance. Here, the focus is on the optical design of the payload instrument to be integrated onboard of the orbiting platforms. The study trades-off the current state-of-art of optical detection technologies, by assessing their performance against a set of specific requirements: (a) the minimization of the uncertainty associated to the image resolution; (b) a field of view that maximizes the extent of the monitored area; (c) an optimal exposure time to avoid under or overexposure of the image; (d) minimization of the effects of light diffraction and above all, (e) the maximization of the signal to noise ratio to detect the smallest and dullest objects possible. Several configurations of optical systems are then chosen as suitable for the ORCA constellation, also considering the system design implications of its integration into a CubeSat, such as size requirements. Commercial Off-The-Shelf hardware are explored and performances of the optical system are evaluated through numerical simulations in order to estimate the detectable sizes of space debris while taking into account their potential distances from the sensor. This paper concludes with estimates of the impact of the ORCA mission on space situational awareness for decades to come. • We defined requirements for an optical system for space debris surveillance. • We formulated conditions allowing detection of space debris from optical images. • We explored feasible design of optical systems for a CubeSat size implementation. • We compared our results with available commercial-off-the-shelf optical systems. [ABSTRACT FROM AUTHOR]

In this paper, different compressible correction strategies for k-ω turbulence model are proposed. The dilatation dissipation term, which is significant in compressible flow, is taken into account and modeled in the Favre-averaged k-equation and ω-equation. We propose correction model for the dilation dissipation term based on the gradient Mach number in this paper, the coefficients of which are calibrated from experimental data on compressible turbulent mixed layers, and compare it to the Wilcox, Sarkar, and Heinz correction models. In slope-cavity flow and oblique shock wave/boundary layer interactions, the performance of these corrections is evaluated and compared to experimental data. The new correction has a similar or even better effect in separation prediction, heat flux prediction, and other aspects, and the uncertainty of model form and coefficients can be investigated further. • A general compressible correction model is conducted with the consideration of dilatation and structural compressibility. • The experimental data of compressible turbulent mixed layers is employed to calibrate the model coefficients. • Verification and validation are carried out in slope-cavity flow and incident oblique shock wave/boundary layer interactions. • The performance of the model and its comparison with other correction models are discussed. [ABSTRACT FROM AUTHOR]

This paper treats the problem of positioning and navigation in the absence of GNSS. Given recently raised vulnerabilities for GNSS both on Earth and in space, the work revisits the old problem of the sextant in new light. The combination of stars and planetary horizons was the popular tool for autonomous navigation on-board spacecraft, but given the rise of GNSS receivers, this solution has been largely disregarded. New spacecraft missions beyond Earth have made new progress in visual-based navigation. The work utilises these new methodologies in the scenario that RF-derived positioning is unavailable, achieving a performance below 100 m. An additional important consideration is the development of new navigation infrastructure in LEO and the Moon. Current methods seek to use GNSS and RF-derived sources for orbit determination, however, to be seen as 'redundancy infrastructure' for critical Earth and beyond applications, these signals cannot be the primary and only source of navigation reference. This paper utilises derived performances for visual-based methods and applies them to the ranging service problem. Most user scenarios in maritime, aviation and lunar domains are satisfied. This work illustrates implementation using the simple architecture of a star tracker camera. Known as CROSS, the technology is a new navigation tool in development by the University of Sydney. As a star tracker is common device to many spacecraft platform, it simplifies implementation, given that redundant systems are always not a priority to the manufacturer. • Vision-based approach for orbit determination using only stars and the horizon. • Application of this approach to future LEO PNT mega constellations. • Comprehensive review of requirements by different market segments. • Proposed methods to overcome key challenges including stray light and planetary atmospheres. [ABSTRACT FROM AUTHOR]

The cooperative orbit determination (OD) using inertial inter-spacecraft measurements is of great value for the constellation. Previous studies concluded that the cooperative OD system in the two-body dynamics is observable using the inertial angle measurement. This paper performs an observability analysis for the inertial angle-only cooperative OD system based on the observability matrix (OM), and finds that the previous conclusions do not hold for some special configurations. Eight special configurations are proposed, and for each configuration, the unobservable orbit elements are analytically revealed by analyzing the linear relationship among different columns of the OM. Moreover, a method of improving the observability of the two-spacecraft system is proposed by adding another spacecraft. An extended Kalman filter is employed to numerically illustrate the observability of the proposed special configurations and to show the relationship between the observability and the convergence ratio of estimations. Monte Carlo simulations show that the convergence ratios of the observable orbit elements and state combinations are higher than 97%, while convergence ratios of the unobservable states are lower than 70%. With one non-planar elliptic orbit added, all the elements of the three-spacecraft system are observable, and the estimated errors converge within 6 h. The results found in this paper can provide a reference for the design and management of the constellation. • Observability analysis for the angle-only cooperative OD is performed. • Eight unobservable configurations are revealed using the observability matrix. • A method of improving the observability is proposed by adding another spacecraft. [ABSTRACT FROM AUTHOR]

The use of modular, yet reusable and interchangeable components offers flexibility for system designers and thus leads to a higher degree of reconfigurability with cost effective, adaptable and scalable solutions. The development of a robotic system remains, however, a complex task with increasing demands depending on the targeted capabilities, functionalities and operational constraints of the robot. This requires expert knowledge in mechanical, electrical and software engineering, as well as tools that can collect, process and use relevant data. As a result, the amount of available and managed information is increasing, while its relevance, usability, and knowledge that can be extracted from it are consequently decreasing: a situation described as information overload paradox. The design of a modular system can be based on the use of standardized modules with universal interconnects, but it should be aligned with a strong formalism, e. g. the semantic representation of components, to allow for a better management of the overall complexity. An ontology-based knowledge representation, for instance, which provides domain and application specific knowledge for the design of robotic systems, can be used as a method of storing and sharing data in a uniform, machine-readable and standardized way. In this paper, we introduce the Knowledge-based Open Robot voCabulary as Utility Toolkit (korcut) as a core component to support the ontology-driven development of robotic systems as part of a reference implementation of the Q-Rock development cycle. Q-Rock gathers methods to assist users in designing robot configuration by: (a) automatically exploring robot capabilities based on robot hardware, (b) proposing robot designs to meet the user's needs, (c) refining the proposed robot designs. We use korcut to improve the robot design process in orbital and planetary robotics from the requirements definition to the development phases of complex modular robotic structures (e. g. composition of an unmanned ground vehicle with robotic arms). This is implemented by embedding semantic component descriptions in a state-of-the-art open-source 3D modeling software. Furthermore, the korcut ontology family includes various sub-ontologies to address specific space-related tasks, such as the modeling of a Standard Interconnect (SI) for On Orbit Services (OOS) and Orbital Factory. The ontology design follows criteria that have been derived from a survey conducted as part of this work. This paper also covers an evaluation of its current applications from a methodological, knowledge representation, and software tool perspectives. In the application part, we introduce the use of the ontology to model a multi-functional SI that enables mechanical connections between various (modular) robotic components, as well as transfer of power and data. Finally, we provide a critical analysis of our work and outline its future work. • The development of complex robotic systems in space domain needs expert knowledge. • Increasing information about a robotic system can lead to an information paradox. • The information paradox can be improved by standardized and inferable knowledge. • Ontology is used in a case study to facilitate the robot development workflow. • Ontology is used to provide a metric for evaluating a space interface. [ABSTRACT FROM AUTHOR]

On-Orbit-Servicing missions aim to improve the operative life of orbiting satellites. Among the studied technologies, the most promising are satellites equipped with one or more robotic arms (space robots). One robotic arm might be employed to establish a rigid connection with the target satellite, while the other performs the servicing operations. Nowadays, the control of the capture tool is managed by the computer of the space robot. However, it is possible to distribute the control tasks between the capture tool and the space robot. For this purpose, this paper presents the development of the SMArt Capture Kit (SMACK). SMACK is a standalone capture system equipped with sensors, actuators, and an integrated computer. In this fashion, the capture tool executes the guidance and navigation algorithms, and the control of the actuators; so that the space robot acts as the "positioning" system for the capture tool. The paper provides a description of all the subsystems of SMACK and their experimental validation. The tests of SMACK at the system level, confirm its capabilities to autonomously perform the capture the target interface. • Smart capture tool. • Autonomous tool with integrated sensors, actuators and computer. • Distributed control. • Use of fuzzy logic to develop the model of a custom sensor. • Use of fuzzy logic to create the approach trajectory. [ABSTRACT FROM AUTHOR]

By the time of the 2022 IAC, NASA's Mars2020 Perseverance rover will have spent 18 months on the surface of Mars, during which time the MOXIE experiment (M ars OX ygen I SRU E xperiment) will have produced oxygen at night and in the day during both the annual maximum and minimum atmospheric density periods, as well as at many other times during the year. MOXIE is the first demonstration of the use of indigenous resources (ISRU = In Situ Resource Utilization) on the surface of another planet. This paper will present a summary of what MOXIE has accomplished, how its performance on Mars has changed with time, and plans for the future. The paper will also present results from an optimization study of a human-scale MOXIE-type system capable of providing the oxidizer for a 6–person Mars Ascent Vehicle. As an experiment carried inside the rover, MOXIE had to satisfy many constraints that would not apply to an independent, full-scale system. Other potential oxygen-producing technologies should be compared to the optimized human-scale system results summarized in this paper rather than to a simple linear scaling of the mass, power consumption, and oxygen production rate of MOXIE. • MOXIE (M ars Ox ygen I SRU E xperiment) has successfully demonstrated a solid oxide electrolysis technology for extracting oxygen from the carbon dioxide-rich Martian atmosphere. • The in-situ production of oxygen is necessary for sustainable human Mars exploration, mainly for the combustion process powering a rocket to return from the Martian surface to space. • Bringing all this oxygen to Mars from the Earth would require multiple launches of costly Saturn 5-class rockets, which would be a serious impediment to sustainable human exploration of Mars. • In-Situ Resource Utilization (ISRU) for planetary exploration (i.e., "living off the land") has been discussed and written about for decades, but MOXIE is the first experiment to actually demonstrate ISRU on the surface of another planet. • The article summarizes the purpose and method of operation of MOXIE, shows the results of oxygen production so far, and shows how operating MOXIE informs the design of future larger-scale oxygen-producing systems. [ABSTRACT FROM AUTHOR]

This paper discusses a system for Titan exploration enabled by nuclear fusion power. Titan is one of the most interesting locations in the solar system with a thick atmosphere, surface oceans, under-ice oceans and complex terrain. This paper provides a conceptual design of a fusion-powered system to explore many parts of Titan and enable the use of high-power instruments. The design includes a fusion-powered orbital transfer vehicle and an electric Titan science aircraft. A Direct Fusion Drive (DFD) propulsive engine could bring a sizable spacecraft to Titan orbit in less than two years. A second fusion reactor, configured as a closed-loop power generator, would be used for an electric Titan science aircraft. Both reactors are based on the Princeton Field-Reversed Configuration (PFRC) concept which combines an FRC with a magnetic mirror. PFRC uses a novel radio-frequency plasma heating system and deuterium-helium-3 fuel. A lower temperature plasma flows around the closed-field FRC region removing the fusion products. In the DFD propulsive configuration, this secondary flow permits direct and variable thrust and exhaust velocity. The science aircraft would do a powered entry to Titan and then have the capability to fly anywhere on the moon at subsonic speeds. The DFD-powered transfer vehicle would allow the in-orbit transfer stage to change inclination as needed to cover different areas of the surface. • A PFRC fusion rocket could deliver 1 MT payloads to Titan in about two years. • A PFRC fusion reactor could power an electric aircraft to explore Titan's surface. • Fusion power allows the aircraft to utilize novel high-power instruments. • The PFRC transfer vehicle doubles as a science platform and communications relay. [ABSTRACT FROM AUTHOR]

All current rehydratable food and drink items prepared for U.S. astronauts on missions to the International Space Station (ISS) must be compatible with the station's potable water dispenser, called the rehydration station. This highly validated system ensures that water is compartmentalized and safely transferred to custom-made food and beverage pouches to prevent free-floating food particles and water in microgravity. However, the food and beverages packaged in these single-purpose pouches require complex, time-consuming, and costly procedures before flight, which is not feasible for future commercial space stations seeking more affordable life-support systems. Furthermore, even though a vast number of rehydratable food and beverage items available at supermarkets have appropriate nutritional content and shelf life for spaceflight, their incompatible consumer packaging renders them unusable by astronauts. In this paper, we present a low-cost solution that draws on NASA technology to adapt these commercially developed food and drink items for spaceflight and makes them compatible with the water dispensing unit and consumption in a microgravity environment. The 3D-printed adapter is based on the NASA septum adapter with the addition of a piercing end and one-way barb that stabilizes the adapter inside the food pouch and ensures it is not released due to the force generated by the water when it is dispensed. The dimensions of the adapter are 20.00 (L) x 12.50 (W) x 37.24 (H) mm and the mass is approximately 1.5 g. The design is compatible with at least three types of packaging commonly used for commercially available rehydratable foods: kraft paper and polymer laminate (KP), polymer aluminum laminate (PA), and compostable polymer laminate (CP). Based on these three packages, we present a table of commercial items that could be available to astronauts if the adapter were to be used. To validate the adapter's use in the space environment, we also present data on the structural properties against loads experienced during launch and re-entry, as well as two functionality tests to determine piercing and rehydration effectiveness. The results show that the prototype of the adapter meets structural specifications for spaceflight and can effectively pierce KP and PA materials with minimal tear propagation. The rehydration test with KP showed that the adapter can effectively rehydrate a typical dehydrated food package. In summary, the adapter is an effective method to increase menu diversity while simultaneously eliminating costs associated with space food preparation and re-packaging. • A new proof-of-concept adapter for food and beverage rehydration in space was demonstrated. • Adapters can be made compatible with U.S. galley equipment on the ISS. • Enabling commercial food on the ISS could significantly reduce costs of human spaceflight. [ABSTRACT FROM AUTHOR]

Tether deployment of the space tether system is the first step of related on-orbit experiments, which requires an advanced and reliable control scheme to realize rapid deployment without overshoot. In this paper, we propose a fractional-order fixed-time controller based on hierarchical sliding mode control to be applied to the deployment process of the tether system. The application of fractional calculus improves the local effect of the controller. Through strict mathematical derivation, we prove that the proposed control law satisfies the global fixed-time stability. So the closed-loop system has the performance of fixed-time convergence. Part of the simulation uses a fractional sliding mode control law for comparison, which further illustrates the superiority of the control law proposed in this paper. In view of the uniqueness of tether system dynamics, the research ideas in this paper are also of reference value to similar under-actuated systems. • Aiming at the underactuated dynamics, the hierarchical sliding mode is adopted. • Multiple sliding surfaces are used, and they all satisfy the fixed time convergence. • The rapid deployment of tethered satellite without overshoot is realized. [ABSTRACT FROM AUTHOR]

One of the main challenges for spaceflight safety is preventing crack growth. These defects can occur both in the skin of spacecraft and during the combustion of solid fuels. An important problem is the propagation of cracks at high temperatures. This paper proposes a study of a single polygonal crack under a constant load perpendicular to the crack plane. The solution is constructed numerically using the boundary element method. The main goal of the work is to calculate the stress intensity factors in the vicinity of the crack. The first intensity coefficient was calculated, and calculations were carried out for various crack parameters. • A symmetrical and asymmetrical polygonal cracks were studied. • An asymmetrical crack is more resistant to growth than a symmetrical one. • For a symmetrical crack the stress intensity factor increases in proportion to the size of the crack. • For an asymmetric crack, the stress intensity factor does not change on increasing one shoulder. [ABSTRACT FROM AUTHOR]

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

This paper explores the relevance of a new class of upper stages commonly referred to as Orbital Transfer Vehicles (OTVs) or Kick-Stages, and the importance of green propellants in the development of such systems. Orbital Transfer Vehicles are currently developed for diverse applications by many entities and are expected to have a major impact in the future of space missions and overall space sector. Such systems are commonly connected to the topic of On-Orbit Servicing with the promise of unlocking new missions, including active space debris removal, multi-payload to multi-orbit delivery, in-orbit experiments, in-orbit refuelling and other generic services. Green propulsion is an increasingly prevalent trend in the space industry that has experienced significant growth in recent decades with the main objective of identifying suitable alternatives to commonly used, toxic liquid propellants for in-space applications. The study investigates the synergies between the two themes, underlining the advantages that a widespread use of green technologies could bring to the overall sector, but also examining the final benefits to the system design of OTVs. Due to the numerous and challenging propulsive system requirements associated with these systems, green technologies are analysed to determine their advantages and disadvantages in comparison with current concepts, outlining current trends and expected future developments. The focus is on attainable performances of propulsion systems with respect to the required dry and inert mass. Various architectures of the different green alternatives based on hydrogen peroxide and the more peculiar self-pressurized propellants are proposed, exploring green propellants-based designs that can offer synergies and advantages over classical ones. These new architectures offer performance comparable to classical and widely available toxic alternatives and present the important advantage of using easy-to-handle propellants. The development of such new options has the potential to enable innovative future designs by improving mass and size requirements of new propulsion systems, thereby enhancing the current landscape. • Orbital Transfer Vehicles are emerging in the space sector as enabling technology. • The propulsion systems of Orbital Transfer Vehicles are crucial for the success of their missions. • Green Propellants offer various synergies possibly advantageous for new OTV designs. • Green Propellants perform well against toxic alternatives, but have less heritage. • Multiple design choices based on green propellants are proposed and compared. [ABSTRACT FROM AUTHOR]

Orbiting solar reflectors (OSRs) are flat, thin and lightweight reflective structures that are proposed to enhance terrestrial solar energy generation by illuminating large terrestrial solar power plants locally around dawn/dusk and during night hours. The incorporation of OSRs into terrestrial energy systems may offset the daylight-only limitation of terrestrial solar energy. However, the quantity of solar energy delivered to the Earth's surface remains low due to short duration of orbital passes and the low density of the reflected solar power due to large slant ranges. To compensate for these, this paper proposes a constellation of multiple reflectors in low-Earth orbit for scalable enhancement of the quantity of energy delivered. Circular near-polar orbits of 1000 km altitude in the terminator region are considered in a Walker-type constellation for a preliminary analysis. Starting from a simplified approach, the equations of Walker constellations describing the distribution of the reflectors are first modified by introducing a phasing parameter to ensure repeating pass-geometry over solar power farms. This approach allows for a single groundtrack optimisation to define the constellation, which was carried out by a genetic algorithm for single and two reflectors per orbit with an objective function defined as the total quantity of energy delivered per day, to existing and hypothetical solar power projects around the Earth. When full-scale constellations are considered with a number of reflectors, the quantity of solar energy delivered is substantial in the broader context of global terrestrial solar energy generation. • Terrestrial solar energy enhancement by constellation of orbiting reflectors is proposed. • Scalable, predictable and continuous solar energy from space is aimed. • Walker constellation equations are modified for repeating reflector passes. • Groundtrack optimised by genetic algorithm to maximise daily energy delivery. • Quantity of daily solar energy delivered is considerable in the global context. [ABSTRACT FROM AUTHOR]

Electric Solar Wind Sails (E-Sails) are a new type of spacecraft propellantless propulsion system that gathers its energy from solar wind protons and is potentially useful for interplanetary missions. Although optimal interplanetary trajectories have been the subject of thorough research, the substantial variability of the solar wind necessitates the adoption of active guidance strategies, an area that has received significantly less scholarly attention. This paper proposes guidance algorithms for E-Sails based on Model Predictive Control (MPC), a modern control methodology based on online re-planning of the trajectory. To this end, first, properties of E-sail time-optimal orbits are studied applying Pontryagin's Minimum Principle, and then time-optimal orbits for missions to Mars and Jupiter are computed via direct transcription methods. Next, solar wind perturbations are modeled, posing a challenging saturation problem due to their high variability. Guidance strategies based on Shrinking Horizon and Receding Horizon Model Predictive Control (RHMPC) are developed, analyzed and compared using Monte Carlo simulations, successfully implementing MPC to E-sail guidance. Lastly, the RHMPC strategy is successfully tested with accurate historical solar wind data from the WSA-Enlil model. • Solar wind fluctuations implies need of guidance in E-sail interplanetary missions. • Optimal planning for E-sails is performed using direct and indirect methods. • Control saturation due to solar wind pressure is reduced by leaving control margins. • Model predictive control guidance methods are compared with Monte Carlo analysis. • Receding horizon strategy in tested with semi-empirical historical solar wind. [ABSTRACT FROM AUTHOR]

With the ever increasing number of active satellites in space, the rising demand for larger formations of small satellites and the commercialization of the space industry (so-called New Space), the realization of manufacturing processes in orbit comes closer to reality. Reducing launch costs and risks, allowing for faster on-demand deployment of individually configured satellites as well as the prospect for possible on-orbit servicing for satellites makes the idea of realizing an in-orbit factory promising. In this paper, we present a novel approach to an in-orbit factory of small satellites covering a digital process twin, AI-based fault detection, and teleoperated robot-control, which are being researched as part of the "AI-enabled Cyber–Physical In-Orbit Factory" project. In addition to the integration of modern automation and Industry 4.0 production approaches, the question of how artificial intelligence (AI) and learning approaches can be used to make the production process more robust, fault-tolerant and autonomous is addressed. This lays the foundation for a later realization of satellite production in space in the form of an in-orbit factory. Central aspect is the development of a robotic AIT (Assembly, Integration and Testing) system where a small satellite could be assembled by a manipulator robot from modular subsystems. Approaches developed to improving this production process with AI include employing neural networks for optical and electrical fault detection of components. Force sensitive measuring and motion training helps to deal with uncertainties and tolerances during assembly. An AI-guided teleoperated control of the robot arm allows for human intervention while a Digital Process Twin represents process data and provides supervision during the whole production process. Approaches and results towards automated satellite production are presented in detail. • An automated production system for modular small satellites is proposed. • Central element is a force-sensitive robotic assembly, integration & testing system. • AI and machine learning approaches are applied to make the process more robust. • A digital process twin supervises the autonomous production. • A multi-modal shared control approach allows human teleoperation if required. [ABSTRACT FROM AUTHOR]

This research paper presents a thorough examination of shock wave train phenomena in various duct structures, utilizing a machine learning-optimized k-ω model. The focus of the study is to apply machine learning techniques to adapt the constant coefficients of the k-ω turbulence model, improving its accuracy and computational efficiency in capturing the dynamics of shock waves. An important aspect of this investigation is the impact of diverging sections within the duct, specifically how changes in the divergence angle, while maintaining a constant ratio of the exit area to the throat area, affect compressible flow parameters, shock wave positions, and other related characteristics of shock trains. The study systematically explores these effects and provides fresh insights into the behavior of shock wave trains under different divergence conditions. In a departure from the usual practices in supersonic research, this study compares the phenomena of shock wave trains in rectangular and circular duct geometries. The findings contribute valuable data and analyses to a field that has traditionally focused on rectangular configurations. The results indicate that the shape of the duct has a significant influence on the shock wave train, emphasizing the importance of considering geometric diversity in the study of supersonic flows and shock wave phenomena. This research sets the stage for further investigations into the interaction between duct geometry, shock wave propagation, and the dynamics of compressible flows. • Improved k-ω SST model accuracy to 95.5 % with machine learning. • Longer duct diverging sections intensify the first lambda shock. • Duct shape critically affects shock wave and boundary layer interactions. [ABSTRACT FROM AUTHOR]

This article presents a study on design and testing of space sails using Al-polyurethane Shape memory polymer. The properties of the polymer are simulated and tested in the design phase, and the results are compared with the fabricated sail, which is considered nominal. The efficiency of the polyurethane sail is found to be approximately 86.5%. The controllability of the sail is enhanced using a heating system that uses light rays to heat the polymer at specific control spots. The relative deflection at these spots causes a deviation in solar momentum, resulting in a controlled displacement of the sail. This control system is found to be 42% more efficient than conventional vane-type deflection systems. The paper also presents the fabrication of a ground testing apparatus designed to test the sail specimen in laboratory conditions. The apparatus is capable of producing a vacuum chamber pressure up to −400 mm/Hg and includes a magnetic levitation platform with strong permanent magnets to neglect local gravitational effects. The platform provides support for fixing the solar sail frame made from carbon fiber composite material. The testing apparatus consists of a 1000-W halogen light source with intensity ranges of approximately 200,000 lux. Sail material with a reflective surface is exposed to controlled luminance under vacuum conditions to measure the deflection of the sail. Deflection values are recorded by means of a resistance type strain gauge and strain indicator arrangement. Plotting deflections with time period gives the acceleration of the sail inside the test apparatus. A coolant chamber with a ceramic base is also provided to reduce temperature effects inside the vacuum chamber. In conclusion, this study presents a promising approach for the design and testing of solar sails using shape memory polymer and a ground testing apparatus. The results of this research can contribute to the development of more efficient and versatile solar sail technology for fuel-free aerospace propulsion. • A solar sail can travel to stars and planets without carrying fuel. • AI-polyurethane shape memory polymer is used for testing the space soils. • The ground-based testing is used to predict sail movement accurately. • A coolant chamber reduces the temperature within the vacuum chamber during testing. • Simulates and compares the properties of the polymer with the fabricated sail. [ABSTRACT FROM AUTHOR]

Laser-induced plasma ignition is a new ignition method to achieve stable and reliable ignition of liquid rocket engines. It has the advantages of accurate and controllable ignition position, broadening ignition lean-burn limit and small disturbance to flow field. In this paper, the laser ignition process and combustion characteristics of gas oxygen/gas methane under rocket engine conditions are studied from four aspects, namely different ignition positions, ignition energy, fuel-oxygen momentum ratio and chamber pressure. Using a high-speed camera, the dynamic process of ignition was captured by the shadow method, and the generation and propagation of laser-induced flame nuclei were recorded. The results show that at the airflow velocity of 300–400 m/s, the time required for the flame kernel to develop to the entire combustion chamber is about 300 μs, and the time for the flame to reach a steady state is tens of milliseconds. When the ignition energy is increased by 1.5 times, the time required for the flame kernel to develop from the recirculation zone to the mainstream zone is shortened to 1/3 of the original; with the increase of the fuel-oxygen momentum ratio, the shape of the initial flame kernel changes from round to oval, and the development speed of the flame kernel also increases. The fuel-oxygen momentum ratio increases from 0.59 to 2.75, and the time required for the flame kernel to develop from the recirculation zone to the mainstream zone decreases from 167 μs to 56 μs＀ Except chamber pressure, other factors have no effect on the pressure rise rate of the ignition process, and the higher the chamber pressure, the faster the pressure rise rate; in order to avoid deflagration, the shear layer is more suitable as the ignition position. With the increase of ignition energy, fuel-oxygen momentum ratio and chamber pressure, the development speed of flame kernel increases, but the shape and development process of flame kernel will be different under the influence of different factors. • Laser ignition is carried out in a highly complex engine environment with high pressure and non-premixed flow field. • The higher the chamber pressure is, the faster the pressure rises during ignition. In order to avoid deflagration, the shear layer is more suitable for ignition. • The trend of flame kernel development rate increases with the increase of ignition energy, fuel-oxygen momentum ratio and chamber pressure. [ABSTRACT FROM AUTHOR]

The study of unsteady combustion modes and, in particular, the deflagration-to-detonation transition, is of particular importance for the aerospace industry. Knowledge of the basic mechanisms, as well as the criteria for the development of detonation, helps to control this process in the framework of its applications for novel propulsion devices and safety issues. In this paper, a series of calculations are carried out in order to distinguish the role of channel geometry (planar or circular) in the flame dynamics at all the stages of acceleration prior to the deflagration-to-detonation transition. The combustion of a stoichiometric acetylene-oxygen mixture diluted with nitrogen is considered. It is established that in channels with a circular cross-section, the flame front is stretched and the combustion process evolves up to the transition to detonation, in accordance with the experiment. At the same time, in planar channels (slits), the primary flame acceleration leads to the establishment of a quasi-stationary mode. The peculiarities of the compression and rarefaction waves evolution in the cylindrical and planar geometry are responsible for such a difference between the flame dynamics in channels of different geometry. The detonation onset takes place via the formation of a "chocked flame" structure characterized by permanent joint compression and acceleration of the combustion. As well as in experiments the deflagration-to-detonation transition occurs exactly at the flame front. • Numerical study of flame acceleration in C 2 H 2 / O 2 / N 2 mixtures using reduced chemistry. • The effect of the channel geometry on the flame acceleration is demonstrated. • Flame dynamics in planar and circular channels is described. • Detailed analysis of deflagration-to-detonation transition is provided. [ABSTRACT FROM AUTHOR]

This paper introduces a novel approach for regulating the pose of a free-flying dual-arm anthropomorphic space manipulator system (SMS) using a finite-time state-dependent Riccati equation (SDRE) controller. The proposed system finds applications in on-orbit satellite inspection, servicing, space structure assembly, and debris manipulation. The dual-arm SMS presented in this work consists of two 7 degrees of freedom (DoF) robotic arms mounted on a free-flying spacecraft, resulting in a complex 20-DoF system. Due to the high number of DoFs, advanced controller design and efficient computations are necessary. The finite-time SDRE controller relies on the state-dependent coefficient (SDC) parameterization matrices, which are nonlinear apparent linearizations of the dynamics. Conventionally, the computation of SDC matrices is offline and relies on the a priori derivation of the analytical equations governing the dynamics of the system. However, this strategy becomes computationally impractical for high DoF plants. To overcome this issue and deliver a more viable solution, a numerical method to construct and update the SDC matrices at each time step is presented. This approach relies on a screw-theory-based recursive Newton–Euler algorithm designed to reconstruct the manipulator inertia and Coriolis matrices. These quantities are the building blocks of the SDC parameters used in the synthesis of the SDRE controller. Simulation results demonstrate the performances of the finite-time SDRE controller augmented with the online update of the state-dependent coefficients. • Automatic derivation of the dynamics of free-flying robot with multiple open chains. • Implementation of a Newton-Euler approach for updating state-dependent coefficients. • Finite-time SDRE to control a 20 DoFs, anthropomorphic dual-arm space manipulator. [ABSTRACT FROM AUTHOR]

This work is devoted to generating optimal guidance commands in real time for attitude-constrained solar sailcrafts in coplanar circular-to-circular interplanetary transfers. Firstly, a nonlinear optimal control problem is established, and the necessary conditions for optimality are derived by Pontryagin's Minimum Principle. Under some mild assumptions, the attitude constraints are rewritten as control constraints, which are replaced by a saturation function so that a parameterized system is formulated. This allows one to generate an optimal trajectory via solving an initial value problem, making it efficient to collect a dataset containing optimal samples, which are essential for training Neural Networks (NNs) to achieve real-time implementation. However, the optimal guidance command may suddenly change from one extreme to another, resulting in discontinuous jumps that generally impair the NN's approximation performance. To address this issue, we use two co-states that the optimal guidance command depends on, to detect discontinuous jumps. A procedure for preprocessing these jumps is then established, thereby ensuring that the preprocessed guidance command remains smooth everywhere. Meanwhile, the sign of one co-state is found to be sufficient to revert the preprocessed guidance command back into the original optimal guidance command. Furthermore, three NNs are built and trained offline, and they cooperate to precisely generate the optimal guidance command in real time. Finally, numerical simulations are presented to demonstrate the developments of the paper. • New dataset generation method for training neural networks used in optimal guidance. • A preprocessing method ensuring the smoothness of the preprocessed guidance command. • Using the sign of one costate to convert the preprocessed guidance command. • Improved fitting performance enabling the precise generation of the guidance command. [ABSTRACT FROM AUTHOR]

This paper presents a methodology for the design of deployable tube flexures made of ultra-thin carbon fiber composite. The proposed process aims to enhance viscoelastic resilience on recovery time after long-term stowage through the design of cut-out geometries. Viscoelasticity is modeled using an experimentally determined Prony series master curve. Bayesian optimization is employed to determine cut-out shapes based on the recovery time and Hashin failure criteria. Partial dependent plots are implemented to interpret the effect of cut-out formation on deployment performance. These reveal optimal profiles of cut-out shapes to achieve a desired operational performance. The efficiency of the proposed methodology is validated by experiment. • Viscoelastic and failure performance of tube flexures is enhanced via design optimization. • Bayesian optimization permits exploration of a highly non-linear design space. • A design framework of tube flexures for desired stowage and deployment is presented. [ABSTRACT FROM AUTHOR]

The Iris mission is a student-led 3-U CubeSat comprised of a science payload of 20 geological samples placed in low-Earth orbit (LEO) to study the effects of space weathering (i.e., those aspects that can be investigated in LEO). This experiment is the first of its kind to take place in the vacuum of space. The science objective of the mission is to investigate how space weathering processes affect the surfaces of airless planetary bodies by monitoring the spectral reflectance properties of geological samples over time. Two three-color cameras acquire spectral reflectance data within the visible spectrum to measure any spectral and/or textural changes exhibited throughout the ∼1-year orbital mission. This mission will provide valuable data to help understand how the space environment may affect the surfaces of the Moon, Mercury, and main-belt and near-Earth asteroids. The acquired data will help constrain meteorite-asteroid links and provide insight into the spectroscopic analysis of samples returned from the Moon, and near-Earth asteroids Itokawa, Bennu, and Ryugu. The results of this mission will also complement the results of laboratory simulations of space weathering. All samples selected for flight on the CubeSat have undergone several tests to ensure suitability for launch and operations. This paper describes the payload subsystem, sample selection and preparation. • First-of-its-kind geological space weathering investigation of lunar and asteroid materials. • The Iris CubeSat geological samples will be directly exposed to the LEO environment. • This study will observe changes in the visible spectral range of the samples over short time scales. • The results of this mission will complement laboratory studies of space weathering. • The CubeSat mission is led by post-secondary students to gain hands-on experience. [ABSTRACT FROM AUTHOR]