The Clean Space industrial days aim at providing an insight to the technological advancement achieved to date in the fields of ecoDesign for space, technologies for Space Debris Mitigation (SDM), Active Debris Removal (ADR) and Space Servicing Vehicles (SSV).
Through its Clean Space initiative, ESA is pioneering an eco-friendly approach to space activities. On the ground, that means adopting greener industrial materials, processes and technologies. In space, it means preserving Earth’s orbital environment as a safe zone, free of debris.
Clean Space is considering the entire lifecycle of space activities, from the early stages of conceptual design to the mission’s end of life – and even beyond, to removal of space debris. Clean Space has three branches, reflecting its mission to assess the environmental impact of Agency programmes as a first step to finding ways to address them in future, and contributing to a more sustainable and competitive European space industry:
ESA has pioneered EcoDesign for space by performing environmental life-cycle assessments of its launchers and missions. Applying LCA to space activities is a first step towards a cleaner space sector. Based on this work, ESA has developed a space-specific handbook to perform LCA. A great challenge is to use the results of the LCAs to develop cleaner technologies which will be at least as performing than existing technologies. The industrial days will be an opportunity for the European space industry to acquire knowledge on ecoDesign for space missions.
ESA is globally leading the pursuit for Active Debris Removal through its e.Deorbit mission. The objective is to catch a large spacecraft in Low Earth Orbit and to bring it back to Earth in a controlled manner, in order to reduce the risk of growth of space debris. Through this mission, European industry is developing cutting edge technologies for monitoring, rendezvous, capture and deorbiting of space systems.
There are clear synergies between the technologies required for the e.Deorbit mission and those needed for the Space Servicing Vechicle missions. Indeed, all of these mission concepts require that the space servicing vehicles perform close proximity operations, including rendezvous and capture. ESA and the European space industry are looking into the technical understanding of these synergies with the objective to pave the way for all future SSV.
The CleanSat project aims to develop technologies to comply with space debris regulations and avoid the generation of new debris from future missions. It seeks to steer this evolution in a coordinated manner, in direct collaboration with Europe’s leading satellite integrators and their suppliers. By developing technologies that enable future satellites to comply with standards, Europe’s space industry will have an international competitive advantage.
By bringing together a variety of stakeholders including the public sector, industry and new potential participants, the Clean Space Industrial Days offer a platform to focus on the development of the required technologies, encouraging all actors to exchange their advancements and difficulties, thus strongly promoting European cooperation.
Mr Franco Ongaro, Director of Technology, Engineering and Quality will launch the 2018 Clean Space Industrial Days
Airbus has been developing a mission and chaser concept for capturing and deorbiting the defunct Envisat satellite with different partners in the frame of ESA studies (e.Deorbit Phases A and B1). In the last e.Deorbit study called ‘Consolidation Phase’ which aims at implementing findings from the intermediate SRR, Airbus along with their partners CBK, MDA, SENER and GMV was pursuing the mission and chaser definition, but then based on the Airbus Spacetug vehicle for GEO servicing.
The assumption for this analysis was a Spacetug vehicle adapted for the specific e.Deorbit mission rather than generic electrical Spacetug for orbit transfer or GEO servicing missions. As the same Eurostar Neo bus platform would be used in both cases, the potential for the reuse of the e.Deorbit Spacetug solution for further servicing missions would be therefore very high.
Starting from mission analyses, an e.Deorbit Spacetug baseline configuration was defined, bringing various options depending on the mission scenarios. As with the Airbus Spacetug vehicle, this e.Deorbit Spacetug vehicle aims at being launched with a medium or heavy launcher. The first considered scenario is the one inherited from e.Deorbit Phase B1 study, i.e. based on motion synchronization with Envisat, capture via robotic arm and fixation and stabilization using clamping mechanism, before deorbiting with controlled re-entry. The Airbus Spacetug preliminary design is based on the Eurostar Neo platform from Airbus and made of a 2.5kW electrical power system, a chemical propulsion system with associated bi-propellant tanks, a set of sensors dedicated to AOCS/GNC, a robotic arm system, gripper and clamping mechanism for servicing and an On-Board Computer and avionics bus reused from Eurostar Neo.
The preliminary assessment of the synergies between the e.Deorbit and GEO Spacetug systems already shows very good opportunities for a mutual fertilization of both systems. A great opportunity pointing out here is the high flexibility of the Spacetug platform to adapt to different mission configurations.
The EC FP7 RemoveDebris mission aims to be one of the world’s first Active Debris Removal (ADR) missions to demonstrate key technologies in-orbit in a cost-e_ective ambitious manner, including: net capture, harpoon capture, vision-based navigation, dragsail de-orbitation. The mission will utilise two CubeSats as artificial debris targets to demonstrate the technologies. In early 2018, the main 100 kg satellite was launched to the International Space Station (ISS) and deployed via the NanoRacks Kaber system into an orbit of around 400 km. The mission comes to an end in early 2019 with all space entities having been de-orbited.
This presentation reports on the LEOP and commissioning phase of the mission, along with initial operational results from the first of the demonstrations with a view to next steps in the operations planning.
This presentation will provide updates on the last results of the GreenSat study currently performed by Thales Alenia Space in France and Deloitte Sustainability. A Life Cycle Assessment (LCA) of the Sentinel 3 mission has been carried out and the results investigated in order to identify the environmental hotspots of the Sentinel 3 mission. Based on these results, brainstorming sessions have been organized in order to identify potential ecodesign options aiming to reduce the environmental footprint of the Sentinel 3 satellite. The different options have been traded-off against a series of objectives/indicators. The presentation will in particular highlight the challenge of rating different ecodesign options based on multicriteria results (environmental performance, technical performance, customer value, etc.). More generally, the presentation will address the relevance of performing ecodesign at mission level and show the necessity of a broader approach to include the impacts of satellite design on the ground segment and the launch segment.
With the GreenSat project, ESA wants to evolve from an assessment of the environmental impact of an existing satellite mission to a significant reduction of the environmental impact through a redesign of this mission. During this presentation, the approach and the first results of this exercise for the PROBA-V mission will be shown.
PROBA (PRoject for On-Board Autonomy) is a family of small satellites developed for ESA by QinetiQ Space. PROBA-V, which is tasked to perform global vegetation monitoring, was launched in May 2013 and is fully operational since then.
The first step in the ecodesign of the PROBA-V mission was the development of an LCA (life cycle assessment) model for this mission. Based on this model, the environmental hotspots were defined.
As a second step, a brainstorm and workshop were organized, involving relevant experts, to identify ecodesign options for the PROBA-V mission. A selection process is followed to select the most promising options. A first step in this selection process included a trade-off analysis, taking into account criteria like potential environmental impact reduction, feasibility, system level impact, expected cost impact. In the next steps a semi-quantitative screening of the environmental reduction potential and possible burden shifting is performed.
The currently running third step is to further elaborate the selected ecodesign options. Each option is technically developed to a higher level of maturity and the environmental impact and benefit are calculated with LCA. Finally, in a fourth step, the environmental impact of the new space mission, including all selected ecodesign options, will be determined and compared to the baseline. The cost, performance, risk, schedule and feasibility of the options will be further evaluated to select the three most promising options. For these options, a roadmap will be developed.
The presentation will explain the "Tech for Space Care" initiative aiming to develop technological elements to ensure the sustainable use of space and the security of space operations in synergy with CleanSpace ESA activities
In a second part will focus on the update of the Overview of CNES activities related to the compliance of the satellites with FSOA in the 2017.
The introduction of Space Debris Mitigation requirements at national, European and international level has to be taken into account in the platform design: as a consequence, Airbus has adapted the end of life strategies for its different LEO platforms in order to comply with the different SDM requirements.
Parallel to this, several technologies aiming to support the platform compliance have been identified. The priority technologies for Airbus will be presented during the Clean Space Industrial Days.
The presentation will outline Thales Alenia Space’s vision to enhance the compliance of its LEO platform productline to the space debris mitigation requirements. A review of these requirements as well as the technical challenges they pose is addressed. The presentation then moves to illustrate existing strategies and ways to build upon these strategies to improve the current compliance status. The enhancement of the compliance status can be achieved through building blocks which are divided in the following categories: deorbit, design for demise, passivation, and reliability. A screening of potentially interesting technologies is done for each category. Following this, a list of high-interest building blocks is proposed together with a mapping between those building blocks and platform classes.
Since the beginning of the space era, a significant amount of debris has progressively been generated in space. In addition, the rise of large low Earth orbit (LEO) constellations mean in the near future, space populations could significantly increase in key orbits. Founded in 2013, ASTROSCALE’s mission is to secure long-term spaceflight safety, and to become a key provider of reliable and cost-efficient spacecraft retrieval services to satellite operators.
The presentation will provide an overview of the ELSA-d end-of-life (EOL) mission, ASTROSCALE’s first semi-cooperative spacecraft retrieval technology and capability demonstration mission, with a targeted launch of late 2019. ELSA-d consists of two spacecraft – a chaser and a target. The chaser is equipped with proximity rendezvous technologies and a magnetic capture mechanism, whereas the target has a docking plate (DP) which enables it to be captured. This is Astroscale’s first step in the maturation of a range of key capabilities which would be present in a full customer mission, including: target search, target inspection, target approach and rendezvous, and target capture.
The presentation will briefly discuss the InnovateUK funded National in-orbit Servicing Ground Segment Facility, developed by Astroscale, hosted at the Satellite Applications Catapult, and considered a key capability for future services. Finally, there will be an examination of how ESA supports ELSA-d through a joint strategic agreement, and can continue to support in future missions.
In Orbit Services is a rising potential market, both in GEO and in LEO orbits, creating baseline technologies for active debris removal (ADR) missions. Life extension, repair and inspection, refueling and transport are a few of the ideas that have been mentioned. The business feasibility is reflected in the new Satellite Servicing Vehicle Concept within ESA‘s Clean Space Initiative.
Effective Space, headquartered in the UK, is designing, deploying and operating the SPACE DRONE™ spacecraft to provide services such as station-keeping and attitude-control, relocation, orbit correction and inclination correction, deorbiting and Bringing-Into-Use (BIU). Its Phase One deployment will provide life-extension services to satellites in GEO. First commercial contract will see two SPACE DRONE™ spacecraft serving two host satellites starting 2020.
The SPACE DRONE™ spacecraft has a launch mass of 400 kg and deploys a highly efficient electric propulsion system. The SPACE DRONE™ spacecraft is designed to be a semi-autonomous, with multiple docking/undocking operations using four docking arms, and to provide full orbit and attitude control for the joint stack. A total mission life time of 15 years is offered by the SPACE DRONE™ spacecraft for a typical 2 ton GEO satellite.
By incorporating some innovating technologies and operational solutions the SPACE DRONE™ spacecraft can service host satellites ranging from 1500kg to 4000kg of dry mass. A non-intrusive (patent pending) docking system was developed allowing the SPACE DRONE™ spacecraft to dock to the launcher adapter ring of the host satellite.
System constraints, such as maintaining station-keeping while handling the SPACE DRONE™ spacecraft and the host satellite shadowing on each other have been incorporated in the design. Joint operations of joint stack have been designed. Method for station-keeping (patent pending) using four electric propulsion thrusters to be mounted on four deployable thruster arms provides all orbit maneuvers with high efficiency.
The SPACE DRONE™ spacecraft can be adapted for providing orbit services in LEO orbit, servicing Post Mission Disposal for LEO satellites and Mega constellations. The SPACE DRONE™ spacecraft can provide orbit disposal and orbit re-location for stranded satellites due to launch errors. With minor changes, the satellite can be also used for Active Debris Removal, utilizing many of the previously developed technologies and lessons learned.
In this presentation the SPACE DRONE™ spacecraft design and system constraints will be presented. Analysis done towards joint stack operations will be shown as well as the different mission concepts that can be performed using the SPACE DRONE™ spacecraft concept.
The announcement of mega-constellations reinforces the debris issue and the need for operators to address the problem of removal of failed spacecraft from orbit. Future satellites shall be compliant with Space Debris Mitigation requirements. In any case, failure before end-of-life could occur, endangering the capability for the satellites to respect those requirements, with increasing collision risk and detrimental effect on operators’ constellation business.
With the number of operating satellites increasing, solutions for mega-constellations have been compared from increasing reliability to in-orbit servicing or removal in the frame of the ESA phase 0 study.
The most cost effective Active Debris Removal solutions have been identified with the associated impact on the constellation business plan.
Today the application of Eco-Design can benefit business, users and society at the same time because it responds to a common interest in obtaining more efficient products in term of both, economic and environmental perspective.
Enhancing the current efforts made by the European space sector dealing with Ecodesign, ArianeGroup is carrying a Life Cycle Analysis of the future European launcher Ariane 6. Since the Launcher System is currently in the design phase, the goal of this study is to evaluate and to anticipate the environmental impact in exploitation phase adopting a proactive approach. With the aim of determining the environmental contribution of the Ariane 6 Launcher System, the major hotspots are identified for a set of relevant environmental indicators.
The scope of the study covers the overall exploitation phase of the A6 Launcher System, including production & assembly, launch campaign/ground operations and launch campaign/flight event. This work focuses on “1 launch of Ariane 64 version” defined as the reference flow. Considering the payload mass put into orbit, a capacity similar to A5 ECA is assumed.
This study is a first attempt for ArianeGroup to assess the overall life cycle of the A6 launcher. Applied methodology, main assumptions, preliminary environmental indicators evaluation as well as conclusions regarding the comparison between A5 ECA and A64 will be presented. A second iteration considering final Launcher System definition and industrial set-up will be completed for the Ariane 6 first flight.
In this way, ArianeGroup is committed to the ESA Clean Space Initiative and contributes to the big challenge for space industry: keeping a competitive advantage for Europe and minimizing the environmental footprint on Earth and Space.
The “LCA Ground Segment” project (ESA Contract No. 4000123991/18/NL/GLC/as) aims to assess the environmental performances and the applicability of eco-design principles to Ground Segment through the elaboration of a specific methodology, the involvement of ground segment experts and the in depth evaluation of the most promising options.
The Life Cycle Assessment (LCA) is a standardized methodology, which assesses environmental impacts associated with all the stages of a product/service life cycle, from raw material extraction and materials processing to manufacture, distribution, use, repair and maintenance and disposal or recycling.
The Ground Segment’s (GS) main functions are the management of the communications between the in-orbit satellite and the mission control centre as well as the archiving and processing of the mission-specific scientific data collected.
The one-year study (started in June 2018), managed by RINA with the support of Spanish Space Company Deimos Space and French Consulting Company Bertin Technologies, is going to:
• Identify and define various “generic families” of GS representatives for Telecommunication (TC), Navigation (NAV), Scientific, Earth Observation (EO), CubeSat missions, covering their specific infrastructures and operations.
• Perform LCA of the environmental impact of the various GS families.
• Provide datasets and methodological guidelines about LCA methodology applied to GS in order to update/complete the ESA LCA Handbook and Database.
• Investigate innovative eco-design options (technical solutions, spin-ins and/or new technologies, innovative processes, etc.) by also considering non-technical aspects (cost and risks, TRL, implementation roadmap, etc.) which can be applied to the various GS family’s infrastructures and operations in order to reduce their environmental impact.
Actually the consortium is completing the analysis and the identification of different GS families, starting with collecting data through a questionnaire from selected GS facilities: ESTRACK Cebreros, Redu and Kiruna; ESOC Darmstadt; HISPASAT Arganda del Rey; ESAC Villanueva; DEIMOS-2, DeSS and LEAF SPACE Puertollano.
The project is part of the ESA Clean Space Initiative, which promotes an eco-friendly approach to space activities.
ARA-Atmospheric Re-entry Assessment is the TASinI-led study aimed at the investigation of the potential impacts on the atmosphere and climate, caused by gases and particles released during the re-entry of spacecrafts and rockets.
The study is divided into two phases:
- Phase 1, dedicated to the definition of the substances released in atmosphere during re-entry
- Phase 2, dedicated to the assessment of the atmospheric impact of released substances
The presentation will provide a specific description of Phase 1 (which had been already carried out), at the same time addressing the main features of potential results from Phase 2.
The increasing number of space debris has become problematic for the sustainment of space activities in Earth orbit. Several spacecraft breakups have been observed in the past and some of them are due to a battery breakup. In order to mitigate the risk to generate debris in the future, passivation of the spacecraft after end of mission is now required. One way to achieve passivation is to deplete the spacecraft battery. However if the latter is to be implemented a very careful assessment of the behaviour of the battery at such conditions should be made, because, how can we be sure that the battery is 100% safe once depleted?
Such an assessment can be very complex considering that temperature, radiation, state of charge and ageing of the battery could play a role among other parameters on the risk and possibility of battery explosion. The actual probability of a thermal runaway leading to a cell explosion is impossible to assess without testing. The ESA TRP objective, performed in collaboration with Airbus DS, SAFT, ABSL and CEA, is to understand and demonstrate through testing the behaviour of a spacecraft lithium-ion battery under extreme conditions which the battery can experience after EoM or during/after passivation if any. More than 200 battery cells used in spacecraft batteries, both fresh and aged, have been tested from an early stage of the project in 2016 and so far in order to assess the impact of ageing during LEO and GEO missions under harsh conditions in space that the battery might experience.
The tests performed in this TRP include External short-circuit, Internal short-circuit, Over-charge, Over-discharge, Accelerating Rate Calorimetry (ARC) test, Over-temperature test and Micrometeoroids. Cells and modules from the two main battery manufacturers in Europe, ABSL and SAFT, are used for testing including ABSL18650-HC, -HCM, -NL and SAFT VES16, VES140, VES180. The test campaign is currently coming to an end along with a series of recommendations for passivation strategies to ensure battery safety.
Since human started to send active satellites on the orbit they needed some energy source and means to store it. Since them there are thousands of active and inactive satellites flying on orbits around the Earth. In recent years most of the spacecrafts send in to space are equipped batteries made of lithium-ion 18650 cells. They have many advantages compare to older designs. Lithium-ion cells are about times more powerful than nickel-metal hydride cell. Their stored energy to mass ration also it is better. Their charge discharge cycle is also remarkable and makes possible operational life of a satellite for years without any significant loss of performance. With all-of those good things Lithium-ion cells come with one big trade-off: hazard of explosion. Batteries which are subjected to elevated temperature might go into thermal runaway where the cell starts to rapidly get fire and rupture. If such battery will explode it might cause to fragment its or satellite structure. To avoid this scenario European Space Agency from 2020 will implement new space mitigation requirement in which there will be set of rules regarding battery Passivation (deplete energy of batteries at the end of life and additional contamination of them from thermal runaway). Up to this day explosive forces and nature of lithium-ion cells is not well explored. Jakusz SpaceTech with ABSL are preparing research having goal is to understand what are explosive forces of the batteries. Test specimen will be single cells, 8-pack of cells, and 88 block cell with and without walls. In Batteries will be placed inside detonation chamber on heated table to cause thermal runaway. Pressures generated in thermal runaways will be compared to pressures generated by conventional explosive materials. Based on this data there will calculated TNT equivalent (which serves as unit of explosive force) for lithium-ion cells. future this will information will be key element for designing contamination device. Next company will plan to make tests in vacuum with witness sheet to have a reference for future tests especially in future with contamination device. If under thermal runaway all contents will be contaminated without damaging witness sheets device will be declared as serving it goal.
We have established and applied the software tool “PHILOS-SOPHIA” that enables a non-expert user to perform hydrocode simulations of hypervelocity collisions on orbit. It relies on high-fidelity hydrocode simulations, which is the best method for systematically studying the processes and effects of orbit fragmentations for a wide range of impact conditions and for complex space systems. Other numerical methods such as empirical tools or analytical models are limited to a narrow validity range. Their computing time is much faster due to their simplifying nature, but they cannot meet the quality and precision of physics-based methods, nor do they allow complex modelling.
“PHILO-SOPHIA” includes a graphical user interface that can be used to 1) setup complex collision scenarios, 2) start the simulations using the existing FE/SPH solver SOPHIA, 3) monitor its outputs, and 4) visualize and analyze the results. The software tool outputs the characteristics of the generated fragments, including mass, size and the velocity vector of each fragment.
We performed comprehensive numerical simulations using “PHILOS-SOPHIA”. This includes the simulation of 1) spacecraft shielding performance on component level, 2) simple validation with experimental results, and 3) complex collisions using ESA LOFT spacecraft. The spacecraft shielding analysis demonstrated the tool capability for studying specific hypervelocity features but also showed the demand for adequate material models when new materials like Kevlar and CFRP structures come into play.
The validation against experimental data and a commercial general-purpose hydrocode showed very good performance of PHILO-SOPHIA in terms of quality and computing time. Since the existing experimental data (high-speed images) allow for a more qualitative comparison of the fragment cloud size and its evolution in time and space, we propose to perform new experiments using advanced particle methods to gain quantitative data for validation.
We performed detailed fragmentation analyses for different complex scenarios using the LOFT spacecraft. When comparing the results with the empirical NASA Standard Satellite Breakup Model, we found both good agreements and clear deviations. Due to the strong influence of the collision geometry, we did not find a strongly noticeable breakup limit depending only on the energy-to-mass ratio. More research is recommended to define generalized criteria for catastrophic collision conditions. PHILOS-SOPHIA, thoroughly backed by advanced experiments, can be the tool for this purpose.
Germanium is the semiconductor of choice for the production of high-efficient multi-junction space solar cells. Solar cells technology is, by nature, a large surface area semiconductor application and therefore Germanium is the most important semiconductor material, in weight, of all Space missions. Ge has been identified as one of the important environmental hotspots of space missions. This presentation will outline the ambitions of Umicore Electro Optic Materials to reduce the environmental impact of germanium.
Primary Germanium is extracted from coal ash or as a by-product of base metals refining processes. Depending on the source material, the carbon footprint of this Germanium extraction process varies widely. The presentation will cover the entire scope of our approach from LCA study to sustainable supply of Ge to recycling of Ge in internal and external processes and increasing the efficiency of Ge use in the substrate and solar cell manufacturing processes. We will cover a description of the current status and a visionary outlook for the years to come.
A screening program to test different forms of weld (EB, WAAM, FSW, OTIG, RFW) with a variety of materials (aluminium alloys, titanium alloys, stainless steel) against green propellants LMP-103S and Hydrogen Peroxide (HTP) has been undertaken to ascertain the compatibility and usability of the weld/material combinations for future propulsion systems. The results from the accelerated compatibility testing, including leaching of materials, decomposition of propellants and materials properties tests are presented, with recommendations for further research into the most promising combinations.
Polyurethanes (PUs) are versatile materials applicable across many industries namely for their excellent resilience and applicability in different forms: flexible and rigid; monolith and foam. In space industry they are used for instance in spacecraft as coating and potting materials for protection of electronic compounds and further in launchers as rigid foams for thermal insulation of cryogenic tanks for liquid propellants.
Results of a screening study (1) are confirming a great potential to replace the conventional PUs in the selected space applications by the newly developed eco-friendly hybrid non-isocyanate polyurethane materials (HNIPUs). The HNIPU materials are formulated with the aim to minimize health and ecological issues related to the use of toxic isocyanates that are main chemical compounds in production of traditional PUs. Sustainability aspect is also achieved using renewable resources, such as vegetable oils and/or their derivatives.
Our recent research in the framework of this project confirmed possibility to prepare non-isocyanate conformal coating and potting systems with bio-sourced mass content exceeding 50 wt. % and content of non-isocyanate-based hydroxy urethane bonds mass per total bond mass up to 100 %. Processing, thermo-mechanical and electrical properties are adjustable to meet requirements for industrial applications.
The currently developed HNIPU foam has been compared to the reference PU system CRE210VS (provided by Latvian State Institute of Wood Chemistry) and displays ca. 3.5 times higher density (0.163 g/cm3), about 10 times higher compression strength at 10 % deformation (1.85 MPa @ 25 °C) and almost 2 times higher conductivity (0.042 W/m·K @20 °C). Enhanced mechanical properties and increased thermal conductivity are related to higher density of the developed foam.
Scale-up of the HNIPU foam production by industrial PU spraying equipment as the next step of the study is feasible by optimization of foaming parameters with the aim to reach requirements of industrial processing and finer, closed-cell structure.
The developed HNIPU is a good candidate for further development of non-isocyanate based conformal coatings, potting and thermo-insulation PU foam materials applicable in space industry.
(1) ESA Contract No. 4000119685/17/NL/KML: „Development of „Green“ Polyurethane Materials for Use in Spacecraft and Launcher Applications“.
When looking at the complex value chain of space systems, the obsolescence risk due to the use of critical raw materials (CRM) is a critical issue for the space industry and needs to be investigated and anticipated.
The European space sector needs to find ways to anticipate obsolescence risks related to CRM use and develop technological solutions that minimize those risks as early as possible in the design of future space systems.
ESA has adopted the eco-design approach to design future space missions in a more environmentally friendly way: eco-design is a preventive approach to mitigate the environmental impacts of a product (good or service) as early as possible in the design phase. To do so, ESA successfully applied Life Cycle Assessment (LCA) to space activities. Life Cycle Assessment (LCA) is a powerful method, standardized at international level by ISO, to evaluate the potential environmental impacts of products and services in a comprehensive and objective manner and from a multi-criteria life cycle perspective.
ESA deemed these obsolescence risks as a fully-fledged dimension of its eco-design framework towards more sustainable space activities. This is why ESA commissioned a consortium led by Deloitte Sustainability to develop an LCA-related methodology to identify, flag and classify the obsolescence risks due to CRM use through the complete life-cycle of space products.
The project’s first step consists of a literature review and interviews aiming to identify existing similar initiatives in other sectors. Secondly, the project team has calculated the main raw materials’ criticality for the European space sector. Thirdly, a systematical methodology was developed to map the CRM obsolescence risks to the use of certain space components/materials in the design of a space mission. This methodology now needs to be tested on several test cases to check its applicability.
In this presentation, we will recall the context and challenges related to the use of Critical Raw Materials in the European Space Sector. Then, we will present the methodology used to recalculate the criticality of main raw materials and the factors that were specifically adapted to the case of the European space industry. Finally, we will present the methodology proposed to map the obsolescence risks related to CRM use in the early design of a space mission and discuss about the next steps of the project.
The REACH requirements affect the European space sector to a great extent, both from a regulatory compliance and commercial perspective. The processes for Registration and in particular Authorisation of Substances of Very High Concern (SVHC), which aim at their substitution with suitable alternatives, pose continuous risks that have to be actively monitored and mitigated by the space industry to secure the reliable continuation of space activities and the EU’s independent access to space as a key element of the EU’s space policy
In the context of REACH, an obsolescence risk in the space industry can be defined as any possibility of impairment of quality and reliability or even loss of critical technologies for qualified materials and processes, which is induced by a chemical’s unavailability or substitution threat.
The European space sector has launched significant activities to identify and anticipate the REACH-related obsolescence risk for several years. The European space sector needs to make a further step forward and find ways to anticipate REACH-related obsolescence risks and develop technological solutions that minimize those risks as early as possible in the design of future space systems.
ESA has adopted the eco-design approach to design future space missions in a more environmentally friendly way: eco-design is a preventive approach to mitigate the environmental impacts of a product (good or service) as early as possible in the design phase. To do so, ESA successfully applied Life Cycle Assessment (LCA) to space activities. Life Cycle Assessment (LCA) is a powerful method, standardized at international level by ISO, to evaluate the potential environmental impacts of products and services in a comprehensive and objective manner and from a multi-criteria life cycle perspective.
ESA deemed these obsolescence risks as a fully-fledged dimension of its eco-design framework towards more sustainable space activities. This is why ESA commissioned a consortium led by Deloitte Sustainability and REACHLaw to develop a LCA-related methodology to identify, flag and classify the obsolescence risks due to REACH through the complete life-cycle of space products.
The project’s first step has consisted in a literature review and interviews aiming to identify existing similar initiatives in other sectors. On this basis, the project team developed a systematical methodology to map the REACH-related obsolescence risks to the use of certain space components/materials in the design of a space mission. This methodology now needs to be tested on several test cases to check its applicability.
In this presentation, we will recall the objectives of the project, present the key principles of the methodology developed in the frame of this project and discuss about the next steps and possible implications and uses for the European space industry.
Our idea addresses the two biggest problems of space activities: their extreme costs as well as space debris reduction. Beside comments and feedback from the audience, we were looking for research & industry partners who would be interested in working with us on this concept.
Space debris is a serious issue. Although only few humans reached out into space, humanity polluted the orbit already with thousands of tons of waste. The current approach of voluntarily deorbiting space debris within 25 years or by moving it to graveyard orbits seems to be not adequate anymore in the era of satellite mega constellations. One day in the future, even graveyard orbits will be filled up like waste dumps today. And beside negative ecological aspects, the exponential waste growth put any future space mission at risks of unplannable collisions.
Like waste on Earth, space debris should not be considered only as a burden but as an opportunity, too. The recycling industry in Europe has become a multi-billion Euro industry in the past few decades. The concept could be applied to space, too. It’s only important to understand, what the value of space debris would be and what commercial use case lies beneath it: From our perspective, the core value of space debris lies in its location in space as it doesn’t require to be launched into the orbit again. A simplified calculation clarifies this advantage: A single Ariane 5 launch transports around 10 tons of material into orbit. With 8,000 tons of space debris, more than 800 Ariane launches would be needed. With estimated costs of 100 Million Euro each, this would result into 80 billion Euro launch costs alone to take an equivalent weight into space (ignoring the additional costs for this material).
Unfortunately, not all space debris material could be reused / recycled with affordable efforts: either the pieces are far too small to be collected, or the material cannot be reused (like fuel droplets), or the objects consist of complex material mixtures or individual shapes and forms like many scientific satellites which makes automated recycling processes a nightmare.
Luckily, there are “perfect” objects which seems to be ideal for recycling: rocket upper stages. Rocket upper stages have simple, geometric shapes (without large solar panels or antennas), have dedicated, well described connectors (where the payload was attached) and consists mainly of fuel tanks. Especially the European ESC-A upper stage from the Ariane launcher consists of 4 tons of aluminum for its two LH2 and LOx tanks. More than 60 of them are circulating in our orbits for the next decades or even centuries. Simplified, their shape is a geometric cylinder with a connector for the lower payload adapter and / or the Sylda on the top. By using currently developed technology described below, upper stages could be caught in space (e.g. with a net) and transported to the Moon (with a “space tug”). On the Moon, either the aluminum could be recycled and used as new construction material for the planned Moon station or the upper stages could be reused for the same purpose they had been designed for initially: as fuel tanks. Ice (water) on the Moon could be split into hydrogen and oxygen and used in fuel-cells to power the Moon station. To store the hydrogen and oxygen, the tanks of the ESC-A upper stages are ideally suited. Each ESC-A stage could store 14 tons of liquid hydrogen and oxygen in the right ratio. With the identified 60+ upper stages, the whole storage capacity would be 850 tons resulting in approx. 5500 MWh energy supply, enough for the largest Moon station.
Our concept could be applied to upper stages from different types and sizes to accommodate the needs in different scenarios like larger tanks for space / planetary stations or smaller ones for asteroid mining facilities in the future. Our proposal helps to reduce construction costs wherever water (ice) will be processed in space.
Our concept is based on special space tugs: space ships capable of catching an upper stage e.g. with a net and moving it afterwards to the Moon. The space tug concept would be based on the following ESA projects:
Our proposed concept is an additional use case for all of these projects and complements the activities. Over time, technology will improve and affordable metal recycling might become feasible in space, too. This would allow to recycle the 4 tons of aluminum of the ESC-A upper stage to construct new aluminum elements. Aluminum could be reused as foils, foam, structures or powder for 3D printing for all kind of construction purposes. By using material from space instead of from the Earth’s surface, similar launch cost savings like in the tank scenario above would be possible. With 250 tons of aluminum already in orbit today, commercial recycling would be feasible if the commitment of space agencies to use recycled material in their space construction would be made. This would allow the space industry to purchase the recycled material and reuse it in space at lower costs than comparable Earth material.
While the removal targets should be selected from a global perspective, legal constraints dealing with the ownership of space debris objects, and the validation thereof, cannot be neglected. In the case of Ariane ESC-A upper stages, the ownership is clear (Ariane Group with France as the corresponding country). With strong relationships, especially in the development of the Ariane launchers, combined ADR activities between ESA, France and the Ariane Group should be more than realistic, including all legal aspects.
ESA has performed feasibility analysis on a proposed baseline GNC system during a CDF study to design a Space Servicing Vehicle. The main objective of the GNC design is to be able to serve multiple missions with minor adaptations. This is achieved by defining a common rendezvous and capture/docking strategy and sensor suite. Based on ATV experience and previous studies a common framework for the definition of the GNC functionalities is proposed, being the main difference between missions the distances to the target. The significant differences between cooperative and uncooperative rendezvous and/or the mission requirements are absorbed by close proximity navigation and guidance algorithms. It worth noting that the proposed baseline reuse most all the algorithms and only different configuration parameters are needed to adapt from one mission to another. High-fidelity simulation of the most important building blocks (image processing, on-board navigation) has being used to confirm preliminary feasibility for the scenarios considered in the CDF. Some examples are camera-based LOS-only relative navigation outside the Keep-Out Sphere, or camera-based relative pose estimation navigation using control-points in the outer envelope of the target satellite.
cosine Measurement Systems BV is developing visual and infrared optical systems ranging from 0.4 to 14 µm suitable for relative navigation, on-orbit servicing and debris removal. We report on the development status including ongoing and planned In-Orbit Demonstrations of the technology, and tailored developments of three vision-based sensor systems for the I3DS suite: a stereo camera system in the VNIR range, itself comprised of two camera heads; a high resolution VNIR camera; and a thermal infrared camera in the 8 µm to 14 µm range.
The RVS® rendezvous & docking sensor technology has successfully served as operative LIDAR approach sensor on eigteen re-supply vehicle missions to the International Space Station (ISS). Drawing from the experience with this robust and reliable technology, the next generation time-of-flight scanning LIDAR sensors, RVS3000, has been developed and qualified by Jena-Optronik with the support of ESA and DLR – German Space Agency. The development successfully achieved:
The RVS3000 sensor family has been tested thoroughly, including an on-orbit demonstration flight of the optical head on ATV-5 in February 2015 and two campaigns with the complete sensor on robotic test ranges of DLR (EPOS) in December 2017 and of NASA in June 2018. In order to achieve the mentioned results, a number of new technologies have been incorporated:
Today, RVS3000 is a leading edge rendezvous & docking sensor suitable for various applications, including ISS approach, on-orbit servicing, planetary landing, and exploration. The convincing technology and the competitive pricing have allowed to contract already nineteen flight models for both application cases, cooperative and non-cooperative targets.
Extending life or repairing damaged on-orbit assets is not only a very attractive economic option for satellite operators as it could potentially increase margins for commercial services or increasing delivered value of scientific missions, but it would also help reducing the number of debris objects in space.
These types of servicing missions pose technical challenges never faced until now. Of utmost relevance is the autonomous control of several movable devices, whose dynamics are inter-coupled (e.g., spacecraft platform, robotic manipulator, and end-effector), needed to safely and effectively achieve the mission objective.
In the frame of ESA-supported COMRADE study, fully combined control (single control system controlling simultaneously all movable devices) is proposed due to its higher improvement potential (propellant saving, performances increase, safety) w.r.t. tele-operation, decoupled and/or collaborative control (the last one characterized by the use of two different control systems for the spacecraft platform and robotic manipulator respectively but, differently to the decoupled version, with information/feedback about what the other control system intends to do). Two independent combined control designs are developed in COMRADE (H∞ and nonlinear Lyapunov-based), and tested. Each of them is applied for both Active Debris Removal (ADR) and servicing/re-fuelling mission scenarios.
The : the processes of scenario analysis and derivation of COMRADE system requirements; a description of the design and setup for a Simulator, which included at its core the selection, prototyping and integration of algorithms for Guidance, Navigation and Control (GNC), Modes Management (AMM) and Failures Detection, Isolation and Recovery (FDIR) (all three together compose the COMRADE system) and the outcomes of the simulation phase of the Verification & Validation process.
Life cycle assessment (LCA) is the systematic estimation and evaluation of environmental impacts over the life cycle of technical systems and provides the framework and data to support eco-design. ESA, as part of their eco-design initiative, in the last few years has initiated several projects for LCA data regarding materials, components, manufacturing processes, space propellants and missions. The Agency is also making LCA as part of development contracts, providing a source for expanding the LCA library for space systems. In this presentation, we describe the opportunities and challenges for the space industry from an aligned LCA database for space systems, with discussions based on an ongoing ESA contract to review, evaluate and harmonize existing ESA LCA data.
Prior to harmonization, the ESA database contained ~ 1500 individual LCA datasets. This will represent a vital library of environmental data for eco-design, and includes subsystems, materials, components and processes. However, the management and use of the database brings challenges regarding confidentiality of industry data and requires quality criteria for existing datasets and for additions to the database.
The presentation will show an overview of the LCA database and illustrate how these are used to model space technology. Moreover, we discuss major gaps identified in the finalized database, and the formats to handle confidentiality and continuous expansion of new data through ESA contracts. The space industry’s perspective on these discussions is greatly welcome
In 2017 the United Nations Committee on the Peaceful Use of Outer Space (COPUOS) reached a consensus on the first set of sustainability guidelines for the space sector named the ‘Guidelines for the long-term sustainability of outer space activities’. These guidelines perfectly aligns with the mission of the European Space Agency’s Clean Space Initiative by pioneering an eco-friendly approach to space activities. In this regard, the Ecodesign branch of the Clean Space Initiative has identified Environmental Life Cycle Assessment (E-LCA) as being the most appropriate tool to measure the environmental impacts of the entire life cycle of a space mission.
However, Guideline 27.3 of the COPUOS sustainability guidelines state that in the conduct of their space activities, actors should take into account “the social, economic and environmental dimensions of sustainable development on Earth”. This suggest the need for a more encompassing form of sustainability assessment to address the three pillars of sustainability by incorporating a multitude of important environmental, social and economic aspects of space missions into the assessment. This notion directly aligns with the development of E-LCA over recent years and the growing consensus within the environmental sector for a move towards Life Cycle Sustainability Assessment (LCSA). Despite this, until recently, no LCSA for space had ever been attempted.
To address this, the Strathclyde Space Systems Database (SSSD) was built to become the first LCSA database for the space sector to allow the industry to become more accountable and responsible for their operations by taking into account the full spectrum of life cycle impacts and sustainability issues associated with the operation of space systems. Since it’s unveiling at the 68th International Astronautical Congress, the SSSD has considerably grown and has now run a variety of post-mission studies ahead of its full integration into concurrent design in February 2019.
This paper will present a status update on the development of the SSSD as well as providing LCSA results of the MÌOS mission which was run through the SSSD. The results will be based on two scenarios; (1) full system level results, and (2) the impact of switching propellant types on the full system level results. The analysis shows the importance and practicality of all three sustainability dimensions in assisting decision-makers to select the most sustainable technologies and products through the play-off between cost-eﬃciency, eco-eﬃciency and social responsibility.
The CleanSpace One mission aims at de-orbiting EPFL’s CubeSat SwissCube. The motivation behind the CleanSpace One project is to increase international awareness and start mitigating the impact on the space environment by acting responsibly and removing our “debris” from orbit. SwissCube is on a polar type orbit at 710 km altitude, 98.3-degree inclination. Launch is planned around 2023.
To be in line with the project’s sustainability intentions, the project started two new threads in its development: demisability investigations and eco-design. The goal would be, according to the funding available, to integrate as much as possible of these threads as early as possible in the project, as a test platform. The eco-design thread has the purpose of creating the frame for gathering the right data all along the project to be able to perform a Life Cycle Assessment. Thus it takes into account the impacts of each material and each process related to design, create, test, use and dispose the product.
The work started via a student part-time project, at a time where the ESA database was not yet available. The implementation consists of an excel parametrisation and an attempt was done to insert the excel tool into Cameo, a SYSML model based system engineering tool. The tool includes flows created by EPFL’s Sustainable Campus for the transport, infrastructure and electricity. This paper will shortly discuss the progress made with the tool.
The “Space system Life Cycle Assessment (LCA) guidelines” (or ESA LCA Handbook) was developed by ESA with the goal of establishing methodological rules for performing space-specific LCAs. General rules for LCA are set in the ISO 14040 and 14044 standards. However, these standards are not sector specific and leave many options open for the LCA practitioner to decide. To obtain comparable results, further guidance is thus needed. ESA also developed a LCI (life cycle inventory) database with space specific materials and processes.
During the ecodesign of the PROBA-V mission in the GreenSat project, the ESA LCA handbook and database are applied. This presentation will discuss the advantages and added value of both tools when performing space system LCA studies. The presentation will also touch upon possibilities for improvement of the ESA LCA Handbook and the ESA database, using concrete examples encountered while performing the PROBA-V life cycle assessment.
Recent studies have highlighted reaction wheels as critical elements which are likely to survive re-entry and contribute to casualty risk. The major element of concern is the flywheel which is often constructed of stainless steel in larger reaction wheels, although the survival of the steel ball bearing unit due to its shielded location is also an issue. Baseline simulations with the SAM tool using a seven-component compound model are consistent with these studies.
Wind tunnel tests have been performed on an engineering model of a 120mm diameter reaction wheel with a steel flywheel and ball bearing unit. An initial test was performed at low heat flux and the failure of the aluminium housing was observed. Running at higher heat flux on the surviving steel flywheel and ball bearing unit parts shows a heat flux profile which is dependent on the local heating. The parts were tested at heat fluxes higher than would be expected in re-entry and demise was not observed, supporting the suggestion that these objects can be a casualty risk. Rebuilds of the heat fluxes on the surfaces have been performed, and an energy balance using data from the demisable materials studies has shown consistent results.
A wide variety of potential design-for-demise techniques have been simulated using SAM to assess the reduction in the casualty area produced. Each assessment has been performed using 1000 runs accounting for uncertainties in the release conditions and the aerothermodynamic heating. Improved demisability is most clearly seen with changes to the flywheel material. Copper is identified as a viable alternative with demise performance to close to that of aluminium but the capability to maintain the flywheel inertia due to its higher density.
Finally, a small benefit from a spoked flywheel is seen using the inclination-based Modified Lees approach for the aerothermal heating. However, if the local length-scale SAM heating model is used, which is much more in line with the test observations, significantly higher heat fluxes are seen on the spoked model. A reduction of under 10% in the casualty area is seen for the standard wheel, but more than 20% for the spoked wheel. In this case the casualty area is still driven by the steel flywheel, and the demise probability of the ball bearing unit is almost doubled. Allied to a change in the flywheel material, this design shows potential.
Design-for-demise (D4D) looks at technical solutions to reduce the casualty risk on ground of re-entering satellites and their components by promoting demise during atmospheric re-entry. Earlier studies have shown that the early release of the satellite structure will also help to improve the overall demise of the satellite. In order to understand betterbetter understand the behaviour of current joining technologies and how to improve them for demise, tests have been performed in a high enthalpy wind tunnel and static heat chambers. A broad range of samples were prepared and tested under a range of conditions in order to broaden the current understanding. A broadwide range of phenomena was exhibited by the samples and a number of different failure scenarios were seen to be dependent upon technologies tested, heat flux profiles, and the mechanical loads, among other influencing factors. These results are currently being fed into developments of new demisable joining technologies which will undergo bread boarding development and testing similar to the earlier phase. These activities as a whole both help us to understand the processes of demise and also will lead to more informed decisions in terms of increasing the altitude at which break-up occurs and reducing on-ground casualty risk.
In the e.Deorbit Consolidation Phase CBK PAN is responsible for development and validation of a cartesian force/torque compliant control of the robotic arm. This controller will be used during the clamping operation. The goal of the compliant control is to ensure that during the motion of the arm the forces and torques (especially acting on the gripper) will be kept within a certain limit and that the gripper, the robotic arm and the clamp will not be damaged. In this case, it is necessary to control not only the position of the end-effector, but also the forces that arise when the manipulator interacts with the environment. The approache selected to solve this problem is the impedance control approach which is a type of the compliant control. The target impedance is defined as a mass-spring-damper system and this approach allows selection of such trajectory which ensure that the target impedance is achieved.
Numerical simulations were performed to verify the proposed compliant control. These simulations were carried out using the ‘Simulation tool for space robotics’, which is being developed at CBK since 2009. The tool is based on the SimMechanics model of chaser satellite equipped with a 7 DOF robotic arm (SimMechanics is software based on the Simscape, the platform product for the MATLAB Simulink). The model of Envisat is attached to the gripper. For the purpose of compliant control simulations in the frame of the e.Deorbit mission the simulation tool was updated. The first modification is related to the robotic arm elasticity. In space elasticity is an important issue due to the lightweight structure of the arm. Preliminary simulations (based on mathematical models) showed that for the e.Deorbit robotic arm the influence of elasticity in manipulator joints on the motion of the end-effector is much higher than the influence of elasticity of manipulator links. Therefore, in the compliant control simulations only the elasticity of manipulator joints is considered. Another new feature, added to the simulation tool, is the model of contact. The contact between the clamp and LAR, occurring during the clamping operation, has significant influence on the behavior of the system and must be taken into account.
Validation of the proposed compliant control was performed using the Monte Carlo method. There were three main objectives of these analyses : (i) to test the proposed compliant control for a wide range of parameters and initial conditions, (ii) to estimate forces and torques induced by the motion of the robotic arm when the gripper is attached to LAR, and (iii) to identify parameters that have the most significant influence on the overall performance of the compliant control.
Nowadays, robotic spacecraft on orbit servicing or so called “Space Tug” capabilities are limited. This is mainly due to the relative poor availability of exteroceptive sensors for space navigation and the poor on-board processing resources preventing the design of ambitious autonomous systems.
Goal of the I3DS project is to fulfil this technology gap and realise a suite of perception sensors that will allow localisation and map-making for robotic inspection of orbital assets.
I3DS is a generic and modular system answering the needs of near-future space exploration missions in terms of remote and contact sensors with integrated pre-processing and data concentration functions. It consists in state-of-the art sensors and illumination devices integrated in a coherent architecture as inter-changeable building blocks and targeting an on-orbit missions such as non-cooperative target capture such as debris removal missions and cooperative rendezvous: servicing & Space Tugs.
I3DS Integrated 3D Sensors is a project co-funded under Horizon 2020 EU research and development program and part of the Strategic Research Cluster on Space Robotics Technologies as the Operational Grant n°4 among 6.
An important application is seen in the collection and avoidance of space debris, up to active de-orbiting of LEO satellites which have failed or reached the end of their lifetime.
For active de-orbiting of unusable satellites, safe and reliable capture technologies have to be designed and developed. Clamping technologies are one of the building blocks of ADR missions, and an interesting development for in-orbiting servicing in general. The e.Deorbit particular demanding case (large object, etc.) can indeed serve as a good demonstrator of the capabilities developed and the lessons learned can be transferred to other ADR activities.
A Clamping Mechanism (CLM) technology is currently being developed under an ESA contract in the framework of the e.Deorbit mission within ESA CleanSpace activities.
The objectives of the activity are:
• To design a clamping mechanism (CLM) for clamping a chaser spacecraft to an interface on a Launch Adapter Ring (LAR) of Envisat spacecraft during a debris removal mission
• Manufacture & test a breadboard of the CLM to reach TRL 4 of the CLM technology by the end of the activity
The CLM presents several features which make the project a very interesting example of what can be accomplished for future ADR missions:
• Large size of the target debris object, with its about 10 m and 7800 kg, which translates to corresponding large efforts and demanding stiffness to cope with the about 1.7 kN de-orbiting thrusts (and corresponding bending torques)
• The robotic arm grasps the LAR and approximates the CLM to its nominal position within certain accuracy, which results on potential misalignments of ±25 mm and ±2.5º, which results in a relatively wide region to be covered by the CLM closing motion and requires specific means to adapt to the potential variations
• The CLM has to compensate these misalignments against the backdriving reaction efforts of the robotic arm, up to 20 N and 80 N.m
• Actual LAR status is unknown, so the CLM has to be able to deal with manufacturing tolerances, presence of (potentially degraded) thermal tape, micrometeoroid-caused damage and large thermal gradients (which, for the hot scenario of +140ºC limit significantly the aluminum LAR structural strength)
• Once clamped, the CLM has to provide stack alignment capability up to ±20º to compensate for the uncertainties on the position of the target center of mass. This realignment capability has to be also available between de-orbiting thrusts
The presentation will address the status and main challenges overcome during the activity.
In the course of the e.Deorbit Consolidation Phase study led by Airbus Defence & Space, SENER is responsible for the development of the clamping mechanism aimed to clamp and lock on the launch adapter ring of ENVISAT – the inactive ESA satellite. Afterwards, the mechanism has to sustain the main loads induced during the de-orbit phase.
One of the main goals of this phase was to select and improve the final design of the clamping mechanism based on the two concepts analysed in the previous stages.
In general, the mechanism consists of the following sub-assemblies:
The current e.Deorbit concept is represented by the rotary clamps – their main function is to hold on the Launch Adapter Ring (LAR) of ENVISAT. In addition, the alignment mechanism provides capability of changing angular position of the clamping mechanism in order to point Space Tug’s thrust vector onto ENVISAT’s CoG prior disposal phase performance. The main design drivers include: loads transfer capability, stiffness, interface and its condition, redundancy, sensors and control system.
The space sector is a new area of development for Life Cycle Assessment studies. Considering guidelines for the evaluation of environmental impacts of space activities, several actors of or related to the European space industry, such as ArianeGroup and the European Space Agency, have identified Life Cycle Assessment (according to ISO14040/44) as the most appropriate methodology to measure and minimise their environmental impact. To demonstrate the value of life cycle thinking at complete space mission level and more particularly during the design stage, a Life Cycle Assessment (LCA) study of Ariane 6 in exploitation phase is currently performed by the Design for Environment team of ArianeGroup.
Current LCA studies adopt a Cradle-to-Launch pad approach due to lack of modeling for use and disposal phase. Therefore, to cover End-of-Life (EoL) aspects, a priority has been given by the ArianeGroup to the integration of space debris into the LCA framework. LCA studies of space missions should indicate trade-offs not only between typical impact categories, e.g. toxicity and climate change but also with regard to the potential impact of space debris on the orbital environment as a particular challenge for the sustainable use of the orbital resource.
Given this context, a specific LCA indicator is currently being developed and will be discussed against work previously proposed and published by Space debris experts. This indicator called orbital scarcity is fully in compliance with the LCA framework and particularly the ISO standard (14040/44). The presentation will highlight the relevance of considering on-orbit stages of space systems in LCA as a way of measuring the resource security for orbits.
A case study will be presented to demonstrate the applicability of such methodology. The assessment of the potential trade-off occurring in term of environmental impacts on both the Earth and orbital environment will be discussed, comparing several Post-Mission Disposa (PMD) scenarios. Active and Passive deorbiting options that occur during this stage should be discussed through the prism of LCA environmental evaluation.
In this way, ArianeGroup and its partners are committed to the big challenge for the European space industry: keeping a competitive advantage for Europe while decreasing the environmental footprint on Earth and Space.
The evaluation of the environmental impact of a mission can be extended to consider not only aspects such as resources depletion and toxicity, but also its impact on the space environment and, in particular, the potential contribution of a mission to the creation of new space debris objects. The approach described in this work is based on a risk indicator that quantifies which is the probability of an object to be involved in a fragmentation and which would be the severity of such an event, considering the resulting increase in the collision probability for operational satellites. The index is computed along the whole mission profile in order to take into account whether disposal strategies are implemented and the reliability of such strategies.
A self-standing index, in the frame of life cycle assessment or with a specific use such as ranking ADR targets, can only serve as an actionable metric if it can be placed in the broader context. This includes the space environment of a whole as well as shifting from single object to mission such as constellations, including launchers. As the index proposed here relies on high-level mission parameters (e.g. operational orbit, mass, disposal strategy), it can be evaluated early in the mission design phase and it allows, for example, different design choices (e.g. operating altitude) to be compared. This leads to using the index as a metric of how the environment capacity is consumed, indicating with this term the set of space objects and missions that are compatible with a stable evolution of the environment. It will be shown how long term simulations of the environment can be used to define a threshold value of the cumulative index across the whole environment. Finally, a capacity allocation scheme, similar to the frequency one regulated by ITU, will be outlined and examined considering the actual launch traffic and compliance to mitigation guidelines observed in the last years.
Space debris is today a constant threat to all space activities. According to recent requirements and guidelines, operators of space missions have to ensure that space vehicles do not become space debris at the end of their mission. It becomes compelled to design all new space missions considering End of Life requirements in order to ensure a sustainable use of space orbits.
In addition to Space Debris Mitigation good practices, several actors of, or related to the European space industry have identified Life Cycle Assessment (ISO14040/44) as the most appropriate methodology to measure and minimize their environmental impact. However, the space sector deals with strong particularities which complicate the use of LCA. Strong efforts still have to be done in order to fully characterize the complete life cycle of space systems, particularly the on-obit stages which include the interactions with the orbital environment.
Given this context, a priority has been given to the integration of space debris related impacts on the orbital environment into the Life Cycle Impact Assessment (LCIA) framework. To address space debris issues in a comprehensive way, a causal chain (i.e. Impact pathway) linking orbital occupation to environmental mechanisms (midpoint) and damages (endpoint) has been already defined considering the near-Earth orbits as a resource. Consequently, the Endpoint characterization shall be performed considering socio-economic damages into the orbital resource asset caused by the potential emission of space debris.(1,2)
A preliminary work dealing with the valuation of orbitals regions has been achieved focusing on the revenue generated by satellites into the Sun Synchronous Orbital (SSO) region (3)(as a proxy for the valuation of the resource for human’s activities). Broadening this approach, we propose a new methodology to encompass the whole satellite’s population in the LEO region. We define several archetypes for space missions in order to perform a systematic valuation of the orbital asset: Communication, Earth observation, Technology demonstration satellites etc. Our work focuses on the assessment of the economic value (especially the direct and indirect use value seen from an anthropocentric point of view (4)) of communication satellites as well as Earth observation satellites as they both cover 75% of the whole population of the active satellites.
The presentation will highlight the relevance of applying economic valuation to the orbital resource for the complete life cycle of space missions. We will focus on Communication and Observation satellites depicting the potential cumulative economic value of the orbital bins based on the 2018 reference year. It’s a first step to consider the global loss of value for the society allowing a damage characterisation at the Endpoint level for LCA of orbital systems. Further work could be envisaged covering the whole current population of active satellites but also anticipating the future trends in term of launches. In this way, we will have a more accurate overview of the orbital resource valuation particularly taking into account the mega constellations expected for the next years.
1_Verones, F., Bare, J., Bulle, C., Frischknecht, R., Hauschild, M.Z., Hellweg, S., Henderson, A., Jolliet, O., Laurent, A., Liao, X., Lindner, J.P., Maia de Souza, D., Michelsen, O., Patouillard, L., Pfister, S., Posthuma, L., Prado, V., Ridoutt, B., Rosenbaum, R.K., Sala, S., Ugaya, C., Vieira, M., Fantke, P., 2017. LCIA framework and cross-cutting issues guidance within the UNEP-SETAC Life Cycle Initiative. J. Clean. Prod. 161, 957–967. doi:10.1016/j.jclepro.2017.05.206
2_Sonderegger, T., Dewulf, J., Fantke, P., Souza, D. M. De, Pfister, S., Stoessel, F., … Hellweg, S. (2017). Towards harmonizing natural resources as an area of protection in life cycle impact assessment. The International Journal of Life Cycle Assessment, 22(12), 1912–1927
3_Esteve, R. (2017). A Valuation Framework for the Orbital Resource.
4_Bontems, P. & et Rotillon, G. L’économie de l’environnement. La Découverte, 2013
It is now well established that spacecraft manufacturers and launch
service providers have a duty of care to understand and mitigate the
ground casualty risk posed by their products such that this risk can be
demonstrably managed within acceptable levels. In order to achieve
this, an ability to determine the likely ground casualty risk that a
vehicle poses given its as-designed configuration and applicable
missions is required. Further, a design capability which enables design
changes to be implemented, which aim to ensure that system hardware
will fragment and demise in such a way as to mitigate ground casualty
risk, is highly desirable. These types of assessments rely heavily on
analysis, but ensuring robustness of the conclusions drawn from basic
physical understanding, limited ground testing and a small number of
existing flight observations is only tentatively achievable, and could
be significantly enhanced by a carefully conceived flight experiment or
set of experiments. This presentation provides some initial
considerations for the design of such an experiment. These include:
choice of target (launcher vs spacecraft), experiment type
(phenomenological, quantitative or remote observation) and host type
(sub-scale, full-scale or even real-mission). The pros and cons of a
number of initial concepts have been preliminary examined in
preparation for a systematic trade and feasibility assessment.
According to international safety guidelines the on-ground casualty risk of a re-entering object shall not exceed 1 in 10,000. The casualty expectancy can be reduced in two ways (1) by selecting a suitable impact area and population density within, or (2) by reducing the casualty area of the surviving fragment. Due to the high cost associated with a controlled re-entry the latter option has attracted alot of attention. The fragments which survive are often from recurring spacecraft components (such as Propellant tanks, Reaction wheels, Solar array drive mechanisms, Magnetorquers, etc.), therefore the interest of applying designs which increase the demisability of these components is high. In the frame of ESA’s activity “High Fidelity Re-Entry Simulations on Critical Spacecraft Platform Equipment”, HTG conducted detailed SCARAB simulations in order to assess the breakup and demise behavior of solar array drive mechanisms (SADMs) and, together with the manufacturer, identified feasible design for demise options and evaluated their impact on the causalty risk.
In this abstract the final results on the ESA TRP study (ITT AO/1-8301) “demiseable Propellant Tanks Materials and Technologies” are presented.
The replacement of Titanium with an aluminium alloy is confirmed to be the most promising approach given the major impact in terms of tank demiseability. Based on the design trade-off, which was confirmed by the material testing performed, it was concluded that the best approach will be to focus on conventional technologies, available from launchers (e.g. 2219 alloy, TIG welding), so that LEO platform evolutions can be achieved with a reduced effort and at minimum risks. The report presents the results of the material tests performed (demise characteristics, compatibility with hydrazine, PMD wettability) and provide a set of recommendation for future developments.
The object of this study is to further investigate the break-up processes of a reaction wheel during re-entry. The ball bearing unit (BBU) was already identified as a key element (demising late) during the demise process in previous studies. Potential design changes are analyzed with regards to the optimization of demisability. First simulations show that a reaction wheel with an aluminum rotating mass (and with no design modifications) is demisable at and above release height of 87 km. A reaction wheel with a steel rotating mass is demisable at the same release height with only minor design modifications.
In order to improve reduce ground casualty risk on ground, technical solutions need to be identified which promote atmospheric demise of the spacecraft and respective components. One approach is to enable the early opening of external structural elements in order to begin the demise of internal components earlier. Taking into consideration the results of testing of state-of-the-art joining technologies, a broad range of concepts for applications of technologies which would result in earlier joint demise were identified for potential further investigation. This showed a considerable preference towards leveraging the introduced heat and enabling clear and decisive separation. Six of these concepts were then selected and further refined and developed, taking not only design considerations into account but also potential effects on safety, processes, and potential costs. Of these, two part inserts using solder material to enable early melt and release, cleats bonded to the surfaces of the panels, and composite inserts will be developed into breadboard models and undergo a range of tests in order to compare them to current state-of-the-art technologies and assess their demise performance.
Previous D4D studies at system level have clearly shown that, in order to reduce the risk posed by re-entering satellites, an integrated approach to design for demise is more efficient or, for large satellites, even required. Re-entry simulations demonstrated that solutions at component or equipment level might not be enough to effectively reduce the re-entry casualty risk, and that a system-level approach is almost invariably better, or even necessary. The early break-up of the spacecraft main structure (in particular if adopted in combination with re-designed critical elements), or the early separation of critical payloads, can improve the overall demisability, significantly reducing the casualty risk on ground.
The solution presented here is the innovative system, referred to as Demisable Joint (covered by the TAS-I patent N.TO2014A000998.), based on a standard Aluminium joint in which two modifications are implemented:
• one of the aluminium cleats is modified introducing an opening,
• a washer, made of a material with low melting point, is adopted replacing standard aluminium washer.
During the operative life the elements are kept together by the force exerted by the screw, but during re-entry, when the demisable washers demise, the two elements of the joint are free to slide transversally w.r.t. the screw axial direction. Therefore, due to the mechanical loads experienced during re-entry, they will separate. The detailed re-entry simulations performed, showed that replacing current joints for connecting S/C external panels with this re-designed joints will cause the S/C to break up at an altitude higher w.r.t. the one achieved with current joints design.
The main advantages and the novelty of the solution proposed are the extreme simplicity of its design, that consists is a minor, but crucial, modification of S/C current joints, with very low system level impacts and very low cost. The Demisable Joint can be implemented in current satellites without need of modifying the current overall S/C design, manufacturing and integration process.
In conclusion, the Demisable Joint is a simple solution which, if proved effective by test, could be largely applied to future satellites.
ALTRAN Research Team created in 2013, initially intented to focus on space safety topics. Due to the emergence of space debris recommentations anad regulations for space industry, the research team initial activities adressed the flight and ground aspects (satellite robustness to micro-meteroids and orbital debris & trade-off between atmospheric controled reentry and design for demise).
With involvement inside THALES consortium on ESA DESIGN for DEMISE activities on satellite and optical payloads , the team proposed several Building Blocks to the CLEANSAT Initiative and was selected to investigate further the BB10_SMA Dismantlement Mechanisms & BB11_Demise Reaction Wheels.
The intention of this presentation is to remind the outcomes of those building blocks and the evolutions and progress since to prepare the roadmap for the upcoming CLEANSAT Development Phase . A lot of spin-off topics and related activities have been identified and investigated. The presentation will present those assets as well and the potential benefit for the CLEANSPACE community. In addition, the presentation intends to highlight the evolutions of those RESEARCH activities in line with CLEANSPACE topics :
In parallel, the initial ALTRAN Research Cannes team has enlarged its Space Debris topics to the launcher industry mostly with an activity with CNES_DLA launcher directorate concerning debris mitigations scenarios for dual launch structure (AR6_SYLDA) . The team has since investigated several preliminary innovative scenarios in CNES RFP concerning launcher debris mitigations and reductions :
- ST1R : Suborbital Reentry / Recovery & Reuse of 1rst Stage (ie Launch Booster)
- ST2R: Orbital Reentry / Recovery & Reuse of 2nd Stage (ie Upper Stage)
- ST3R : No Orbital Reentry / Reuse in Orbit of 3rst stage (ie Upper Stage or spacecraft vehicle)
Recently, several new activities have emerged within ALTRAN-FR in parallel of MMOD topics (Mitigations Measures for Orbital Debris) adressed by original ALTRAN Cannes Research team . This trend is now so-called “GREENSPACE” intending to adress same perimeter as CLEANSPACE. It is then since including :
ADR : Active Debris Removal topic with CIR Projhect “Space Cleaner” (an investigative scenario of small size debris collector in a large quantity and high cadence rate)
GREENSAT : “Life Cycle Assessment” for satellite LEO & GEO at preliminary developpement phase in line with evolutions Industry 4.0 trends.
As more and more man made space objects orbit around the Earth and crucial orbits are crowded with end-of-life objects. The use of deorbiting maneuvers is a solution to free crucial orbits from near end-of-life space objects by sending back on Earth while controlling the location of the reentry in order to minimize the risk for human populations. Yet, even during a controlled reentry, the reentry is subject to uncertainties, coming from the initial conditions, the material response to atmosphere forces and heatflux, the atmosphere conditions etc. In order toproduce a robust measure of the Gound human risk, the model uncertainties mus be included. In this work, we propose to use a probabilistic predictor to increase the robustness of the risk predictions while maintaining the computational cost at a reasonable level. The approach includes uncertainties coming from unknown model parameters or inputs and the uncertainties coming from our lack of understanding about the breakup. The breakup event is modeled as a random event sampled from an underlying distribution. Our predictor naturally provides the ground risk distribution from which can be derived statistics of interest. In order to keep the computational cost at minimum, the generation of the ground risk distribution is accelerated with a surrogate model.
In the frame of the French Space Operation Act (LOS) signed on 3rd June 2008, CNES is particularly interested by re-entries of space debris. CNES is indeed in charge of ensuring the right application of the law, for every mission launched or operated from the French territory. To predict the debris survivability during their re-entries and assess the prospective risk on ground, the development of complete multidisciplinary tools is required.
With this in mind, CNES develops its own spacecraft-oriented tool named PAMPERO since 2013. PAMPERO aims to simulate the complete atmospheric reentry of an entire satellite, launcher or the associated fragments due to the breakup process. These characteristics are the following:
The purpose of this communication is to highlight the latest developments in the PAMPERO code, such as:
Validation test-cases will be also presented.
Historically, re-entry risk assessment has followed a process which defines the calculations required to be made. Recently, the limitations of this relatively simple process, with respect to both the physical phenomena and the attempts to specifically alter the design of spacecraft to promote demise, have become evident.
Significant recent progress has been made in the understanding of spacecraft heating and break-up during re-entry though recent studies and test campaigns. This includes experimental testing of basic material properties, joint fragmentation phenomenology, and demise of complete spacecraft components. Theoretical work on aerothermodynamic heating to compound shapes has also provided further evidence that the local length scales drive much of the heating rate.
These new findings allow an improved assessment of the uncertainties in both the physics and the modelling of destructive re-entries to be performed, as well as the assessment of potential biases in the current generation of destructive re-entry tools. In particular, this provides a potential basis for the capture of relevant physics by which design-for-demise techniques could be properly verified.
Construction of a statistical tool which will allow this new knowledge to be captured, and the uncertainties to be respected, is part of a new ESA study. This study has the objective of assessing how a procedure for assessment of the casualty risk from destructive re-entry can be realised in a pragmatic way commensurate with the available research findings. To assure this, the activity will make use of the scientific experts in the field as well major system integrators. This procedure will be tested on a range of spacecraft and re-entry types, simulating the mission development phases, in order to ascertain its usefulness for risk assessment.
Spacecraft that survive their nominal mission lifetime are generally proposed for a mission extension in order to maximize their return on investment, to validate new technologies or to have a higher flexibility in the replenishment of a constellation.
However, a life extension has some limits and threats, especially when satellites units are operated well beyond their qualified lifetime: the performance may be degraded and the probability of having unexpected behaviors may increase over time. This may lead to an unsuccessful End of Life (EoL) disposal.
In fact, a low success rate of the disposal manoeuvers is currently observed in LEO, despite the requirements dictated by the international Space Debris Mitigation (SDM) standards. The main reasons have been often linked to:
- being old satellites not designed to be disposed and passivated;
- a disposal decision based mainly on the remaining propellant mass;
- a not-attempted start of the EoL manoeuvres or a too late decision, after the occurrence of major failures.
It is therefore clear that satellite design, reliability models and decision-making process have to be improved in order to dispose satellites in a safe manner and at the right time, before completely losing the disposal capability.
In order to address these challenges a new reliability model is proposed where systems engineering aspects, innovative reliability approaches including Risk Assessment, real operating conditions and events observed by satellite operators are taken into account to support the decision on how and when the satellite has to be disposed.
A new short-term reliability criterion is also proposed in addition to the current one based on consumables. The disposal success rate and the reference duration are currently being chosen in order to decide for a disposal initiation only when a too high risk of losing the satellite exists, avoiding to interrupt a mission that could be extended.
This reliability criterion will be even more important for future missions since the post mission disposal of constellations is the major contributor to the future presence of debris in LEO, and because the propellant mass criterion could become less adequate or at least useful to decide when to dispose the satellite with future on-orbit refueling and servicing missions (space tug).
In conclusion, the final goals are to have satellite designs fully compliant to SDM requirements and to be able to dispose the satellite at EoL in a safe and reliable manner and therefore limit the proliferation of space debris in already crowded LEO and GEO orbits.
Most medium and large LEO spacecraft will contain parts which are likely to survive a re-entry. When too many of such parts exist on any one satellite, this can pose a risk to humans on ground, above the regulatory threshold, and a controlled re-entry, with a well-defined impact footprint in an uninhabited area becomes necessary. However, this is implemented at a cost in complexity and mass that may well drive the satellite design beyond the capability of low-cost launchers, such as VEGA.
Over the last few years, the concept of semi-controlled re-entry has been considered and studied in further detail. In such a scenario, the probable debris fallout zone is comparable to that of an uncontrolled re-entry, but the certainty of its location can be smaller than one orbit ground-track of can reach a few orbit ground-tracks. The semi-controlled re-entry minimises the casualty risk by ensuring that these successive ground tracks are uninhabited. This is possible by controlling the satellite down to a final orbit which will re-enter the atmosphere within a day or two, and by reducing and/or managing uncertainties during this uncontrolled final descent.
Airbus has been investigating this concept through internal research as well as studies for CNES and ESA. This presentation highlights the result of the latest study for ESA, including the simple sizing rules for determining applicability of the approach to future missions, and evaluation of the performance of the semi-controlled re-entry applied to three study cases.
With the growing number of satellites launched in Space, Agencies have implemented rules to remove satellites at their End-of-Life from space by themselves or actively with Active Debris Removal missions. However, this poses a new threat on Earth, with surviving parts after re-entry, and in space, as they can collide with other objects while decreasing their altitude. Complying with these regulations also impacts the overall mission as it requires more fuel and more complex systems which have to be functioning at End-of-Life, meaning sometimes decades after being lunched.
In this study, one asses different systems which can be embedded on a satellite in LEO, evaluates their impacts on the system and, then, proposes a trade-off of the different options. Specifically, one focuses on three categories of spacecraft: small (750kg), medium (1500kg) and heavy (4000kg). The main propulsion system can be monopropellant-based (hydrazine), Hall Effect Thrusters or an Arcjet. However, one considered that the final burn had to be made with monopropellant-based thrusters in order to have a thrust high enough to perform a controlled re-entry. It may require a second independent propulsion system, or one using elements from the main propulsion system. Pressurization is done in blowdown mode or with an active pressure regulation. Finally, one analyses the impact and potential benefits of injecting a satellite at a lower altitude and then increasing it with the main propulsion system of the spacecraft.
Considering all those elements, one proposes a recommended configuration for the different categories of spacecraft chosen.
An overview of the current status of Pre-Qualification of Aluminium-Free Solid Propellant is given in this presentation. Advantages of solid propulsion for satellite de-orbit are shown and proposed mission scenarios are the source of requirements of the system, motor and propellant itself. The first step in development of Solid Rocket Motor for deorbitation was made with elaboration of a dedicated propellant composition. Specific demand for low burn rate and typical aim for high performance required an extensive trade-off. Basing on a literature review, suitable methods of reducing linear regression rate were defined. The optimal composition selection, basing on initial laboratory test results is presented. Safety assessment results are given as a vital part of the pre-qualification process. Long in-orbit storability concerns were shown and analysed. Results of initial tests are given, including radiation impact. Final verification is done via SRM hot firings with reduced burn duration. An important part of the project is dedicated to the preliminary Solid Rocket Motor for deorbitation design since propellant performance can be analysed only with respect to the motor. Concept of a Thrust Vector Control is given. Finally, system level aspects and integration with the spacecraft is proposed as an outline for future development.
One major source of new space debris are spacecraft that are not removed from orbit after the end of their operational lifetime. Many regulations (e.g. ISO 24113) require the removal of spacecraft at the end of operation - known as Post Mission Disposal (PMD) - with a compliance rate of 90% to ensure that spacecraft do not become a new source of space debris. An analysis performed by ESA shows that the success rate of PMD in 2013 was in the range of about 50%-60%.
The goal of TeSeR (Technology for Self-Removal) is to take the first step towards the development of a cost-efficient and highly reliable PMD module. This PMD module is to be attached to the spacecraft on ground and it shall ensure the PMD of the spacecraft at the end of the operational lifetime. This PMD module shall be scalable and flexible, thus, enabling the PMD of any future spacecraft in an Earth orbit. Ultimately, the gap between the required 90% PMD success rate and the current success rate can be closed.
The technological enhancements and developments required for successful PMD are addressed and analysed in TeSeR. The project’s primary aims are:
• Develop, manufacture and test an on-ground prototype of the PMD module
• Develop three different removal subsystems (solid propulsion, electro-dynamical systems and deployable structures) for easy plug-in/plug-out implementation to the PMD module
This is the first step to demonstrate the main aspects of such a PMD module and the required main technologies. The technical activities are supported by non-technical tasks, e.g. investigation of legal issues relating to a PMD module, execution of a market study and consideration of this technology as a leverage to advance ISO norms. This double tracked approach ensures that the technological developments are embedded into the needs of the space community right from the start.
Up to now the prototypes of the three removal subsystems have been developed, manufactured and tested with a common interface for implementation into the PMD module prototype. The PMD module prototype will be manufactured until summer 2018. Afterwards the removal subsystems will be integrated via the same interface.
Airbus is the coordinator (and potential launch customer) of TeSeR. The project is conducted together with 10 notable institutes and companies from all across Europe with experts who have been working on the space debris issue for many years.
The presentation describes the project status in greater detail.
Deorbiting a large spacecraft at the end-of life from LEO requires a robust system. For a defunct satellite, which cannot rely on its primary systems any further, a dedicated deorbiting package is preferred. Providing high thrust, a solid rocket motor can be a reliable controlled deorbiting method, which requires an equally dependable component for steering to the required trajectory.
Almatech was selected to investigate, and design a thrust vectoring mechanism as part of the Solid Propellant Autonomous Deorbit System (SPADES). In the scope if this project, Almatech has explored several solutions for thrust vectoring at the end-of-life. Constraints of the application drive the design towards a flexible and modular concept, with compliant parts to avoid degradation in orbit, and with solutions that decouple the structure from the thermal environment induced by the rocket motor.
A frictionless flex-gimbal mechanism with novel custom flexures was chosen for detailed design due to its relative simplicity, good performance, reliability, and ease of testing. The technology – with innovative, customizable pivots that have built-in movement limiters and allow the use of a compact, dedicated launch lock system – provides a versatile solution for autonomous trajectory control, eliminating the need for a system that is deeply integrated into the vehicle.
We show the main design drivers of such vectoring mechanism, the resulting novel thrust vectoring concept, as well as present outcomes and future steps of the project.
In the last years D-Orbit has been involved in five major projects at European level on the topic of space debris.
Three projects were in the frame of the ESA'e Clean Space Initiative, while the other two were founded by the European Commission in the frame of the Horizon 2020 programme.
All those programmes are completed or about to be completed: these allowed D-Orbit to consistently advance in technology and expertise about space debris mitigation and remediation. D-Orbit operates in this market since 2011. The core product that was developed through those activities is an autonomous propulsive decommissioning device that would allow a reliable and efficient end-of-mission manoeuvre, leading to End-of-life.
Within the ESA CleanSat project, D-Orbit carried out a concurrent design activity, together with ESA and the three major European LSIs (i.e. Airbus, OHB and Thales-Alenia Space), exploring the feasibility and analysing the requirements of an autonomous decommissioning device, based on solid rocket motor, able to provide large satellites with a direct-re-entry capability. A key technology of solid rocket motors is the Thrust Vector Control mechanism: this feature was studied and developed in another ESA Clean Space activity.
An in-orbit demonstration mission of such an autonomous decommissioning device was partly funded through the Horizon 2020 SME Instrument programme: the D-SAT mission. D-SAT was a 3U CubeSat Satellite equipped with the D-Orbit’s “D3” decommissioning device, that was launched into orbit in June 2017. The D3 device was successfully fired on October 2017, demonstrating all the key capabilities, although the target orbit was not reached.
The D3 device was further developed thought an activity funded thought the Horizon 2020 Protec programme: that was the TeSeR project (Technology for Self-Removal), led from Airbus and carried out by a consortium of 11 companies and academia, of which D-Orbit was the second largest contributor.
D-Orbit believes that the Space Debris are an enormous business opportunity, and is constantly monitoring evolutions from the market and regulatory points of view, refining accordingly its business plans. This expertise was capitalised in another ESA Clean Space activity for assessing the size of the Space Debris Mitigation market.
From those activities emerged a clear need for an high level of autonomy and reliability of decommissioning systems, the need to develop thrust vectoring systems for solid rocket motors, as well as large demand for service-oriented end-of-life strategies. For the future, D-Orbit is working for optimising its technology for larger spacecraft and mega-constellations, looking at the same time to the Active Debris Removal and In-Orbit Servicing market.
The ADEO subsystem is a scalable drag augmentation device that uses the residual Earth atmosphere present in Low Earth Orbit applicable for passive de-orbit of satellites between 1 kg to 1000kg. For initiation of the de-orbit maneuver a large surface is deployed which multiplies the drag effective surface of the satellite. Thereby the drag force is increased as well causing accelerated decay in orbit altitude. Advantageous about a drag augmentation device is that it does not require any active steering and can be designed for passive attitude stabilization thereby making it applicable for non-operational, tumbling spacecraft as well. The ADEO subsystem consists of four deployable CFRP booms that span four sail segments in a truncated pyramid shape configuration. While the sails are made of an aluminum coated polyimide foil, its coating thickness was chosen such that it provides sufficient protection from the space environment. To prove the survivability of the sail material in the space environment over 25 years de-orbiting time, multiple environmental tests were performed at material and sample level, including mechanical strength and stiffness tests, thermal cycling, atomic oxygen exposure tests, UV exposure tests, and high velocity impact tests, as well as crack propagation tests at room and reduced temperature. A fully functional full scale 25 m2 sail demonstrator with one sail and two booms has been subjected to environmental testing including, vibration, rapid decompression, deployment testing in thermal-vacuum environment and ambient conditions showing great performance during the ADEO-1 activity concluded in spring 2017. Furthermore, a dynamical de-orbit analysis has been carried out as part of an ESA De-Risk activity confirming the functionality of the subsystem to de-orbit passively. The next step is now the development of a proto-flight model of the ADEO subsystem under a follow up ESA GSTP program. The activity commenced in August 2018 and will finish at the summer of 2020 with the full qualification test of the de-orbiting subsystem PFM. The presentation at hand will give a summary on the ADEO demonstrator activity, the outcomes of the ADEO De-Risk Dynamical De-Orbit Analysis as well as the reference mission and IOD mission selection as well as the preliminary design of the ADEO dragsail de-orbiting PFM.
ARTICA is a project started in 2015 with the aim to develop a ultra-compact passive deorbiting sail for cubesats. The scope was to parallely embrace the two antagonist aspects of the last decade: on one side the foreseen explosive growth of cubesat launches and on the other side the necessity to pose a limit to the consequent LEO orbit population and space debris generation.
ARTICA has been conceived thus with the goal to provide an affordable and reliable deorbiting system for cubesat compatible with user and satellite developer needs, meaning thus a compact and user-friendly device.
The first prototype of the system has been launched in 2017 on board Cubesat UrsaMaior, consisting in a module of 0.35U deploying a sail of 2.1 m2. Thanks to the experience acquired, a novel device as been realized reducing the occupied volume to 0.2U in order to further encourage the use even on board 1U cubesats. With the same philosophy, an open platform approach has been implemented permitting to the user to differently approach the system according to the mission strategy, from a satellite controlled subsystem to complete stand-alone device.
The novel system has currently been ground qualified and NPC is currently looking for possibilities for in-orbit test and validation.
Solar and drag sailing have been proposed as passive end-of-life deorbiting methods, and technological demonstrators are under development. In the drag dominated regime the required area-to-mass-ratio for deorbiting a sail spacecraft is primarily dependant on the semi-major axis, growing exponentially with increasing altitude. In the solar radiation pressure dominated regime, the required area-to-mass ratio strongly depends on both semi-major axis and inclination of the initial orbit. The deorbiting phase, at least in the first phase, is achieved on an elliptical orbit, not a circular orbit like in the case of drag sail with inward deorbiting.
During deorbiting the satellite passes through the debris environment. The cumulative collision risk can be quantified as a function of the collisional cross-section present in orbit and the time of exposure of this cross-section to the flux of debris present in the environment. The objective of this study, funded by the European Space Agency, is to understand the net effect of using de-orbiting technologies on the future debris population around the Earth. Indeed, the increased cross sectional area will decrease the deorbiting time, however they will increase the collision risk over the deorbiting phase with respect to a standard satellite. We assess the collisions risk of deorbiting satellites using these deorbiting techniques, and the consequence of such a collision in terms of global effects onto the whole debris population. To do that fragmentation models have been devised to define when a catastrophic collision will take place and to characterise the following fragments distribution. Long-term simulation of the whole space object population environment are used to evaluate the net effect of using these strategies by means of the definition of an environmental index.