bluuebunny/crossref_metadata_2025_split huggingface dataset:
Everything you always wanted to demand from a digital Sunsensor but couldn't get
Authors:   Lens R&D B.V., Noordwijk, Netherlands, Leijtens Johan , Broekmans Dick , Stelwagen Frank 
Published: January 2024

True digital Sunsensors have been under development for a number of decades but until now it has proven to be very difficult to produce a radiation hard device that is cost effective enough to find widespread use.
Analogue fine Sunsensors would fit the bill for every mission if it weren't for the Earth albedo sensitivity and the analogue interface.
Adding the processing electronics adds significant more cost and complexity that the costs associated with the sensor itself and avoiding significant albedo errors is only possible by very carefully placing the sensors in combination with restrictions on the satellite pointing.
Having an albedo insensitive sensor with a digital interface available would therefore lead to significant cost savings and reduction of application constraints.
Producing a cost effective radiation hardened true digital Sunsensor is only possible if a dedicated single chip solution is found. Currently Lens R&D B.V and Systematic Design B.V with assistance of ESA experts are working on the development of such a solution within an ESA ARTES program.
In order to remain technically feasible but still cost effective, a number of compromises needed to be made which will be addressed during the presentation.
Although the cost target is twice the price of an ESA qualified Analogue fine Sunsensor, cost savings in the processing electronics will lead to significant cost savings at system level if this target can be achieved.
Cis-Lunar Transfer Vehicle: Mission Analysis for an ESA Transfer Vehicle to the Gateway
Authors:   Airbus Defence & Space, Toulouse, France, Locoche Slim , Delage Rémi , Leblond Philippe , Martens Waldemar , Bucci Lorenzo , Grenouilleau Jessica , Vivenzio Salvatore , Lamy Alain 
Published: January 2024

In 2018, the International Space Exploration Coordination Group proposed international agreements on space exploration: expand human presence into the Solar System, with the surface of Mars as a common driving goal and prepare for space exploration missions beginning with the ISS and continuing to the lunar vicinity, the lunar surface, then on to Mars. The development and deployment of a lunar orbital platform gateway concept, called the Gateway, has been proposed as an intermediate step towards deep space travel. The Gateway would be placed in a Near Rectilinear Halo Orbit (NRHO) which has been identified as a suitable orbit for fulfilling the various mission requirements. The Gateway would, for instance, host crews on their way to the Moon surface or serve as an assembly point for the various elements of lunar landers.
In order to fulfil some of these exploration objectives, the study of a Cis-Lunar Transfer Vehicle (CLTV) concept has been performed at Airbus Defence and Space under the supervision of the European Space Agency: in its initial configuration, the CLTV would be an unmanned expendable space transport vehicle designed for logistic servicing of the Gateway. It provides the capabilities of delivering pressurized and unpressurized cargoes to the Gateway and serves for the Human Landing System Refuelling. The CLTV uses the ARIANE 6 EVO launch services for the launch and then performs the transfer to the Gateway’s NRHO. Once this orbit is reached, the CLTV performs autonomously all the operations until docking to the Gateway.
This study investigates the available Earth to NRHO transfer options with the objective of providing useful information for a mission architecture and system trade-off in order to define one or more transfer baseline. In particular, 1) the launch options – launch into Lunar Transfer Orbits or into Earth bounded orbits such as GTO, launch inclination and 2) transfer types – from short transfers to the NRHO with and without powered gravity assist to long low-energy transfers are optimized and assessed in a high-fidelity model. The feasibility and robustness of these options under realistic operational requirements are analysed and the overall mission performance (payload delivery, transfer duration, launch window, etc …) is assessed and traded providing a broad scope of future transfer options from Earth to the Gateway. Finally, a preliminary assessment of the operational concept of the far rendezvous with the Gateway are shown for the investigated transfer options.
Hardware-in-the-Loop Simulations of a Navigation Camera in a Cometary Dust Environment
Authors:   Danish Technical University (DTU Space), Kgs. Lyngby, Denmark, Toldbo Christina , Henneke Jesper , Leif Jørgensen John , Bodin Per , Courson Robin 
Published: January 2024

This paper presents a novel hardware-in-the loop (HIL) simulation tool with the capability to accurately and realistically test the response of optical systems during a fly-by of a comet through the central coma region i.e. in a heavy dust environment. The HIL simulations are performed using the Dynamic Optical Ground Support Equipment (DOGSE) developed at DTU Space which can expose representative optical sensors to high fidelity, dynamic, optical stimulations on a small, high-resolution screen, thus enabling flexible, yet complex HIL simulation tests.
The simulations are part of a predevelopment activity within the Comet Interceptor AOCS/GNC preliminary design with OHB Sweden leading the effort. The Comet Interceptor mission aims to intercept a Dynamically New Comet or an interstellar body with the closest encounter at a distance of less than 1000 km from the comet nucleus and a relative velocity between 10 and 70 km/s. The spacecraft will be launched towards the Sun-Earth L2 (SEL2) Lagrange point where it will wait up to three years until an interesting target object is identified. The spacecraft will then initiate a transfer phase which will last between 0.5 and 4 years. The final approach starts approximately 60 days before the encounter and mission success depend on the navigation system ensuring that the spacecraft can track the comet nucleus throughout the encounter. The navigation solution is realized with navigation cameras in combination with star trackers, and with an on-board navigation filter which processes and fuses the measurement data.
Due to the inherent uncertainties associated with intercepting an a priori unknown, fast-moving comet with an unknown and likely dynamic dust environment it is essential that new, effective, and flexible methods for testing fly-by scenarios are developed and refined. The DOGSE tool is partially automated to test mission scenarios with a fast turn-around and it can realistically display the translucent coma against an accurate projection of the starry sky (i.e. stars can be observed through the dust). The tool takes as its input a trajectory file and generate synthetic images of the comet based on a detailed cometary dust distribution model in the vicinity of the comet nucleus, including distributions of different particle sizes, their phase angle dependent albedo and Sun-distance dependence.
This paper presents current results from the HIL simulations based on the navigation architecture of the Comet-I mission and describes how the dynamic optical stimulator is calibrated and handled. Results include the operation of the navigation camera and the performance of star tracker functionality from afar to the closets encounter in the coma dust environment.
MSR-ERO Rendezvous Navigation Sensors and Image Processing
Authors:   Airbus Defence and Space, Toulouse, France, Kanani Keyvan , Marchand Alex , Falcoz Alexandre , Kracht Christian , Röll Tristan , Kämpfe Thomas , Lecocq Antoine , Richert Michaël , Goux Pierre , Bomer Thierry , Novelli Léo , Lorenzoni Leila , Sanchez Gestido Manuel 
Published: January 2024

The ESA/NASA Mars Sample Return campaign aims to return to Earth samples of Mars materials. The samples will be collected by the NASA provided perseverance rover, assembled in the Orbiting Sample which will then be injected in the Mars orbit by the NASA provided Mars Ascent Vehicle. The ESA provided Earth Return Orbiter (ERO) vehicle will then autonomously detect and rendezvous with the Orbiting Sample (OS) container in low Mars orbit, capture it, seal it, and safely bring it back to Earth. Airbus Defence and Space, under ESA contract, is designing and developing the ERO spacecraft and, in particular, its vision-based GNC system to be used during rendezvous (RDV).
The GNC system includes two types of vision sensors: a Narrow Angle Camera (NAC) and a Light Detection and Ranging device (LiDAR), integrating their own image processing to provide ERO navigation with measurements of OS relative position.
From the navigation standpoint, the MSR-ERO rendezvous is a non-cooperative rendezvous scenario with a target vehicle (the OS) characterized by an uncontrolled fast tumbling motion. Choice has been made to divide rendezvous into a far range phase (from ~50km to ~500m) and a close range phase (from ~500m to capture). In the far range the vision system is required to provide only OS relative line of sight (LoS) using the Narrow Angle Camera (NAC), provided by Sodern, as main sensor. The NAC has a 4.5deg field of view (FoV) and 1Mpx detector (Faintstar2). It integrates a centre of brightness algorithm robust to proton impacts on the detector. In the close range the main sensor is a scanning LiDAR provided by Jena-Optronik with an adaptable FoV from 0.5° to 40° and a maximum scan frequency of 2Hz which integrates a 3D barycentring processing to assess the OS position from 3D point cloud directly measured by the LiDAR.
This paper focuses on the vision sensors of MSR-ERO spacecraft. In a first section, it describes the MSR mission, including its requirements, and recalls the vision sensor trade-off done in the early phases of MSR-ERO project. Then, the vision sensor architecture within the ERO system and the sensors operation concept (board/ground) are presented in the second section and, in the third section, the selected sensors as well as their image processing are described in more details. An overview of their development plans is provided in the conclusions, together with the preliminary performance test results.
Ariel - AOCS overview, Safe Mode and Normal Mode design
Authors:   Airbus Defence and Space Ltd., Stevenage, United Kingdom, Chapman Patrick , Hamilton Warren , Aguiar Goncalo , Broquet Renaud , Sechi Gianfranco , Saavedra Criado Gonzalo 
Published: January 2024

Ariel (Atmospheric Remote-sensing Infrared Exoplanet Large-survey) is an ESA scientific mission whose mission is to measure the spectrum of a large population of known transiting planets in our galaxy. To achieve this, it has a large telescope and an infra-red spectrometer payload, which has challenging AOCS requirements.
The main driving requirements for AOCS are sun angle exclusion zone requirements, which apply from launcher separation, and the fine pointing mode requirements for the observation phase at the sun-earth Lagrange L2 orbit. Fine pointing requires reaction wheels with built-in speed controllers, and using payload Fine Guidance Sensors in the loop.
This paper presents an overview of the Ariel mission and the AOCS design, particularly the safe mode and fine pointing mode.
Using a Star-tracker in degraded modes: outputs of the STEAM R&D study
Authors:   Airbus Defence And Space, Toulouse, France, Lagadec Kristen , Aguiar Gonçalo , Rudolph Tobias 
Published: January 2024

Star trackers have become ubiquitous on space missions. Their versatility is unparalleled, while their track record belies their apparent internal complexity. One counter-intuitive benefit arising from said complexity is that they have no trivial failure mode in which an invalid attitude measurement can be wrongly flagged as valid. This allows to trust them enough to include STR measurements in the mix of AOCS observables in safe mode.
The approach inspiring the STEAM study was to investigate the possibility of using star trackers as the backbone for a generic safe mode architecture (which can be instantiated with or without the STR). We will be presenting the main outcomes of this study:
- Verification of robustness of real STR HW to high angular rate and high radiation levels with optical stimulation test benches
- Architecture definition of a generic safe mode, with a versatile attitude determination module that can incorporate STR measurements seamlessly in the overall attitude determination mix, while still capable of sun-pointing convergence in the absence of STR or gyro data.
- Detailed performance simulations for the safe mode, when instantiated for 2 very different and very challenging science mission safe modes: an Ariel-like and an Envision-like use case.
- Real-time closed-loop tests with the STR in the loop confirming the good behaviour.
A main conclusion of the study was that all of the perceived obstacles to the use of STR measurements in safe mode have been successfully addressed and overcome, and that future projects could greatly benefit from the added flexibility and functionality.
Low-thrust deorbiting after dual reaction wheel failure: autonomous attitude guidance and hybrid actuator control mode
Authors:   Airbus Defence And Space, Toulouse, France, Aymon Rémi , Caravaca Rémi , Ozenne Antoine 
Published: January 2024

For large constellations comprising several hundred satellites, the risk of double failure will no longer be extremely improbable. Today a configuration where only two reaction wheels out of four remain functional (in a 3-for-4 redundancy scheme), means a complete loss of attitude control, especially when the bus does not have a chemical reaction control system to fall back to. The resulting inability to perform deorbiting operations would leave a derelict satellite at the mission altitude of the constellation, becoming a permanent collision hazard. Avoiding this risk today for sustainable space development calls for adding a redundant wheel to the avionics design, or deorbiting the satellite after the first failure.
To address this, Airbus DS has developed a new de-orbiting mode called CONDOR (CONtingency DeORbiting) which recovers sufficient controllability to effectively de-orbit the spacecraft using only two functional RWs. The use case is here the ARROW platform, which serves as the generic avionics architecture for constellations and small sat programs, but the mode could be extended to any spacecraft with a magnetic control capability.
The main design principle consists in completing the partial loss of controllability (along the normal to the controllability plane of the two remaining RWs) with magnetic control. As magnetic control authority is generally far below that of reaction wheels, the direction of the magnetic field in the satellite’s frame must be optimized. Along with the need to maintain thrust attitude while protecting the field of view of at least one star tracker against blinding or obstruction at all times, this leads to an over-constrained attitude guidance problem. Using an innovative autonomous potential-based guidance function with a very robust obstacle avoidance feature, we have extensively consolidated the feasibility of guaranteeing 3-axis attitude controllability during the thrust phase, allowing the spacecraft to perform de-orbiting even after a dual RW failure.
Dynamics and Aerodynamic Control of a Cross-Track Tether Satellite System
Authors:   Politecnico Di Torino, Torino, Italy, Aliberti Stefano , B. Quadrelli Marco , Romano Marcello 
Published: January 2024

The objective of this paper is to present and analyze the dynamics and control strategy of a novel type of formation flight architecture, that could enable unprecedented space mission capabilities. In particular we propose and study a tether satellite system composed by two or more satellites, flying in LEO, connected by a linear tether. That tether is maintained in tension along the cross-track direction, with respect to the orbit, i.e., perpendicular to the motion and radial directions, by exploiting the aerodynamic lift acting upon the suitably oriented satellites.
Possible applications of this formation are related, in particular, to remote sensing.
To carry out the study, a three-dimensional multibody model was developed, introducing the equations of relative dynamics centered in an orbiting reference frame.
To simulate the conditions of low Earth orbit, several mathematical models taken from the literature were used, in order to include the Earth’s gravitational potential up to fourth order, solar pressure, third-body perturbations, atmospheric drag and the behavior of aerodynamic surfaces in a free molecular flow.
By utilizing this model and verifying that this configuration is intrinsically unstable, it was concluded that the system needs continuous control to be stabilized.
As a result, the model of the dynamics was linearized and an optimal LQR controller was introduced to calculate the force required to hold the system in place. Once the linear control law was validated by simulating its application on the full system model, aerodynamic surfaces were next introduced in order to generate the force required for control by taking advantage of the rarefied atmosphere of low Earth orbit.
The proposed control strategy is tested by extensive numerical simulations.
Finally, the proposed aerodynamic stabilization approach to continuously maintain a linear tether satellite system along the cross-track direction was compared to a gyroscopic stabilization alternative strategy.
This strategy consists of putting the system into rotation in a direction perpendicular to the cross-track direction. The centrifugal force generated by the rotation stabilizes the system, maintaining constant oscillations in the cross-track direction throughout the orbit.
Both cases have been analyzed as the size of the system and the number of satellites change.
ARIETIS and ARIETIS-NS gyros test results and qualification status
Authors:   Innalabs Ltd, Blanchardstown, Ireland, Torasso Alberto , Diram Romain , Heffernan Alan , Kowaltschek Steeve 
Published: January 2024

ARIETIS and ARIETIS-NS are two gyroscopes for space applications developed by Innalabs based on its proprietary CVG technology. ARIETIS is an ITAR free, Rad-Hard 3-axis gyroscope whose main applications are Earth Observation, Science missions, Telecom, as well as Navigation, Earth Observation, as well as Science missions. It uses class 1 EEE components and meets ESA ECSS standards. ARIETIS-NS is an ITAR free, rad and Single Event Effects (SEEs) tolerant 3-axis gyroscope. It is developed following ECSS standard with the exception of using commercial EEE components that are upscreened by means of radiation testing. ARIETIS-NS is intended for commercial satellites (both LEO and GEO), constellations and mega-constellations, but it has also been selected for science and exploration missions.
ARIETIS and ARIETIS-NS are both based on InnaLabs proprietary Coriolis Vibratory Gyroscope (CVG) technology which is currently used in Space on board both LEO and GEO applications. InnaLabs CVG gyros have already cumulated more than 3,000,000 hours in flight with no failures and no deviation from the specified performance, confirming the suitability of the technology to space environment.
After a brief description of the InnaLabs CVG basic principles and an overview of the CVG technical strengths in comparison to competing technologies, this paper describes the specification. key design features, and budgets of both ARIETIS and ARIETIS-NS. Both gyros are able to provide high performance (ARW <0.005 deg/√hr) and high reliability (~1000 FITs) in a very compact design (~1.2 kg for ARIETIS-NS and ~3kg for ARIETIS) and very low power consumption (<6W). To achieve this, several targeted enhancements and optimisations to InnaLabs gyroscope's design have been introduced, mainly in the component selection, the control loops, the compensation algorithms, and a smaller Sensing Element design.
ARIETIS-NS has recently completed qualification where multiple qualification models were used to characterize different configurations of the product. Qualification test results will be presented to show the gyro compliance to the specification in terms of performance and compatibility to both mechanical and thermal environment. ARIETIS is at the start of the qualification campaign with the qualification model currently under MAIT. Several ARIETIS engineering models have been built and tested and test results are presented in the paper.
Fast Sequential Convex Programming with Relaxed Convergence Criteria
Authors:   Airbus Defence & Space, Stevenage, United Kingdom, Popov Alexander , Contini Cristiano , Zamaro Mattia 
Published: January 2024

The growing interest in lunar exploration demands the development of spacecraft which can land autonomously on the Moon’s surface. Typically, such missions define an ellipse around some selected landing point and any location within this ellipse is deemed acceptable for touchdown. To ensure a safe landing, a spacecraft can employ a hazard detection and avoidance (HDA) system, capable of identifying potential dangers at the target and triggering diversion to another point within the ellipse. The choice of landing site for a mission is a compromise between scientific priorities and landing safety. However, many scientific sites of interest may not be well-suited to landers as there may be hazards, such as craters or rocks, which greatly limit the size of a potential landing ellipse. To enable a broader range of missions, a lander should be capable of diverting to a new target at relatively low altitude and efficiently reaching the target with pinpoint accuracy. This would not only make it possible to reduce the minimum landing ellipse size for a given vehicle design, but it would also allow for a relaxation of the surface hazard requirements within the chosen ellipse. If this can be achieved, then a much more versatile set of lunar landing sites can be considered, prioritising scientific needs over engineering limitations.
Fast and accurate retargeting demands a sophisticated guidance algorithm which can rapidly compute a trajectory to a new point. This paper presents progress on such an algorithm, which solves a trajectory optimisation problem to achieve fuel-optimal pinpoint landing. To enable late diversion, this algorithm must be able to run on board the spacecraft in real time, and so our research has focused on the use of fast, resource-efficient methods based on sequential convex programming.
We recommend two related strategies which can enable fast termination of sequential convex programming algorithms: Hermite interpolation to extract a continuous-time state trajectory from the convex solver, and mesh refinement to enable efficient convergence towards an accurate solution. In combination, these techniques can ensure that the algorithm terminates after only a few iterations, using relatively relaxed stopping criteria, while still producing a feasible solution which can exactly satisfy terminal equality constraints. The benefits of this framework are demonstrated using the case study of guidance for an autonomous lunar lander with a complete model of the system dynamics in six degrees of freedom.