MITM1
Session assets
In the last 15 years the Payloads Department of INTA has developed a variety of compact sensors for different Mars exploration missions. This includes a magnetometer (72 g), a dust sensor (35 g; with UC3M, Spain) and a radiometer (114 g) for the MetNet penetrator [1]; a radiometer (25 g optical head, 56 g processor) for DREAMS (Schiaparelli) [2], [3]; a radiometer plus camera (1 kg) for MEDA on Perseverance [4], [5]; a 110 g dust sensor (with UC3M, Spain) [6] and a radiometer plus spectrometer (180 g) for the METEO package [7] on Kazachock lander (ExoMars’22) and a 0.5 kg nephelometer (with INAF and Politecnico di Milano, Italy) [8] for the Dust Complex on the same lander. Equally miniaturized sensors exist for the measurement of the most relevant environmental variables, such as radiative balance, air temperature, wind, humidity, pressure, dust saltation, electric field, etc. with enough flight heritage (or technology readiness level) on the same sensors’ suites on Perseverance and ExoMars, as well as Insight or Curiosity before [9].
In summary, a large portfolio of miniature sensors for environmental research is available at present. However, a qualitative leap on (in-situ) Mars climate science will only happen through the deployment of networks of environmental stations throughout large areas of the planet. Given the relevance of these measurement not only from a scientific point of view but also because of their importance for future human missions to Mars, this is an objective considered in several Mars exploration roadmaps such as ESA’s Terrae Novae 2030+ [10].
With this aim, we propose a microprobe named MarsConnect. It consists of a 10-12 kg probe with a rigid, deployable aeroshell/TPS and a 5-6 kg impactor/penetrator carrying up to 1 kg of environmental sensors. Many of these probes could be launched to Mars with a single carrier, to deploy meteorological networks. This works inherits different concepts from previous similar proposals, very specially MetNet and MiniPINS [11], but simplifying even more the EDL concept and reducing the mass, at the expense of an increased impact speed.
The probe’s aeroshell is divided into a backshell and two halves of a frontshield that are opened in the low supersonic regime to drop the penetrator. This one is equipped with a drag-skirt that provides some braking and increases stability. The expected impact speed, highly dependent on the atmospheric density profile, entry conditions and landing altitude, ranges from less than 100 to 140 m/s. The whole system is designed to be compatible with a wide range of scenarios and landing sites and is sized to endure more than one Martian year operating on the planet’s surface.
Fig. 1. CAD views of the MarsConnect probe and its eyectable aeroshell, plus some details of CFD and EDL analysis.
The present development status and short-term plans will be presented.
References: [1] Harri A-M et al. (2017), Geosci. Instrum. Method. Data Syst. 6, 103-124. [2] I. Arruego et al. (2017) Adv. in Space Res. 60 (2017). [3] F. Esposito et al. (2018) Space Sci. Rev. (2018) 214:103. [4] J.A. Rodríguez-Manfredi et al. (2021) Space Sci. Rev. 217:48. [5] V. Apéstigue et al., Sensors 2022, 22(8), 2907. [6] A. Russu et al. (2019) SPIE, Vol. 11129. [7] D. Rodionov et al., Vol. 12, EPSC2018-732. [8] D. Scaccabarozzi et al., 2019 IEEE MetroAeroSpace. [9] J. Gómez-Elvira et al., Space Sci. Rev., 2012, 170:583-640. [10] ESA, Nov.’21. [11] I. Arruego et al., LowCostMars2022.
How to cite: Arruego, I. and the MarsConnect team: Mars environmental networks through the MarsConnect microprobes, Europlanet Science Congress 2024, Berlin, Germany, 8–13 Sep 2024, EPSC2024-92, https://doi.org/10.5194/epsc2024-92, 2024.
The “Mars Magnetosphere ATmosphere Ionosphere and Space-weather SciencE (M-MATISSE)” mission is an ESA Medium class (M7) candidate currently in Phase A study by the European Space Agency (ESA) (Figure 1). M-MATISSE’s main scientific goal is to unravel the complex and dynamic couplings of the Martian Magnetosphere, Ionosphere and Thermosphere (M-I-T coupling) with relation to the Solar Wind (i.e. space weather) and the lower atmosphere, and the processes leading to this coupling, which are highly entangled between several regions of the system (Figure 2). The M-I-T coupling controls the dissipation of incoming energy from the solar wind, and therefore, the evolution of Mars’ atmosphere and climate (including atmospheric escape, auroral processes, and incoming radiation). Moreover, understanding the behavior of Mars’ M-I-T system and of the chain of processes that control Space Weather and Space Climate at Mars, as well as the radiation environment, is essential for exploration as it leads to accurate Space Weather forecasts and, thus, prevents hazardous situations for spacecraft and humans.
Figure 1: The “Mars Magnetosphere ATmosphere Ionosphere and Space-weather SciencE (M-MATISSE)” mission is an ESA Medium class (M7) candidate.
Mission goals: The mission has three main goals:
- Characterising the global dynamics of the M-I-T coupling by unravelling its temporal and spatial variabilities. This will be done with simultaneous observations of the solar wind (energy input) and ionosphere-magnetosphere (energy sink), and also, via investigating the coupling of the mesosphere with the ionosphere and solar energetic particles.
- Characterising the Radiation environment, by determining how the M-I-T absorbs the energy that reaches the planet and forecasting near-real time planetary Space Weather
- Characterising the Ionosphere-lower atmosphere coupling, which is a region barely explored but essential for solar energetic particles related phenomena as well as for communications in the HF wavelengths.
In addition, M-MATISSE will significantly contribute to understand Mars climate and the lower atmosphere as two remote instruments have dedicated instrumentation to monitor dust, clouds, and get temperature and density profiles from the surface up to about 50 km. Moreover, the heliophysics community will count with a full-package solar wind monitor at Mars’ distances, contributing to understand solar wind and solar transient propagation in the inner Solar System.
Figure 2: Mars regions that M-MATISSE will focus to understand the spatial-temporal variability of the M-I-T system and its couplings from the surface to space.
Type of Mission: M-MATISSE is one of the current three candidates in competition at ESA at the Medium-size opportunity in ESA's Science Programme from the call in December 2021. From 27 initial responses, ESA down-selected 5 missions in 2022, which went through a Phase 0 study. In that phase, ESA evaluated the expected science that could be achieved with each mission, as well as came up with a preliminary mission design. In November 2023, ESA further down-selected them to three, which are currently in Phase A studies. In this phase, each candidate mission will be studied in detail by ESA, involving European aerospace companies, national institutes, and universities, resulting in a more comprehensive design for each mission. Payload maturation activities are being performed in parallel including breadboarding and test. It is expected that one candidate mission will be chosen by mid-2026.
Figure 3: Simulation of the Martian plasma system and the two M-MATISSE spacecraft with their nominal orbits.
M-MATISSE mission concept: M-MATISSE consists of two orbiters with focused, tailored, high-heritage payloads to observe the plasma environment from the surface to space through coordinated simultaneous observations. It will utilize a unique multi-vantage point observational perspective, with the combination of in-situ measurements by both orbiters and remote observations of the lower atmosphere and ionosphere by radio crosstalk between them (Figure 3).
The fathership, called Henri, has a periapsis below 270 km and an apoapsis of 3000 km with an inclination of 60°, and is intended to spend most of its time within the Martian plasma system. The daughtership, called Marguerite, also has an inclination of 60°, a periapsis below 270 km and an apoapsis of 10,000 km, and is intended to spend most of its time in the solar wind and/or far tail of Mars (a region barely explored before).
M-MATISSE has a nominal mission duration of 1 Martian a year, and the launch date is identified for July 2037.
Consortium: M-MATISSE is the product of a large organized and experienced international consortium in which 15 different space agencies participate. M-MATISSE is a community-led mission, and a legacy concept from pioneer missions, such as Mars Express, Trace Gas Orbiter, MAVEN and experienced teams.
Take home message: M-MATISSE has the unique capability to track solar perturbations from the Solar Wind down to the surface, being the first mission fully dedicated to understand planetary space weather at Mars. It will revolutionize our understanding and ability to forecast potential global hazard situations at Mars, an essential precursor to any future robotic & human exploration.
Acknowledgments: We acknowledge the ESA M-MATISSE CDF team who supported the assessment phase until MDR and the ESA study team, who is managing the Phase A/B1 activities (point of contact robert.buchwald@esa.int). The Science Study Team is managed by Olivier Witasse as M-MATISSE study scientist.
How to cite: Sanchez-Cano, B. and Leblanc, F. and the The M-MATISSE team: The M-MATISSE mission: Mars Magnetosphere ATmosphere Ionosphere and Space weather SciencE. An ESA Medium class (M7) candidate in Phase-A. , Europlanet Science Congress 2024, Berlin, Germany, 8–13 Sep 2024, EPSC2024-545, https://doi.org/10.5194/epsc2024-545, 2024.
In preparation for the human exploration of Mars, several orbital assets will need to be in place. ESA is exploring concepts for very-high-altitude platforms which will facilitate communications links to the Martian surface amongst other duties. Such platforms would have the capacity to monitor most of the Martian surface simultaneously, while also providing mission critical services including global navigation and Earth-Mars data relay. Part of a possible scientific payload will be instrumentation to monitor the Martian climate over the whole planet to enable future weather forecasting. These observations will eventually provide a better understanding of the formation and evolution of Martian dust storms, and therefore their prediction – which is critical for the safety of human exploration.
Here, we present the Broadband Exploration with Bolometric Optics (BEBOP) concept for these missions. This instrument is a thermal imaging system that combines a filter array with heritage from Mars Reconnaissance Orbiter’s Mars Climate Sounder (MCS; McCleese et al., 2007) and Lunar Reconnaissance Orbiter’s Lunar Diviner (Paige et al., 2010) with a fast, wide field of view, compact freeform optical system and uncooled microbolometer detector. The optical configuration, detector, and electronics have heritage from the Lunar Trailblazer Lunar Thermal Mapper (LTM; Shirley et al., 2020; Bowles et al., 2020) and Comet Interceptor’s Modular InfraRed Molecules and Ices Sensor (MIRMIS; Jones et al., 2024) instruments. Fig. 1 shows the completed LTM assembly, which was delivered in 2023 and is awaiting a November 2024 launch. The filter assembly will have 15-19 channels covering a spectral range of 6-25 μm, including the 15 μm CO2 feature. The instrument is compact, low mass and power, and does not require cryogenic cooling for the detectors. On-board time delay and integration (TDI) leads to high sensitivity and low noise. A scan mechanism and internal black-body target calibrates the entire optical chain between observations.
Fig. 1 The fully assembled LTM instrument, now mated to Lunar Diviner and ready for a November 2024 launch.
The spectral range of BEBOP and the necessarily wide field of view at high orbit will allow the measurement of atmospheric parameters across nearly the entire Martian disk. From MCS heritage, we will have spectral bands covering the 15 μm CO2 band, allowing the retrieval of temperature and pressure of the lower Martian atmosphere (Kleinböhl et al., 2009; Smith et al., 2022; Vlasov et al., 2023). To either side of this band, dust and water ice aerosol opacity can be retrieved, providing column opacities over the Martian disk. Other spectral channels that will be included will be thermal bands, providing Martian surface temperatures with high precision, and a series of mineralogical bands over the 7-10 μm region to determine crustal composition via Christiansen feature mapping. We will also be able to monitor surface ice and frost coverage, identify clouds and dust storms, trace the movement of clouds and dust, and extract wind fields. The ability to include a bandpass covering the emission and absorption of water vapour and other gases is under consideration.
Expected spatial resolution is 1.5-2.6 km from a 5700 km altitude orbit. The field of view extends across the entire Martian disc, additionally facilitating limb sounding to retrieve vertical profiles of temperature, pressure, dust extinction, water ice extinction, and possibly water vapour with a vertical resolution of ~5 km. With three spacecraft, this will be done pole-to-pole at six longitudes at high cadence, having nearly global coverage each Martian day. This will lead to a better understanding of the dust and water cycles on Mars, providing insights into contemporary and past climate. A key question is how do dust storms form and how do they transform into global events?
The instrument will also provide valuable information about the surface mineralogy, accessing longer wavelengths than contemporary instruments such as CRISM and OMEGA. This will allow us to address the crucial scientific question: what is the crustal history of Mars? The formation of the Martian crust was a complex process and the origins of its magmatic and volcanic content are unknown, and their study will lead to better understanding of the history and formation process of Mars and, therefore, Earth.
The 6-25 μm range includes emission peaks for silicate mineral Christiansen features and silicate minima within the Reststrahlen bands. These allow the differentiation between plagioclase, olivine, and pyroxene. The surface spectra will inform about mineralogy and help answer the outstanding question of whether phyllosilicates (Fe/Mg) are smectites or the intermediate material in the diagenetic sequence from smectite to chlorite, illite, and other higher-temperature clays.
How to cite: Olsen, K. S., Evans, R., Eshbaugh, H., Warren, T. J., Shirley, K. A., Nowicki, K., and Bowles, N. E.: Returning to Mars with BEBOP (Broadband Exploration with Bolometric Optics) , Europlanet Science Congress 2024, Berlin, Germany, 8–13 Sep 2024, EPSC2024-1093, https://doi.org/10.5194/epsc2024-1093, 2024.
Introduction
Current missions to the Martian surface are marked by the slow and expensive development of complex rover platforms. Although these missions have provided indispensable initial insights about Mars, the datasets are sparse in nature and coverage is limited to the vicinity of the landing sites. The Mars Sample Return campaign has drawn attention to the challenges of the current paradigm, which involves infrequent, high-risk missions based on high-cost platforms.
While Mars orbiter missions offer global coverage, remote measurements are fundamentally limited in resolution. Hence, a new approach is required to provide large datasets at high spatio-temporal resolution to characterise the surface effectively.
Mission Architecture
The Tumbleweed mission was conceived to accelerate the pace of surface exploration and data acquisition on Mars and to fill the capability gap that exists between orbital and conventional rover platforms. By leveraging an architecture based on a swarm of wind-driven Tumbleweeds, extreme terrains can be accessed at a reduced risk of mission failure.
The spheroidal rovers are five metres in diameter with a rigid outer shell and a stabilised inner structure. Sails will provide the motion with flexible solar cells supplying the power. The majority of the scientific payload will be housed in a pod, suspended one metre above the surface. Owing to the simplicity of the rover’s main construction and the use of passive propulsion, manufacturing and development costs can be scaled down considerably. The dispersion of the 90 folded rovers will happen while unfolding in the mid-martian atmosphere (Figure 1, stage 6). Once on the surface, the rovers will start the rolling stage with an intended duration of 90 days. During the rolling stage, the rovers will traverse across rough terrain, utilising the diurnal and katabatic winds to start and stop (Figure 1, stages 7 and 8). After sufficient dispersion, the rovers will transform into weather stations. This would result in the creation of a distributed sensor array spread across the surface of Mars, to provide holistic datasets of the Martian atmosphere for example (Figure 1, stages 9 and 10).
Figure 1: Schematic of the Tumbleweed mission architecture.
Optimizing Instrumentation
About fifty instruments that could address the previously determined open research questions on Mars were identified. Hence, we devised a methodology employing 7 modifiers (A-G) to score instruments based on their alignment with both mission requirements and research questions, enabling the selection of the most suitable instruments. These modifiers assign points for meeting specific criteria, with A, B, and C weighted as 3, 2, and 1 points respectively, considering primary, secondary, and tertiary data relevance. Additional scores (D-G), functioning as binaries, indicate other achievements such as the ability of instruments to produce quality data under a wide array of conditions (D) and whether having multiple rovers with the same instrument increases the value of the findings (E). Moreover, Criteria F is met when an instrument appears to have been used in scientific literature to tackle the same research question. Additionally, Criteria G is awarded when an instrument can perform its task without requiring the entire rover to momentarily stop moving.
Normalisation within each science case yields cumulative scores, representing the expected science return of each instrument.
To optimize instrument selection, we conceived the instrumentation trade-off as a 0-1 multidimensional knapsack problem, which is a special case of 0-1 Linear Programming. More specifically, our goal was to maximize the total science score of the selected instruments to be carried on the Tumbleweed payload, such that the selection does not surpass a total mass of 5 kilograms, a total volume of 6000 cubic centimetres and 20 watts of power consumption. Through performing mixed integer linear programming (MILP) algorithms and a brute force method the optimal set of instruments that maximise science return is shown below, discretised into scientific fields:
- Human Exploration - to identify candidate human exploration sites, based on both resources and scientific interest and deeply characterize them while also delivering a better understanding of potential hazards:
- Stereoscopic camera
- Neutron spectrometer
- Surface Geology - to discover and characterise Martian surface deposits, such as potential turbidite structures and glacier-like forms (GLFs), while also detecting surface compound materials.
- Hand-lens style imager
- Stereoscopic (multispectral) camera
- Interior Geology - to illuminate the remnant crustal magnetism found in the southern highlands of Mars.
- Triaxial fluxgate magnetometer
- Ionizing Radiation - to advance our knowledge of the ionizing radiation environment on Mars, through distributed monitoring of radiation exposure and particle spectra:
- Radiation spectrometer/particle camera
- Neutron spectrometer
- Atmospheric Science - to investigate lower atmosphere dynamics, thermal properties, water and CO2 distribution, surface-atmosphere interactions, and chemical species variability across diverse latitudinal gradients and temporal scales:
- Electric field sensor
- Wind sensor
- Dust sensor
- Pressure sensor
- Temperature sensor
- Soil pH sensor
- Relative humidity sensor
- Astrobiology - to uncover biosignatures ranging from textures to organic molecules and other CHNOPS-containing compounds, where contextual measurements of the surface geology and atmospheric quantities help identify potential zones of interest.
From the optimisations executed it became clear that peak power is the limiting factor when selecting instruments to maximise scientific return. A very large increase in power availability is required (365%) for it to be the non-dominant limiting factor in the optimisation. However, this peak power constraint assumes all instruments are operational simultaneously, where instruments may be idle when others are operational.
Feasibility Studies
The instruments devised above are further filtered by performing extensive feasibility studies. Each potential factor that can be detrimental to the viability of each science case is considered carefully and the most pertinent disturbing factors are simulated. Following this, mitigation strategies are discussed.
Conclusion
The Tumbleweed mission provides a new paradigm for Martian exploration, addressing key limitations of current in-situ missions and providing unprecedented insights into the Red Planet's geology, atmosphere, and potential for habitability. Through an instrumentation trade-off and optimisation, a set of instruments with maximal scientific return are selected for the Tumbleweed mission. Feasibility studies identify and mitigate the technical challenges of the diverse payloads. The findings presented here lay the foundation for further maturation of payload on the Tumbleweed rovers.
How to cite: Kingsnorth, J., Shanbhag, A., Balsemão, M., Bounova, G., Bonanno, L., Pikulić, L., Mikulskytė, O., and Rothenbuchner, J.: Sailing the Martian Winds: Optimizing the Scientific Payload of the Tumbleweed Swarm , Europlanet Science Congress 2024, Berlin, Germany, 8–13 Sep 2024, EPSC2024-1103, https://doi.org/10.5194/epsc2024-1103, 2024.
Over recent decades, laser altimeters have been an essential tool in planetary exploration, enabling precise topographic characterisation of planetary bodies within the Solar System [1]. The working principle of this type of topographic measurement relies on a time-of-flight estimation of laser pulses emitted by the instrument and reflected from the surface of the target. A photon sensor then detects the returning signal, permitting the evaluation of the range between the satellite and the geological features, as well as the variations of intensity and spread of the beam shape. These strategies allow for the reconstruction of altitude, surface roughness, terrain slope, and albedo, thus enabling the creation of accurately calibrated digital maps of planetary landscapes [2].
The European Space Agency's BepiColombo mission to Mercury, as well as the JUICE mission to the Jovian system, carry the laser altimeters BELA (BepiColombo Laser Altimeter) and GALA (Ganymede Laser Altimeter). These instruments, developed by the DLR Institute for Planetary Research in collaboration with national and international partners, use the above-mentioned measurement approach but implement slightly different technical specifics. BELA will study the surface of Mercury from an orbit between 400 to 1000 km, using a 50 mJ Nd:YAG laser with a repetition rate of 10 Hz to cover a 6 to 40 m spot on the target [3]. Conversely, GALA will use 17 mJ pulses with a repetition rate of 30 Hz to determine topography on Ganymede from 500 km of altitude [4].
BELA and GALA have been implemented on large satellites and require significant volume, mass, and power resources. However, there is an increasing trend to employ small satellites and CubeSats not only in near-Earth orbit but also on interplanetary exploration missions [5]. Instruments must be miniaturised to enable laser altimetry for topographic purposes on such missions and open up new possibilities for innovative scientific investigations.
The goal of the current project is to pave the path for adapting the technology developed with BELA and GALA to the new scales being explored by small satellites and CubeSats. Indeed, the mass of the instruments (15 kg for BELA, 25 kg for GALA), their power consumption of about 50 W, and their receiver apertures of about 25 cm diameter [3,4] are not conducive to use on small satellite platforms. Here, we will report on the development of an elegant breadboard for a 3 CubeSat-Units laser altimeter implementing a transceiver design with an 8 cm aperture using Commercial Off-The-Shelf components (COTS) and a SPAD (Single Photon Avalanche Diode) array detector.
In our quest for the most suitable design, we conducted a trade-off study considering 16 possible optics arrangements. The focus was on a Lunar orbiter mission case, aiming to characterise future landing sites and assist in in-situ operations. The driving requirements for the selection were the compactness of the optical system, the laser footprint dimensions, the susceptibility to straylight, the coalignment budget, the transmittance, and the thermal stability. After a preliminary down selection and rigorous analyses of the last candidates in the optical design software Zemax, a final design was identified based on its performance in the above-mentioned requirements. The straight transmitter design combined with a double-deflective receiver performed best regarding internal stray light and coalignment budget, as shown in Fig. 1. This design can be accommodated in a 10x10x20 cm3 volume, equivalent to a 2U CubeSat design, leaving space for the implementation of the other subsystems of the instrument and demonstrating its practical application.
A laboratory setup implementing the design is currently being assembled (Fig. 2), and we will present a preliminary characterisation of the optics performance in terms of the quality of the laser expansion and the coalignment between the laser source and detector boresight axes. These studies will address the technology's readiness level and future development (adjustments) needed by the instrument, currently under Phase 0 study in collaboration with the European Space Agency.
Fig. 1: Zemax model of the final transceiver design. In green, the laser beam expanded by the transmitter optics; in blue, the received signal focused on the detector plane. The ‘L’ labels refer to the lenses, while the ‘M’ labels refer to the mirrors.
Fig 2: Detail from the transceiver design's CAD model showing the different components' arrangement. The ‘L’ labels refer to the lenses, while the ‘M’ labels refer to the mirrors. In green, the laser beam expanded by the transmitter optics; in blue, the received signal focused on the detector plane.
[1] Zhou et al., Infrared Physics & Technology 86, 147-158, 2017. [2] Kallenbach et al., Journal of Geodesy, 95:4, 2021. [3] Thomas et al., Space Science Reviews, 217, 25, 2021. [4] Hussmann et al, CEAS Space Journal 11, 381-390, 2019. [5] Michel et al., Planetary Science Journal, 3:160, 2022.
How to cite: Affatato, V., Althaus, C., Binger, J., Grott, M., Hussmann, H., Hüttig, C., Lingenauber, K., Potin, S., Saathof, R., and Stark, A.: Optical Design of a Miniaturized Laser Altimeter Implementing Single-Photon Counting Detection for Topographic Mapping using Small Satellites, Europlanet Science Congress 2024, Berlin, Germany, 8–13 Sep 2024, EPSC2024-474, https://doi.org/10.5194/epsc2024-474, 2024.
The Compact Lunar Neutron and Gamma-ray Spectrometer (CLUNGAS) is targeting planetary science and in-situ resource utilisation (ISRU) such as mapping sub-surface water on celestial bodies. CLUNGAS is the primary payload in the small lunar mission proposal for Selene’s Explorer for Roughness, Regolith, Resources, Neutrons and Elements (SER3NE).
SER3NE is a small satellite mission performing gamma-ray and neutron spectroscopy, hyperspectral near-infrared spectroscopy, and laser altimetry, roughness, and albedo observations at unprecedented spectral and ground resolution. The aim is to: characterise the lunar surface to unravel its volatile origin and delivery processes, uncover the geological processes that shaped the Moon to prospect lunar resources for ISRU at future landing sites, estimate the neutron lifetime and the orbital evolution of the Earth-Moon system.
In this paper, we present the preliminary design of CLUNGAS tailored for the SER3NE mission, as well as environmental analysis to determine design parameters for the detector.
CLUNGAS is a compact solid-state detector capable of gamma-ray imaging in the range of 30 keV - 8 MeV with an energy resolution of < 4% FWHM at 662 keV as well as the detection of thermal to epithermal neutrons The primary detecting material is a CLLBC (Cs2LiLa(Br,Cl)6:Ce) dual-mode scintillator capable of both high-resolution gamma-ray spectroscopy and thermal to epi-thermal neutron detection. The preliminary detector layout aims for a 4x4 chessboard design where CLLBC is accompanied by lanthanum bromide (LaBr3), as depicted in figure 3, for improved gamma-ray – neutron distinction.
The detector core scintillators are surrounded by plastic scintillators serving as an anti-coincidence shield against highly charged particles. Mitigation techniques for secondary neutrons generated by impact of Galactic Cosmic Rays (GCR) and Solar Energetic Particles (SEP) on the spacecraft will be evaluated as well.
The detector readout is based on the ROSSPAD module provided by IDEAS, hosting the low power space-grade detector read-out integrated circuit (ROIC) SIPHRA coupled to a number of Si photomultipliers (SiPM). SIPHRA is radiation tested and has flight heritage.
The power consumption will be below 4W per module and the instrument weight below per unit, with the scintillator volume and shielding requirements as the main design driver. The proposed instrument consists of two units, one covered with few mm thick Gd shield, enabling the distinct detection of both thermal and epithermal neutrons. It will fit into 2U, with the electronics being compatible with the PC104 format. We plan to keep the form factor also for other missions and demonstrations on other platforms.
Using existing data, such as Apollo and LP, we can estimate the expected lunar radiation environment. This is vital to tune the detector design and estimate the background noise such as secondaries generated by the spacecraft itself. The use of Geant4 allows us to track the creation of secondaries (e.g., protons, electrons/positrons, Kaons, muons, pions, and gamma-rays) and their transport in the shallow planetary sub-surfaces. Figure 1 shows the expected neutron flux generated in different lunar soil types due to Galactic cosmic ray impact. The ratio between thermal and epi-thermal neutrons count rates allows for the identification of different sub-surface elements, such as hydrogen content and thus the potential amount of water in the soil. The gamma-ray leakage flux for different lunar soils is presented in figure 2. With the gamma-ray spectrum the composition of the soil can be characterised.
Further, we simulated the response of the preliminary detector design. Figure 3 shows the simulated design in Geant4, with the green light symbolising scintillation photons. Figure 4 visualises the gamma-ray response for 662 keV (Cs-137) stimulation. In the summed energy spectrum over all scintillators, the 662 keV peak, as well as Compton radiation, clearly is visible. Figure 5 demonstrates the detection of neutrons in CLLBC scintillators by the peak of gamma-equivalent energy at 3.2 MeV, as well as effect of the thermal neutron shield (here Cd). Detected neutrons in the shielded CLLBC scintillators are reduced by almost 50%, which is equivalent to the expected number of thermal neutrons in the non-shielded peak.
First measurements with an engineering model, based on the terrestrial version of the ROSSPAD module provided promising results. We tested the read-out electronics and CLLBC scintillator with various gamma sources, such as Cs-137 and Co-60, as well as neutron source AmBe and were able to demonstrate the spectroscopy capabilities within the targeted specifications, as well as the detection of thermal and epithermal neutrons.
In conclusion, we are developing a hybrid gamma-ray and neutron detector for lunar ISRU onboard the SER3NE mission. Detector simulations and laboratory demonstrations show the capability of neutron and gamma-ray spectroscopy. Environmental simulations support the detector design and prepare for future science data analysis.
Next steps are fine tuning and verification of the detector simulation model, extended measurement campaigns of an engineering model, as well as adaption of the electronics design to accommodate spacecraft interfaces.
Figure 1 : Neutron flux spectra for the following lunar compositions: A11, A17, FAN, L20 and L24. We have used the following Geant4 Physics libraries: 1) QGSP, and 2) BERT. The full neutron energy range from thermal to fast neutron is shown.
Figure 2: Gamma-rays flux spectra for the following lunar compositions: A11, A17, FAN, L20 and L24. We also have used the following Geant 4 Physics libraries: 1) QGSP, and 2) BERT. The different spectra present arbitrary normalisation.
Figure 3 Simulated detector design in Geant4. 2 units with a 4x4 chessboard with CLLBC and LaBr scintillators, one unit covered in 0.5mm Cd foil. Inter-scintillator shielding with 0.5mm Al.
Figure 4: Gamma-ray detection capability with simulated detector design in Geant4. Left: Gamma equivalent energy spectrum for each scintillator for gamma source. Right: summed energy spectra over all scintillators, Compton radiation and expected peak at 662keV.
Figure 5: Neutron detection capability with simulated detector design in Geant4. Left: summed energy spectra for both scintillators. Right: comparison between the Cd covered and the uncovered unit.
How to cite: Kohfeldt, A., Al Jebali, R., and Alves Teodoro, L. F.: A Compact Lunar Neutron and Gamma-Ray Spectrometer on Board SER3NE mission, Europlanet Science Congress 2024, Berlin, Germany, 8–13 Sep 2024, EPSC2024-580, https://doi.org/10.5194/epsc2024-580, 2024.
Introduction. The future exploration of Europa and other Ocean Worlds may involve the direct in situ access and characterization of ice crusts and subsurface liquid water using autonomous melt probes, called Cryobots. The Ocean Worlds Reconnaissance and Characterization of Astrobiological Analogs (ORCAA) project is a multi-institution effort, funded under NASA’s Planetary Science and Technology Through Analog Research (PSTAR) program. ORCAA aims to advance our understanding of cryosphere environments on Earth, enabled by planetary subsurface exploration technologies, while envisioning science operations for future icy subsurface access missions.
Objectives. Our overall objectives include the demonstration of subglacial lake access by a terrestrial Cryobot through two field work campaigns. We aim at sampling and analysis borehole melt and subglacial water to understand the evolution of icy habitable environments and the life they host. Through this work, we also aim to elucidate the importance of the hydrological connectivity that may permit nutrient migration and the establishment of habitable or inhabited niches in planetary ice shells. Uniting these science and technology demonstration objectives, we will demonstrate a science concept of operations by additionally simulating command cycles for a subsurface access science mission at Europa with a remote planetary Science Team. While no terrestrial glacier is a perfect physical, chemical, or biological analog to Europa, the Juneau Icefield offers diverse glaciated systems in which to study glacier microbiomes, hydrology and concept operations around melt probe deployment and icy sample handling (Fig. 1).
Figure 1: Comparison between Europa and the Juneau Icefield, AK, where the analogue Cryobot mission is conducted.
Methods. The ORCAA project comprises two field campaigns, in summers 2023 and 2025, on the Juneau Icefield. During the July 2023 field campaign, we successfully tested separately each aspect of the project. The camp was set up nearby supraglacial lakes (Fig. 2), which play an important part in glacier microbiomes. Geophysical reconnaissance was conducted using radar, and used to measure the glacier bedrock depth at the drilling site. We also conducted a full assessment of the 2025 drilling site, identifying the structure and depth of the subglacial lake we will access. The drilling was conducted using a hot water drill (Fig. 2), which will be replaced by a terrestrial Cryobot from the University of Washington, in 2025. Water sampling and analysis was conducted in several supraglacial lakes around camp, as well as in the borehole. Mission control and decision making was conducted in real time in 2023. Finally, we handled and documented logistics associated with field work in icy environments.
Figure 2: Right: Supraglacial lakes and science camp on the Juneau Icefield, AK. Left: Hot water drill used to reach the glacier bedrock.
Building upon this acquired knowledge, we will conduct a comprehensive Europa-analogue subglacial water access and characterization mission with the follow up 2025 field campaign. The terrestrial Cryobot and autonomous spooling and deployment system will handle melting and descent through the ice, as well as water sampling and pumping to the surface. Samples will be analyzed in a clean environment by a team of micro– and astrobiologists. We will simulate complete command cycles and cryobot autonomy to determine how and when scientists and autonomous systems discover indicators of change in the physical environment, habitability, and biological communities, and to characterize the temporal and spatial fidelity requirements of observations.
Acknowledgements. Portions of this research were carried out at the Jet Propulsion Laboratory, California Institute of Technology, under contract with the National Aeronautics and Space Administration (NASA). This work is supported by NASA’s Astrobiology program through Planetary Science and Technology from Analog Research (PSTAR) program (80NM0018F0613).
How to cite: Lesage, E., M. Howell, S., Campbell, S., Mikucki, J., Winebrenner, D., A. Cwik, T., Perl, S., Smith, M., Burnett, J., Clance, J., Clavette, R., Garner, M., Haq, S., Javier, B., and Holmes, J. D.: Onward to Planetary Ocean Access: The ORCAA Cryobot Analogue Mission to the Juneau Icefield, Europlanet Science Congress 2024, Berlin, Germany, 8–13 Sep 2024, EPSC2024-1283, https://doi.org/10.5194/epsc2024-1283, 2024.
MIST-A is the Middle-Wave Infrared (MWIR) Imaging Spectrometer that will be launched in 2028 aboard the Emirates Mission to the Asteroid belt (EMA). The mission will perform six fly-bys during its journey before reaching 269 Justitia in 2034. The main aim of the project is to study the asteroids' origin and evolution, focusing on the examination of the water, organic matter and minerals stored in them and their possible use as resources for future space exploration missions.
MIST-A operates in the 2-5 µm spectral range with a 9 nm/band sampling and its main objective is the identification of the asteroids' surface composition and thermo-physical properties. MIST-A's performances are specifically optimal to detect diagnostic spectral features of primitive asteroids, typically located in the 2.7-4.0 µm range, indicating the presence of hydrated minerals, organic matter, salts and carbonates.
The calibration process is a fundamental step to ensure the reliability of the measurements performed by the instrument.
MIST-A is composed by two units: the Optical Head (OH) and the Electronics Unit (EU).
The design of the OH was inherited from the JIRAM (Jovian IR Auroral Mapper) instrument. It consists of a modified Schmidt telescope equipped with a flat mirror on a 1-axis steerable mechanism at its entrance and joined to the entrance slit of a Littrow spectrometer. The Hybrid-Thinned HgCdTe photodetector of the spectrometer is housed in a thermomechanical structure that also accommodates order-sorting filters and a coldshield and that maintains a temperature < 90 K through an active cryocooler. The rest of the OH operates at a cryogenic temperature of 135 K, reached by means of a passive radiator. The OH is mounted on isostatic legs that thermally isolate it from the spacecraft and its external walls are covered by Multi-Layer Insulation (MLI).
The OH also houses the Internal Calibration Unit (ICU) which is mounted into the telescope's entrance baffle and consists of a flat diffuser with a golden coating illuminated by two IR emitters. A polystyrene filter is placed in front of these sources so that its absorption bands can be used as reference for checking the spectral response in flight.
The EU includes the power converter and distribution unit, the proximity electronics, the scan mirror drive and the CPU board, which contains the command and process control and the data compression software.
MIST-A's calibration campaign will cover the spectral, geometric, spatial and radiometric calibration processes and the characterization of the internal calibration sources.
During the spectral calibration we will characterize the instrument’s spectral responsivity and resolution and its full operative spectral range. The instrument’s boresight alignment, its field of view (FOV) and instantaneous field of view (IFOV) will be tested in the geometric calibration along with the scan mirror’s performances. To define the uniformity of the instrument’s response we will then perform flat-field measurements which additionally will help us check the presence of a vignetting effect. Finally, the radiometric calibration procedure will cover the characterization of the instrument's responsivity and Noise-Equivalent Spectral Radiance.
All these tests will be performed at the nominal OH and detector temperatures. Further measurements taken at warmer and colder temperatures will be used to characterize any observed deviation.
The laboratory setup currently under development in INAF-IAPS, Rome, for the calibration of MIST-A is an evolution of the configuration previously used for the JIRAM instrument.
Fig. Laboratory setup used for the calibration campaign of the MIST-A instrument.
MIST-A's Optical Head and Electronics Unit are housed inside a thermo-vacuum chamber (TVC) which maintains their operating temperatures of 130-140 K. The EU is connected to the Power Supply placed outside the TVC.
MIST-A is controlled and monitored by means of the Electrical Ground Support Equipment (EGSE) with relative Master PC, which is electrically connected to the EU. The Optical Ground Support Equipment (OGSE) commands the devices mounted on the optical bench (e.g. folding mirror actuators, Black Body controller, monochromator) via Labview S/W and it can be used both as a standalone system for the characterization and tests of the optical bench, or coupled as a slave to the EGSE to control the bench through commands given by the Master. The EGSE is also connected to the Data Processing PC which receives and elaborates data from MIST-A and the optical bench.
An off-axis collimator is placed on the optical bench with its folding mirror and target holder structure which can support different targets and sources according to calibration requirements. The bench is mounted on a hexapod, operated through the OGSE, which can be used to traslate the optical beam within MIST-A's FOV while maintaining the folding mirror's alignment to the instrument's boresight.
The optical bench also houses the monochromator that will be used during the spatial and geometric calibration. This instrument is equipped with a IR source and its exit slit is coupled through a IR fiber bundle with the target holder on the collimator's focal plane. The configuration chosen for the radiometric calibration involves instead a black body placed inside the TVC and controlled by the OGSE. A movable folding mirror, commanded through the EGSE, redirects the black body's signal on MIST-A's entrance pupil.
Our team at INAF-IAPS is currently working on the integration and testing of the described optical bench setup. We will successively focus on the definition of the Calibration Plan and OGSE User Manual, which will include information on the calibration of the optical bench devices, before proceeding with the integration of the OGSE.
How to cite: Cencia, C. and the Team MIST-A: The calibration setup for the MIST-A IR spectrometer aboard the EMA mission, Europlanet Science Congress 2024, Berlin, Germany, 8–13 Sep 2024, EPSC2024-452, https://doi.org/10.5194/epsc2024-452, 2024.
The Mission Analysis and Payload Planning System (MAPPS) has evolved over the past two decades to meet the complex and dynamic planning requirements of various ESA planetary missions [1][2]. Originally designed solely for visualizing Mars Express experiment coverage on Mars, MAPPS now offers science planning capabilities for several missions, including Mars Express, ExoMars TGO (Trace Gas Orbiter), BepiColombo, Solar Orbiter, Juice (JUpiter ICy moons Explorer), and EnVision.
These missions present diverse and sometimes conflicting planning needs, such as contrasting power, data volume, or thermal requirements. For instance, Juice requires considerations for long-distance communication and thermal stability in cold outer regions, while BepiColombo prioritizes managing thermal loads and power generation due to its proximity to the Sun.
The primary tool available to the Science Operations Centres (SOC) in ESAC during the science planning process is MAPPS, that has been adapted to these complexities by becoming highly configurable and continuously updating to meet the unique demands of each mission. It operates by consolidating observation requests from the instrument teams, merging them into a comprehensive plan, and extensively simulating and modeling instruments to detect conflicts or constraint violations. The validated plan results in the generation of a multi-instrument operations timeline for uplink to the spacecraft's Mission Operations Centre (MOC).
Recently, MAPPS has included support for multi-body and constellation coordination, crucial for missions like Juice, which orbits Jupiter and its moons for extended periods. Adapting MAPPS for such missions required significant modifications to accommodate observations initially designed for a single central body. Additionally, MAPPS has facilitated collaboration with other Mars missions [3], enabling tasks like radio-science experiments between MEX-TGO [4], communication tests between ESA and NASA orbiters, or regular data relays in support of all surface assets.
Furthermore, MAPPS has been extensively used during the operations studies that were finalized with the adoption of EnVision, by providing essential information about instrument coverage, pointing, or resources consumption.
In summary, MAPPS plays a crucial role in ensuring efficient and conflict-free science planning processes by managing inputs, performing simulations, verifying constraints, generating output files, and executing supporting tasks. Here, we would like to present the main new features and capabilities that make this tool so powerful for science operations and mission analysis.
a) A snapshot of the graphical timeline for one of the Mars Express orbits. It conveys information about the instrument’s modes and actions, along with spacecraft pointing and downlink periods. b) In addition to the graphical timeline, it is possible to validate the planning using 3D visualizations. c) The swaths of the instruments, spacecraft trails, or landmarks can be configured and projected onto 2D maps.
REFERENCES
[1] P. van der Plas et al., MAPPS: a science planning tool supporting the ESA solar system missions, SpaceOps 2016 Conference, AIAA 2016-2512, https://doi.org/10. 2514/6.2016-2512
[2] Cardesin-Moinelo, A., Godfrey, J., Grotheer, E., Blake, R., Damiani, S., Wood, S., Dressler, T., Bruno, M., Johnstone, A., & Muniz, C. (2024). Mars Express: 20 Years of Mission, Science Operations and Data Archiving. Space Science Reviews, 220(2), DOI:10.1007/s11214-024-01059-0
[3] Cardesín-Moinelo et al., First year of coordinated science observations by Mars Express and ExoMars 2016 Trace Gas Orbiter, ICARUS 2021, https://doi.org/10.1016/j.icarus.2020.113707
[4] Cardesín-Moinelo et al. Coordination of Mars Express and Trace Gas Orbiter joint observations of the Martian atmosphere in 2021-2022. EPSC September 2022 Granada
How to cite: Muniz, C., Cardesin-Moinelo, A., Frew, D., de la Fuente, S., Felix, F., Ibarmia, S., Martin, P., Nespoli, F., Mahieux, A., Ashman, M., Lefort, J., and Valles, R. and the ESA Planetary Science Operations Centres: MAPPS: Multi-Mission Science Operations Planning Evolution and Adaptation, Europlanet Science Congress 2024, Berlin, Germany, 8–13 Sep 2024, EPSC2024-252, https://doi.org/10.5194/epsc2024-252, 2024.
Building a full picture of the atmospheric humidity at the Martian surface is crucial for understanding the Martian climate and its implications for the search for life on the red planet. Current humidity sensors on Mars measure the relative humidity (r.h.) with a typical accuracy of a few %r.h [1,2]. This is generally sufficient for determining the volume mixing ratio of water in the atmosphere when the temperature is low enough, however such sensors can run into difficulty when the temperature is higher and the r.h. drops below the sensor measuring threshold, for example during the day or in drier areas. It would therefore be advantageous to deploy a sensor on Mars which is capable of measuring trace levels of water vapour to fill in the gaps where relative humidity sensors reach their limit. A sensor has been developed at the Planetary Analogs Simulation Laboratory at the DLR in Berlin which could provide a compact solution to this problem.
The sensor is based on a coulometric principle and has exhibited excellent accuracy at low humidity levels (range 0.3 to 300ppmV) [3]. It consists of two interlocking comb shaped platinum electrodes printed onto an aluminium oxide substrate using thick film printing technology [4]. The electrodes are spatially separated and an additional glass barrier is deposited between the electrodes to avoid the possibility of a short circuit. The sensor is coated with a patented gel containing P2O5 [5], which is strongly hydrophilic and absorbs water from the surrounding atmosphere. Applying a voltage between the electrodes results in an electrolysis reaction with a current proportional to the water vapour content in the surrounding atmosphere. The reaction can be summarised as:
P2O5 + H2O → 2HPO3
2HPO3 → H2 + 1/2 O2 + P2O5
The sensor layout can be seen in figure 1, as well as the sensor connected to the electrical feedthrough of a CF-25 vacuum flange.
Figure 1. Left: Coulometric sensor developed at the DLR. Right: Coulometric sensor built into a measurement cell CF-25 flange with a Pt100 temperature sensor.
Tests are currently underway in CO2 atmospheres using the sensor calibration facility at PASLAB [6]. The facility has the capability of generating atmospheres of air and CO2 (among other gases), with humidities in a stepless way down to the order of 0.5 ppmV and temperatures down to -70°C. Measurement cells are available for atmospheric pressure and for vacuum measurements.
This contribution will provide details of the sensor design and functioning principle and show the results of tests under CO2 at a range of humidities, temperatures and pressures, providing a solid understanding of the behavior of the sensor in Martian atmospheric conditions.
Acknowledgments: Funding for the results presented in this contribution has been provided through the WIPANO program from the BMWi Deutschland (NORFEUGA-100477081).
References: [1] Hieta, M. et al. (2023) EGUsphere, 1823. [2] Hieta, M. et al. (2022) Planetary and Space Science, 223, 105590. [3] Tiebe, C. et al. (2018) tm - Technisches Messen, vol. 85, no. 12, pp. 746-753 [4] Lorek A. et al. (2010) Proceedings “Aquametry 2010”, Weimar, 289–296. [5] A. Koncz et al., Coulometrischer Feuchtesensor und entsprechendes Verfahren Patent EP2264445 B1 (2017). [6] Lorek et al. (2024) 55th LPSC, 3040.
How to cite: Garland, S., Lorek, A., and Helbert, J.: A trace humidity sensor for measurement of Martian atmospheric humidity, Europlanet Science Congress 2024, Berlin, Germany, 8–13 Sep 2024, EPSC2024-416, https://doi.org/10.5194/epsc2024-416, 2024.
Introduction
Saturn's moon Enceladus gained limelight with the discovery by the Cassini spacecraft of the plumes of ejected gas and ice particles from pronounced linear structures in its South Pole region called “Tiger Stripes". The small (504 km diameter) satellite is believed to have a porous rocky core and an ice shell, separated by a global subsurface saltwater ocean. The tidal heating potentially aids in driving chemical reactions in the moon’s interior which makes it a very promising candidate where the right conditions for life formation may exist. This makes Enceladus a prime target for a future mission. Due to strong gravitational perturbations caused by Saturn, the higher gravitational moments of Enceladus and additional perturbations by the other moons of Saturn, the dynamic environment for artificial satellites around Enceladus is extremely complex. As a consequence, the search for natural stable orbits is far from trivial. A polar orbit is desirable to further investigate the Tiger Stripes region, and for mapping of the global subsurface ocean.
Methodology
To calculate the spacecraft trajectories around Enceladus, all the relevant forces that are acting on the spacecraft are taken into account. We have used a numerical integrator which solves the general equation of motion of the spacecraft. Fig.1 depicts all the relevant perturbations, caused by the Sun, Jupiter, Saturn and its other moons, the higher degrees and order of Enceladus’ and Saturn’s gravity field and solar radiation pressure, which are taken into consideration. The drag experienced due to the plumes is considered negligible (Benedikter et al., 2022).
Fig.1. All the relevant accelerations experienced by the spacecraft in orbit around Enceladus at an altitude of 100 km from the surface.
Orbit Life Time Maps
We searched for suitable orbits in the inertial space by varying orbital parameters such as semi-major axis (350 to 450 km), inclination (40° to 120°), argument of periapsis and longitude of ascending node. The results are presented in the form of orbit life time maps. Fig.2 shows the orbit life time when we vary semi-major axis and inclination with respect to each other.
Fig.2 Orbit Life Time Map: The colour map shows the total life time of various orbits when Semi-Major Axis (320-420 km) and Inclination (77.5°-81.5°) are varied with respect to each other. The grid size of the map is 100 x 100.
Conclusions
Moderately inclined orbits (inclination between 45° and 60°) covering the equatorial and mid-latitude regions of Enceladus were found to be stable from several months up to years. In contrast, the more useful polar mapping orbits were found to be extremely unstable due to the so-called “Kozai mechanism”, due to which a spacecraft would impact the moon’s surface within a few days.
However, an example of a highly inclined orbit was found with inclination of approximately 79°, which had an orbital life time of approximately 13 days. A longer mission in this orbit would require correction maneuvers every few days. This would provide coverage of the tiger stripes region and allow for a global characterization of the ocean.
Orbit Control Strategy
We also determined the delta-v that would be necessary to maintain such an orbit over a mission of several months. This is a preliminary strategy which uses along-track manoeuvres only, i.e., accelerating or decelerating the spacecraft, to maximise the life time of the orbit. The budgets are assessed for an example reference semi-stable orbit solution and various correction manoeuvre frequencies, e.g., every 1, 2, and 4 days.
Also, special attention was paid to satellite formation flying in this orbit to maintain a stable baseline for a distributed radar sounder system (across-track formation of multiple satellites).
References
- Benedikter, A., et al. (2022) Acta Astronautica.
- Hussmann, H., et al. (2012) Planetary and Space Science.
- Ley, W., et al. (2009). Handbook of space technology.
- Enceladus, https://science.nasa.gov/saturn/moons/enceladus/
How to cite: Parihar, T., Hussmann, H., Wickhusen, K., Caritá, G., Stark, A., Oberst, J., Benedikter, A., Rodrigues Silva Filho, E., and Matar, J.: Numerical analysis of polar orbits for future Enceladus missions, Europlanet Science Congress 2024, Berlin, Germany, 8–13 Sep 2024, EPSC2024-707, https://doi.org/10.5194/epsc2024-707, 2024.
Introduction
Brushless DC (BLDC) motors and, in general, Permanent Magnet Synchronous Motors (PMSM), are ubiquitously used in space for robotics as well as other fine motion-control applications [1]. BLDC motors have a high torque density that is practically only limited by overheating. High torques require high currents, which heat up the windings through ohmic effects and, in addition, the increased amplitude of the alternating magnetic field causes the magnets to heat up due to eddy currents. The resulting excessive temperatures can permanently damage the isolation system or demagnetize the permanent magnets beyond the Curie temperature [2]. As robotic motion systems are usually optimized for torque density and accuracy instead of speed, commonly used actuators usually end up having a large gearbox and a small motor. This combination leads to large mass and inertia, and, as a result, sluggish performance. In fact, for a high-performance motion system, one would want to have large motors and small gearboxes instead. In high performance terrestrial robotics, liquid cooled actuators have demonstrated a previously unattained level of performance by avoiding the high temperatures associated with high currents [3] [4]. This study examines the possibility to have the same level of performance also for space robots, as pumped fluid loop cooling systems in general have extensive heritage in space engineering and thermal management [5] [6]. A liquid cooled drive has numerous possible applications in planetary exploration, for example as joints in robotic arms or as wheel or steering drives in rovers. Furthermore, an active thermal controlled actuator can ensure that the motion system remains in its operational temperature range, even in extreme temperature environments, for example permanently shaded regions at the lunar poles, enabling extended mission operations in such places.
Drive Concept
The proposed system consists of a custom BLDC motor with a single-phase liquid cooling system consisting of a motor-fluid heat exchanger, a pump, an accumulator, and a radiator. The motor can be complemented by several sensors and other functional components depending on the application. The thermal management system is flexible, expandable and allows for several actuators to be cooled by one fluid loop. A schematic of this architecture is depicted in the following figure.
Methodology and Results
This work presents a design study for a liquid-cooled robotic actuator by performing coupled multi-physics simulations that can model the drive system in the thermal, electrical, and mechanical domains. This holistic approach allows for an optimal design of the main components considering the complete coupled domain and size the motor and controller according to dynamic requirements (response time, step response, stiffness etc.). The results are used to analyze the influence of active thermal control on the motor performance (torque, power,
short-time as well as continuous) on component level and design an exemplary system for a specific application. The simulation model was verified by sub-scale thermal-vacuum tests of a commercially available BLDC motor.
Conclusion
In summary, active thermal controlled robotic joints have the potential to increase the performance of space robotics systems and can ensure optimal performance in exploration missions with adverse temperature ranges. This study investigates how this proven concept in terrestrial robotics can be applied to in-space applications and presents a simulation-based design methodology to perform preliminary design of such a system.
References
[1] NASA, “Selection of Electric Motors for Aerospace Application,” 01 Feb 1999. [Online]. Available: https://llis.nasa.gov/lesson/893.
[2] X. Wang, B. Li, D. Gerada, K. Huang, I. Stine, S. Worrall and Y. Yan, “A critical review on thermal management technologies for motors in electric cars,” Applied Thermal Engineering, vol. 201, 2022.
[3] T. Zhu, J. Hooks and D. Hong, “Design, Modeling, and Analysis of a Liquid Cooled Proprioceptive Actuator for Legged Robots,” in IEEE/ASME International Conference 2019, Hong Kong, China, 2019.
[4] A. Mazumdar, S. J. Spencer, C. Hobart, M. Kuehl, G. Brunson, N. Coleman and S. P. Buerger, “Improving Robotic Actuator Torque Density and Efficiency Through Enhanced Heat Transfer,” in ASME 2016 Dynamic Systems and Control Conference, Minneapolis, USA, 2016.
[5] D. Gilmore, Spacecraft Thermal Control Handbook, Washington, DC: The Aerospace Corporation, 2002.
[6] A. D. Paris, P. Bhandari and G. Birur, “High Temperature Mechanically Pumped Fluid Loop for Space Applications — Working Fluid Selection,” Journal of Aerospace, vol. 113, pp. 892-898, 2004.
How to cite: Witzel, T., Smolka, A., Plebuch, A., and Guenther, M.: Analysis of liquid-cooled Brushless Motor Actuators for Space Robotics, Europlanet Science Congress 2024, Berlin, Germany, 8–13 Sep 2024, EPSC2024-784, https://doi.org/10.5194/epsc2024-784, 2024.
Introduction. Flux calibration is a key procedure for the full scientific exploitation of the data acquired by planetary remote-sensing cameras. It is necessary to produce high quality, seamless, global and regional color and monochrome mosaics, which are pivotal for the geologic analysis of any planetary surface. It is also fundamental for the quantitative analysis of surface changes, space weathering effects, and for assessing the photometric properties of planetary regoliths. It consists of converting raw data (Digital Numbers, DN) to absolute physical units (W m−2 sr−1 nm−1 or I/F, i.e. the ratio between observed radiance and the radiance of a 100% lambertian reflector with the Sun and camera orthogonal to the observing surface) and requires sources with accurate spectral irradiances or integrated fluxes (i.e, magnitudes). The ESA-Gaia space mission (Gaia Collaboration et al. 2016) is collecting exquisite astrometry and photometry for about two billion stars brighter than G ≃ 20.5 since 2014 (Brown 2021). In the latest data release (Gaia DR3, Gaia Collaboration et al. 2023b), very low resolution spectra (XP spectra hereafter) have been released for the first time, for about 220 million sources (De Angeli et al. 2023). Gaia Collaboration et al. (2023a) demonstrated that remarkably accurate and very precise synthetic photometry can be obtained from flux-calibrated (Montegriffo et al. 2023) XP spectra virtually for any passband whose wavelength range is entirely enclosed within 330 nm≤ λ ≤ 1050 nm. This opens for the first time the possibility to get precise space-based all-sky photometry for a huge number of stars to calibrate the photometric systems of other surveys in the optical range, operating both from space or from the ground. Indeed, synthetic photometry from XP spectra (XPSP hereafter) has been already used for calibration and validation of various photometric surveys (see, e.g., Martin et al. 2023) and is becoming a fundamental photometric reference in the optical domain. In this paper, we assess the potential of using Gaia XPSP for the absolute flux calibration of a planetary remote sensing camera by taking the SIMBIO-SYS instrument on the ESA/JAXA BepiColombo mission as a test case.
Data & Methods. We obtained synthetic photometry in the SIMBIO-SYS photometric system by convolving both XP spectra and well calibrated spectra (hereafter referred as “reference spectra”) from three different libraries of spectrophotometric standard stars through the filter passbands, wavelength dependent detector quantum efficiency, and wavelength dependent mirror reflectivity. We considered the Gaia Spectro Photometric Standard Stars (SPSS; Pancino et al. 2021), The Passband Validation Library (PVL; Pancino et al. 2021), and the latest version of the CALSPEC library (Bohlin et al. 2020). The latter is entirely made of space-based spectra and is generally considered as the best reference stellar flux scale. For each standard, we compare the SIMBIO-SYS synthetic photometry calculated from the GAIA XP with the corresponding synthetic photometry coming from the reference spectra.
Figure 1. Difference between synthetic magnitudes computed from the spectra of the adopted reference set of spectrophotometric standards (magref) and those computed from Gaia XP spectra and corrected for systematics (magcorrXP) for the STC and HRIC filters. Grey triangles are SSP stars, grey square PVL stars and blue circles are CALSPEC stars. The thin horizontal lines enclose the ∆mag = ±0.05 range. The mean and standard deviation of the magnitude difference for the 37 CALSPEC stars having G<10.5 is reported in each panel.
We applied corrections to the original raw XP magnitudes to remove small residual systematics with respect to the CALSPEC reference flux scale. In the following we only refer to these as corrected XP magnitudes.
Results & Discussion. Differences between the synthetic magnitudes computed from the reference spectra of the spectrophotometric standards (magref) and the corresponding corrected XP magnitudes (magcorrXP) as a function of G magnitude and for all the STC and HRIC filters are shown in Fig. 1. We evaluate the average residual (i.e., the accuracy) and their standard deviation (i.e., the precision), on the CALSPEC standards. The mean difference is < 0.001 mag, hence the XP magnitudes in the SIMBIO-SYS system reproduces the CALSPEC photometry with a mean accuracy better than 0.1%. The standard deviation depends on the filter, ranging from 0.004 to 0.010 mag. Hence the corrected XP magnitudes in the SIMBIO-SYS system reproduces the CALSPEC flux scale with a precision < 1%. The calibrated SIMBIO-SYS photometric system is now defined by the XP magnitudes in the SIMBIO-SYS passbands of the standard stars in the SIMBIO-SYS selected fields. To conclude, we show that the availability of high photometric quality Gaia stars can be exploited for providing accurate (below 1%) radiometric calibrations of planetary cameras. For SIMBIO-SYS, in particular, such highly accurate photometry will allow to improve the quality of image mosaics and the quantitative analysis of possible current ongoing surface activity on Mercury, for which SIMBIO-SYS will provide new data after a timespan longer than a decade. Finally, the methodology presented in this paper can in principle be used and evaluated also for other planetary cameras that will explore the Solar System in the next years.
Acknowledgements. MM, MB and PM acknowledge the financial support to activities related to the ESA/mission by the Italian Space Agency (ASI) through contract 2018-24-HH.0 and its addendum 2018-24-HH.1-2022 to the National Institute for Astrophysics (INAF). The study has been supported by the Italian Space Agency (ASI-INAF agreement no. 2020-17-HH.0).
References
Brown, A. G. A. 2021, ARA&A, 59, 59
Gaia Collaboration, Montegriffo, P., Bellazzini, M., et al. 2023a, A&A, 674, A33
Gaia Collaboration, Prusti, T., de Bruijne, J. H. J., et al. 2016, A&A, 595, A1
Gaia Collaboration, Vallenari, A., Brown, A. G. A., et al. 2023b, A&A, 674, A1
De Angeli, F., Weiler, M., Montegriffo, P., et al. 2023, A&A, 674, A2
Montegriffo, P., De Angeli, F., Andrae, R., et al. 2023, A&A, 674, A3
Bohlin, R. C. 2020, in IAU General Assembly, 449–453
Pancino, E., Sanna, N., Altavilla, G., et al. 2021, MNRAS, 503, 3660
Martin, N. F., Starkenburg, E., Yuan, Z., et al. 2023, arXiv e-prints, arXiv:2308.01344
How to cite: Munaretto, G., Cremonese, G., Bellazzini, M., Montegriffo, P., Sordo, R., La Grassa, R., Re, C., Tullo, A., and Messineo, M.: Using Gaia for the flux calibration of planetary cameras: the BepiColombo/SIMBIO-SYS case, Europlanet Science Congress 2024, Berlin, Germany, 8–13 Sep 2024, EPSC2024-804, https://doi.org/10.5194/epsc2024-804, 2024.
Dragliner [1, 2] is an ESA funded study to define the requirements and a preliminary design of a passive Coulomb Drag based deorbit system capable of bringing down LEO spacecrafts in an order of magnitude shorter time than the current regulations for the re-entry time for the spacecraft (25 years), while using significantly less spacecraft resources compared to traditional deorbiting methods like chemical or electric propulsion. Other main requirements for the deorbiting system are low mass and independence from the spacecraft resources. The study is led by Finnish Meteorological Institute, and the other consortium partners are Aurora Propulsion Technology, GRADEL sarl and SpaceR-SnT/University of Luxembourg. The current goal for Technology Readiness Level of the project is 4.
The Dragliner system utilizes a Coulomb Drag microtether, also known as a plasma brake), i.e. tether that utilizes momentum from ionospheric plasma ram flow by electrostatic interaction [3, 4]. In this project, microtether is defined as a tether that does not exceed the mass limit of 200 milligrams per length of 1 meter, which makes it safe to other space assets in the event of a collision (the energy of collision is comparable to micrometeroid impact).
The main mission requirements for the Dragliner are:
- Dragliner shall be able to deorbit a Low Earth Orbit satellite (uncontrolled re-entry) of max. 250 kg mass to 400 km altitude. If the original orbit height is less than 850 km, the deorbit time shall be less than 2 years, with the goal of less than 100 days.
- Dragliner shall be an autonomous, passive bolt-on device that does not use spacecraft resources after it has been triggered to deploy by a ground command. After deployment the host spacecraft does not need to be active.
- The mass fraction of Dragliner to the host spacecraft shall be lower than 5%.
- The reliability of the whole system shall be better than 95%, and the risk of premature deployment less than 0.1%.
Coulomb drag deorbiting time depends on the combination of spacecraft’s mass, orbit height and length of the microtether. To achieve the required deorbiting time for 250 kg spacecraft from 850 km, 10 km long microtether is needed. Currently the practical limit for microtether length is 5 km, so in this case the tether is divided into two separate modules, with 5 km of tether each.
The Main Tether of Dragliner must be electrically conductive to allow electrical charging that enables the Coulomb Drag effect, and tolerant to micrometeoroid impacts. The Dragliner tether is manufactured at Finnish Meteorologial Institute from 4 Aluminum wires of 50 um diameter, mechanically twisted to form a “Hoytether-like” structure [5]. With this configuration, the risk of complete severance of a 5-km tether during 1 year in orbit is low enough to fulfil the 95% reliability requirement of the full Dragliner system.
Figure 1: 4-wire tether from 50 um Al 2024 wire produced by FMI manual tether factory
In addition to kilometers long Main Tether, Dragliner contains also a significantly shorter (tens of meters) conductive Tape Tether that serves as an electron collecting surface for the operation of the Main Tether.
One Dragliner module consists of a Base Unit containing mechanical and electrical interfaces to the spacecraft, and a Remote Unit housing the tethers and the electronics. The size of the Remote Unit is equivalent to 3U CubeSat. During the operational lifetime of the spacecraft, Dragliner remains in stowed configuration, relying on the spacecraft for power and communications via umbilical connectors. When the spacecraft decommissioning starts, the Dragliner system shall be commanded to deploy. All Dragliner functions after the deployment signal has been given will be autonomous.
Figure 2: Physical architecture of Dragliner (not in scale)
In the beginning of the deployment process, as the Remote Unit is released from the Base Unit, the umbilical electrical connectors to the spacecraft are severed, and Dragliner starts to operate completely autonomously, including mechanical deployment of the tethers and harvesting its own energy with solar panels attached to the sides of the Remote Unit. After the deployment (lasting a few hours) is completed, the Deorbit System charges the Main Tether with ∼1 kV high voltage, enabling the Coulomb Drag effect and deorbiting of the spacecraft. The spacecraft does not need to be operational during the deorbiting phase.
The Dragliner project has passed its System Requirements Review and Preliminary Design Review in 2023. The deployment of the Main Tether (scaled-down length) and the Remote Unit, which is one of the critical functions of Dragliner system, is tested with a Dragliner Breadboard model in Zero-G lab of SpaceR-SnT/University of Luxembourg. The lab is is designed to emulate space related scenarios under microgravity conditions [6, 7]. The tests are scheduled for May 2024. In the tests, the Base Unit will be attached to a floating platform and the Remote Unit to a stationary robotic manipulator. When the deployment is executed, the freely floating platform with the Base Unit will start to move away from the Remote Unit simultaneously deploying the Main Tether. The deployment length is limited by the room dimensions.
The Dragliner project is carried out under a programme of and funded by the European Space Agency under the ESA Contract No. 4000138811/22/NL/MM/fm.
[1] Genzer, M. et al. Project DragLiner: harnessing plasma Coulomb drag for satellite deorbiting to keep orbits clean, EGU23-14692
[2] Peitso Pyry et al. Plasma Brake for Deorbiting Telecommunication Satellites, SP2024_paper #490
[3] Janhunen P., Electrostatic plasma brake for deorbiting a satellite, J. Prop. Power, 26, 370-372, 2010.
[4] Janhunen, P., Simulation study of the plasma-brake effect, Ann. Geophys., 32, 1207-1216, 2014.
[5] Toivanen P. et al, Robust Flight Tether for In-Orbit Demonstrations of Coulomb Drag Propulsion, Aerospace 2024, 11(1), 62
[6] M. Olivares-Mendez et al. Zero-g lab: A multi-purpose facility for emulating space operations. Journal of Space Safety Engineering, pages 509 – 521, 2023
[7] B. C. Yalcın et al. Lightweight floating platform for ground-based emulation of on-orbit scenarios. IEEE Access, pages 94575 – 94588, 2023.
How to cite: Genzer, M., Janhunen, P., Yli-opas, P., Peitso, P., Haukka, H., Hieta, M., Laurila, H., Pietikäinen, P., Hallamaa, H., Sinkko, J., Toivanen, P., Polkko, J., Yalcin, B. C., Olivares-Mendez, M., and Macieira, D.: Dragliner – Tether Based System for Passive Spacecraft Deorbiting Using Coulomb Drag, Europlanet Science Congress 2024, Berlin, Germany, 8–13 Sep 2024, EPSC2024-852, https://doi.org/10.5194/epsc2024-852, 2024.
The exploration of planetary surfaces is fundamental to our understanding of planetary evolution, habitability, and the origin of life. The chemical composition is typically amongst the main scientific goals of planetary exploration missions. Current orbital space instruments for probing chemical composition of planetary surfaces such as VIRTIS on Rosetta and NIRS3 on Hayabusa-II rely on classical infrared spectroscopy to measure minerals and organic matter. However, spectrometers for space exploration are more limited in resolution and bandwidth due to constraints in mass, size, power consumption, thermal stability, and space-graded material availability than what can be achieved in a laboratory environment with prepared samples. Beyond these engineering constraints, classical spectroscopy is also intrinsically limited in measuring chemical bonds and not the molecular structure of a given compound.
Alternatively, mass spectrometers such as COSAC on the Philae lander [5], SAM on the Curiosity rover [3], and MOMA on the Exomars rover [1] can deliver outstanding sensitivity, but at the expense of payload mass, instrument complexity, and the destruction of the sample during measurement. As a result of these constraints, these instruments are primarily suited for a select few mission profiles, notably flagships or large-class mission architectures. Furthermore, COSAC, SAM, and MOMA are currently the only space instruments capable of measuring chirality - the proportion of left vs right handed enantiomers in a sample. While studying these enantiomeric ratios provides key data in the context of prebiotic chemistry and understanding the emergence of homochirality on Earth, these instruments can only make a handful of measurements of this kind over their lifetime. Additionally to the limited number of uses of the chiral columns in the gas-chromatographs (GC), this technique is suitable for volatile compounds excluding chiral targets of interest such as amino acids.
In response to these limitations, we present a novel approach to planetary surface spectroscopy, leveraging polarization as an additional observable to complement traditional spectroscopic techniques and mass spectroscopy chirality measurements. Introducing the Chirality Analyzer In-Situ (CHAIS), a near-infrared optical bench (1.5 - 4.0 um) specifically designed to measure Vibrational Circular Dichroism (VCD) for planetary exploration—a technique well-established in terrestrial chemistry and pharmaceutical laboratories. The objective of CHAIS is to validate this approach through measurements in a true-to-life configuration using planetary analogs, with the aim of assessing the sensitivity thresholds necessary for the development of a space-optimized version of the instrument. By analyzing linear and circular polarization signatures as a function of wavelength, CHAIS is anticipated to enable the non-destructive characterisation of the typically featureless felsic primordial crust minerals present on planetary bodies like the Moon and Mars. The addition of polarimetry to CHAIS not only enhances the capability to discern certain molecular structures while measuring chirality, but also takes advantage of the instrument's relative simplicity and the absence of required sample preparation, enabling a higher number of measurements.
Beyond composition analysis, CHAIS, as the name suggests, is built to measure chirality. Our approach would allow for the measurement of enantiomeric excess and homochirality (100\% enantiomeric excess) which paves the way towards a robust biosignature detection tool. Additionally, other existing spectropolarimeters, such as TreePol [4] and LSDpol [2], operate in the visible spectrum using different methodologies, yet similarly demonstrate the use of optics for detecting chirality and biosignatures. Ultimately, this integration of polarimetry into classical infrared spectroscopy will enable the non-destructive characterisation of a wide variety of planetary surfaces and the development of a new type of space instrument optimized for exploring and understanding the origin and formation of biotic matter.
Racemic abiotic samples like plastic films as well as Mars analogs such as gypsum have been measured in transmission as part of the first test cases to validate the bench. Preliminary spectroscopy data is shown in Figure 1. In this work we present initial findings with CHAIS including first light results, first spectroscopy and spectro-polarimetry data from a variety of biotic and abiotic samples, as well as the next steps towards a point-and-shoot space instrument of adequate sensitivity to characterise the chemical composition and chirality of surface minerals.
Acknowledgments
This research is co-funded by the Centre National d’ ́Etudes Spatiales (CNES) and the Aix-Marseille Origines Institute
through the A*MIDEX Program.
References
[1] Fred Goesmann et al. “The Mars Organic Molecule Analyzer (MOMA) instrument: characterization of
organic material in martian sediments”. In: Astrobiology 17.6-7 (2017), pp. 655–685.
[2] Christoph U Keller et al. “Design of the life signature detection polarimeter LSDpol”. In: Space Telescopes
and Instrumentation 2020: Optical, Infrared, and Millimeter Wave. Vol. 11443. International Society for
Optics and Photonics. 2020, 114433R.
[3] Paul R Mahaffy et al. “The sample analysis at Mars investigation and instrument suite”. In: Space Science
Reviews 170 (2012), pp. 401–478.
[4] CH Lucas Patty et al. “Circular spectropolarimetric sensing of chiral photosystems in decaying leaves”.
In: Journal of Quantitative Spectroscopy and Radiative Transfer 189 (2017), pp. 303–311.
[5] Stephan Ulamec, Fred Goesmann, and Uwe Meierhenrich. “Philae Landing on Comet 67P/Churyumov-
Gerasimenko–Planned Chirality Measurements and Ideas for the Future”. In: (2018)
How to cite: Krasteva, M., Carter, J., Madec, F., Challita, Z., Vinogradoff, V., d'Hendecourt, L., Groussin, O., and Brunetto, R.: CHirality Analyzer In-Situ (CHAIS) - A Novel Approach to Planetary Surface Characterisation, Europlanet Science Congress 2024, Berlin, Germany, 8–13 Sep 2024, EPSC2024-881, https://doi.org/10.5194/epsc2024-881, 2024.
The present communication is based on preliminary results within the PRIN 2022 PNRR project entitled “Low-cost, high-safety hydrogen storage into chemically-enhanced clathrate hydrates for energy storage in planetary infrastructures” (Brave New Worlds, CUP D53D2301693000, funded by the European Union – Next Generation EU) led by the University “G. d’Annunzio” of Chieti – Pescara, with the collaboration of University “Aldo Moro” of Bari, and the National Institute for Astrophysics of Catania (INAF) as research units. The aim of the project “Brave New Worlds” is to develop low-cost, high-safety media for the storage of hydrogen, which also have minimal technologic and maintenance requirements. Target storage media will be ideally suitable for space-based infrastructures on near planets (e.g., Mars) or satellites (e.g., Moon), where hydrogen will be produced from planetary water bodies by solar cell-powered electrolysis.
Currently, hydrogen is stored in pressurized cylinders, metal hydrides and similar compounds that require high energy consumption to store and recover H2, or in liquefied form. None of those storage technologies are suitable for planetary infrastructures, because of the high spacecraft payloads needed to carry cylinders, compressors, metallic media and other highly technological devices to be deployed and assembled in situ. Furthermore, the control of the compression of hydrogen into cylinders, or the temperature cycles for the sorption onto metal hydrides, or for liquefaction, require sophisticated, failure-prone control appliances.
The storage media developed in the present work are clathrate hydrates of hydrogen, a class of supramolecular solids consisting of water molecules organized in cage structures that can host one or more gas molecules. These systems represent a safer, technologically simpler, and cheaper alternative for large-scale hydrogen storage than traditional storage methods. Clathrate hydrates are usually formed under conditions of pressures around 5-10 MPa and temperatures of around 250-280 K, or, importantly for the present case, under lower pressures and very low temperatures.
Thus, important features that we are exploiting are the following:
- clathrate hydrates are essentially made up of water, an economical, ecological and safe compound par excellence. Having a potentially infinite life cycle, water is an ideal material for this purpose.
- Water is found on (or just below) planet and satellite surfaces.
- Sun-shaded or deep crater areas of planets and satellites reach temperatures as low as 30 K
- Hydrogen hydrates can form at very low temperatures under mild gas pressures
In the present work, we show how to overcome some critical points of hydrogen storage in clathrates, namely (i) slow capture kinetics, and (ii) low gravimetric content. As for point (i), here we report processes and molecules for improving the kinetics of the process of 1-2 orders of magnitude. The increase of the gravimetric content (point (ii)) has been addressed with the design and test of stabilizers (co-formers) of the hydrate cages, through a combination of rational design, quantum mechanical and molecular dynamics approaches, stochastic methods, and chemical synthesis. Preliminary results reported in the present contribution are: (i) the thermodynamic equilibrium curves of binary hydrogen hydrates in presence of stabilizers (tetrahydrofuran, cyclopentane; Fig. 1); (ii) the kinetics of formation of hydrogen hydrates under water-in-oil or oil-in-water emulsion systems (Fig. 2); and (iii) Raman spectroscopy data showing the presence of hydrogen within the small hydrate cages (512 dodecahedra) and possibly also within the large cages (51262 polyhedra) of sII crystal structures (Fig. 3).
The final goal will be to develop a hydrogen storage medium with a gravimetric H2 content around 4 wt%, which is demonstrably competitive with current top technologies at a fraction of the technological level and economic cost.
Fig. 1: P/T curve for THF/CP/hydrogen hydrate formation under different pressures
Figure 2: Kinetic P/T curves for the formation of THF/CP/H2 hydrate
Figure 3: Raman spectrum of THF/CP/H2 hydrate at 70 bar
Bibliography
[1] Di Profio P, Arca S, Rossi F, Filipponi M. Int J Hydrogen Energy 2009;34:9173–80.
[2] Di Profio P, Arca S, Germani R, Savelli G. J Fuel Cell Sci Technol 2007;4:49–55.
[3] Di Profio P, Canale V, Germani R, Arca S, Fontana A. J Colloid Interface Sci 2018;516:224–31.
How to cite: Di Profio, P., Ciulla, M., Siani, G., Barbacane, N., Wolicki, R. D., Di Giacomo, S., Re, N., Marrone, A., Paciotti, R., Longo, S., Urso, R., Scire Scappuzzo, C., Palumbo, M. E., and Baratta, G.: Investigation of Hydrogen Clathrate Hydrates as Energy Storage Media for Planetary Installations, Europlanet Science Congress 2024, Berlin, Germany, 8–13 Sep 2024, EPSC2024-1169, https://doi.org/10.5194/epsc2024-1169, 2024.