MITM17 | Space Safety

MITM17

Space Safety

Co-organized by SB

Convener: Ari-Matti Harri | Co-conveners: Thomas Statler, Karri Muinonen, Michael Küppers

Orals FRI-OB3

| Fri, 12 Sep, 11:00–12:24 (EEST)

Room Neptune (rooms 22+23)

Posters THU-POS

| Attendance Thu, 11 Sep, 18:00–19:30 (EEST) | Display Thu, 11 Sep, 08:30–19:30

Finlandia Hall foyer, F125–129

Space safety refers to the sustainable use of space, encompassing space-based applications, commerce, science and exploration and focuses on the impacts of space debris, space weather, and planetary defense. It strives towards protecting space infrastructure and human spaceflight missions from orbital debris and from harmful effects of space weather, as well as safeguarding Earth's societies from debris re-entering the atmosphere. Planetary defense involves detecting, tracking, and understanding potentially hazardous near-Earth objects (NEOs), asteroids and comets capable of impacting the Earth.

This session invites presentations on scientific methods, technologies, and missions for surveillance, detection, tracking and characterization of orbital objects around the Earth, for understanding and forecasting space weather and its impact on the space and aviation infrastructure and other technical systems of our society, as well as on asteroids and comets that could threaten the Earth. We also welcome presentations on methodologies, systems and missions for actively mitigating the risks posed by re-entering orbital debris and approaching NEOs to human society.

Session assets

Orals: Fri, 12 Sep, 11:00–12:30 | Room Neptune (rooms 22+23)

Chairpersons: Ari-Matti Harri, Michael Küppers, Thomas Statler

11:00–11:12

EPSC-DPS2025-1733

On-site presentation

Avoidance of Satellite Damage by Collisions in Space with theUse of VLF Plasma Waves in Space to Detect the Location of Harmful Space Debris

Paul Bernhardt, Andrew Howarth, and Bengt Elliason

The number of satellites launched into low Earth orbit (LEO) is increasing at an exponential rate. Launches support deployment of multi-satellite constellations for many applications. Experiments with electric field sensors on Swarm-E and with HAARP in Alaska have been conducted to (a) better locate the positions of satellites and space debris for prevention of collisions
Currently, there are about 27,000 known space objects and over 100 million of unknown pieces of space debris. Collision avoidance requires precise knowledge of the positions for all space objects. New techniques are being developed to detect the small, < 10 cm, currently “invisible” objects by the plasma waves they generate in space. The basis for this technique is that all space objects in orbit around the Earth (1) pass through a magnetized plasma, (2) become electrically charged, and thus (3) produce an electric current that excites electrostatic lower hybrid waves. Orbital kinetic energy is the power source for the lower hybrid waves. When the debris moves through field aligned irregularities (FAIs), the lower hybrid waves are converted into whistler, and compressional Alfven waves. Such whistlers propagate undamped at around 9000 km/s from the source regions and can be detected at ranges of several earth-radii.
This space debris detection process has been tested with the Canadian Swarm-E satellite using the Radio Receiver Instrument (RRI) that measures electric fields in the 10 Hz to 30 kHz frequency range. The RRI makes measurements of plasma waves when near known space objects such as Starlink. An example of these data typically shows an enhancement in electric fields in a band below the local value of lower hybrid frequency. This spacecraft signature are whistler waves in a band between the ion cyclotron and lower hybrid frequencies with an upper frequency cutoff not observed for natural whistler waves. These signals can be used to both detect and track unknown space objects by computing their propagation direction and establishing an orbital state vector of the object. The goal of these measurements is to collect a catalog of space debris with sizes less than 10 cm for collision avoidance.
References
P.A. Bernhardt, M. K. Griffin, W. C. Bougas, A. D. Howarth, H. G. James, C. L. Siefring, and S. J. Briczinski, (2020) Satellite Observations of Strong Plasma Wave Emissions with Frequency Shifts Induced by an Engine Burn from the Cygnus Spacecraft, Radio Science, 56.
P.A. Bernhardt, R.L Scott, A Howarth, George. J. Morales (2023) Observations of Plasma Waves Generated by Charged Space Objects, Phys. Plasmas 30, 092106, https://doi.org/10.1063/5.0155454
Eliasson, B., & Bernhardt, P. A. (2025). The generation of whistler, lower hybrid and magnetosonic waves by satellites passing through ionospheric magnetic field aligned irregularities. Physics of Plasmas, 32(1), Article 012103. https://doi.org/10.1063/5.0225399

How to cite: Bernhardt, P., Howarth, A., and Elliason, B.: Avoidance of Satellite Damage by Collisions in Space with theUse of VLF Plasma Waves in Space to Detect the Location of Harmful Space Debris, EPSC-DPS Joint Meeting 2025, Helsinki, Finland, 7–13 Sep 2025, EPSC-DPS2025-1733, https://doi.org/10.5194/epsc-dps2025-1733, 2025.

11:12–11:24

EPSC-DPS2025-1698

ECP

On-site presentation

Mission concept of CENSSAT-1, a CubeSat for multimodal monitoring of the space environment

Shayla Viet and the CENSSAT-1 team

It is well recognized that knowledge of the space environment is important to society through our reliance on satellites. Navigation and communication systems provided by satellites are vulnerable to solar storms. Electrical power grid networks may also be affected by space weather. The amount of space debris is rapidly growing, giving rise to challenges in spacecraft operations and space-based services. Tracking space debris and monitoring space weather events that may affect human infrastructure is therefore essential.

The Centre for Space Sensors and Systems (CENSSS) at the University of Oslo will address these challenges in space situational awareness by designing, developing, testing, and operating a CubeSat mission, CENSSAT-1, with collaborations from industry and university partners. The primary mission objectives are to monitor space weather in low Earth orbit and to study the response of the Earth’s upper and middle atmosphere to space weather events.

The payload consists of a Langmuir probe for plasma measurements, a multifunctional particle detector, a space debris radar, and a camera to map aurora and airglow as well as ozone variability. The camera will record images in four spectral bands corresponding to atmospheric emission and absorption lines. The observation strategy of the instruments will enable simultaneous measurements of the space radiation environment in the orbit of the satellite, and facilitate the nowcasting of solar storms. The observation principle demonstrated on CENSSAT-1 could also be applied to other planetary missions, for example investigating auroral processes on Mars where forecasting the arrival of solar storms is more difficult.

There is currently a lack of data on small-sized space debris, as detection and tracking from the ground is difficult. While space debris of sub-millimeter size can be detected by in-orbit impact sensors, flux data for objects in the millimeter-size range is limited. The space radar on CENSSAT-1 will therefore aim to detect millimetre-sized space debris in the vicinity of its orbit. CENSSAT-1 will also explore the possibility of measuring neutron lifetime based on directional neutron flux observations, and conduct several optimal control and drag experiments. The launch of CENSSAT-1 is currently estimated to be at the end of 2027.

In addition to tackling key issues in space safety, the CENSSAT-1 mission also serves to train and qualify personnel on MsC, PhD and postdoctoral levels, by providing hands-on experience in satellite mission design and operations. Thus, CENSSAT-1 will also fulfill the needs of the space industry and research institutions.

How to cite: Viet, S. and the CENSSAT-1 team: Mission concept of CENSSAT-1, a CubeSat for multimodal monitoring of the space environment, EPSC-DPS Joint Meeting 2025, Helsinki, Finland, 7–13 Sep 2025, EPSC-DPS2025-1698, https://doi.org/10.5194/epsc-dps2025-1698, 2025.

11:24–11:36

EPSC-DPS2025-547

On-site presentation

Applying the Plasma Brake for Active Debris Removal: revolutionarily high downmass/upmass ratio

Pekka Janhunen, Olli Knuuttila, Maria Genzer, Petri Toivanen, Leo Nyman, Harri Haukka, and Antti Kestilä

The Plasma Brake is a propellantless method for deorbiting objects from Low Earth Orbit (LEO). It is based on charging a long and thin metallic tether (a microtether) to high negative voltage (1 kV) so that the electric field around the tether scatters ionospheric ions of the satellite's ram flow, thus braking the orbital motion. The microtether is made of three or four max 50 micrometre thin aluminium wires. It weighs only 23 grams per kilometre and does not pose a risk to other space assets due to its thinness. The tether is deployed initially to about 100 m length by a spring, and then Earth's gravity gradient affecting the reel unit itself deploys the full 5 km length.

A 4 kg Plasma Brake module has been developed in ESA Dragliner project at TRL 4-5, providing nominally 0.4 mN deorbiting thrust, which is enough to deorbit a 250 kg object from 700 km orbit in 2 years. The deorbiting thrust depends on the prevailing ionospheric plasma density and mean ion mass, but the quoted numbers represent average solar cycle conditions and typical 600-800 km starting altitudes. Up to two modules can be installed in a satellite for its end-of-life deorbiting, one deploying its tether downward and the other one upward. In general the Plasma Brake has to lower the altitude only to about 400 km, after which atmospheric drag on the tether and the object itself complete the deorbiting.

Here we consider application of Plasma Braket technology to Active Debris Removal, i.e., deorbiting of old, uncollaborative debris objects. We have a mothership that has a Hall thruster to make a tour of several debris objects. The mothership attaches a Plasma Brake module to each debris object. The attachment is by a mechanical device such as a flat tape ribbon loop which is opened to e.g. 5 m diameter from a reel by an electric motor. The loop is then passed around the object and slowly tightened by the motor, after which the mothership detaches the Plasma Brake unit and departs to rendezvous the next debris object. Meanwhile the Plasma Brake module opens its tether, first by a spring to about 100 m length and then by Earth's gravity gradient to the full 5 km length. Then it turns on the 1 kV negative voltage on the tether, for which it needs about 1 W of power which it gets from its own cubesat-like surface-mounted solar panels.

Because a 4 kg Plasma Brake module can deorbit a 250 kg mass, if the modules comprise 30% of the launch mass of the mothership, overall we reach a downmass/upmass ratio of nearly 20. This is revolutionary because the current state of the art (chemical propulsion or electric propulsion) yield downmass/upmass ratios not much larger than unity.

How to cite: Janhunen, P., Knuuttila, O., Genzer, M., Toivanen, P., Nyman, L., Haukka, H., and Kestilä, A.: Applying the Plasma Brake for Active Debris Removal: revolutionarily high downmass/upmass ratio, EPSC-DPS Joint Meeting 2025, Helsinki, Finland, 7–13 Sep 2025, EPSC-DPS2025-547, https://doi.org/10.5194/epsc-dps2025-547, 2025.

11:36–11:48

EPSC-DPS2025-1835

On-site presentation

Charon: a system capable of deorbiting satellites with no platform dependencies

Tuomas Simula, Perttu Yli-Opas, Maria Genzer, Leo Nyman, Pekka Janhunen, Jaakko Kaartinen, Aleksandr Gorbachev, Antti Kestilä, Matija Herceg, Leonardo Ghizoni, Kari Mäkiniemi, and Kevin Vainio

The increasing number of satellites launched constitutes a risk in the form of growing space debris, especially in low Earth orbit. With more objects in orbit, the risk of collisions and creation of additional space debris grows significantly. Regulations to address this state that new satellites need to be capable of deorbiting within 5 years of end of mission. This requirement presents a major challenge for nano- and microsatellites, such as CubeSats, which tend to have higher failure rates than larger satellites. In the case of a platform failure, traditional deorbiting systems that are part of the satellite’s AOCS and rely on the platform being functional, cannot fulfill the requirement. As a result, the satellite becomes non-compliant with deorbit regulations and turns into space debris. Additionally, CubeSats are limited in terms of available volume, mass, and power. This makes it difficult to integrate conventional deorbiting technologies such as propulsion systems. To be effective, a deorbiting solution for these satellites must be small, lightweight, affordable, and capable of functioning independently from the satellite platform.Charon is a module designed to fulfill this need: a fully autonomous deorbiting module based on Plasma Brake technology, and in a standard CubeSat form factor of a 0.5U module and an attached TunaCan volume for the plasma brake itself. The system is able to function independently of the host satellite, enabling it to detect a platform failure and deploy its Plasma Brake without any input from the satellite itself. The working principle of the Plasma Brake is based on a long (typically 200 m) microtether charged to a high negative voltage of -1000 V. The charged tether interacts with the plasma in the upper atmosphere, causing a drag force through the Coulomb Drag effect. The Plasma Brake is thus able to provide thrust without any propellant, requiring only constant electric power in the order of 0.1 W from its own solar panels. While the level of thrust is very small, it is sufficient to deorbit a typical CubeSat in a few years.Charon is designed to activate automatically in case of mission failure, also in the case that no communications can be established with a dead-on-arrival host satellite. This is facilitated by a watchdog timer on the module that is periodically reset by the spacecraft; in case of two missed resets of the timer, the module determines that the platform has failed, and commences automatic deorbiting. To be as autonomous and platform independent as possible, the module is equipped with an independent power source and control logic, a small thruster and simple attitude determination system for deployment of the Plasma Brake microtether, and the Plasma Brake itself to deorbit the satellite. The target reliability of above 95% independent of the host satellite is reached by keeping the module architecture simple, and by redundancy of key components. In a standard 0.5U CubeSat form factor and with minimal interfacing to the host satellite, the compact and modular design allows Charon to be integrated into typical CubeSat platforms with minimal impact on payload or mission architecture. By providing a self-contained, platform-independent deorbiting solution, Charon offers a practical path to ensuring regulatory compliance and reducing the long-term risks of space debris for the growing number of small satellites in orbit.

How to cite: Simula, T., Yli-Opas, P., Genzer, M., Nyman, L., Janhunen, P., Kaartinen, J., Gorbachev, A., Kestilä, A., Herceg, M., Ghizoni, L., Mäkiniemi, K., and Vainio, K.: Charon: a system capable of deorbiting satellites with no platform dependencies, EPSC-DPS Joint Meeting 2025, Helsinki, Finland, 7–13 Sep 2025, EPSC-DPS2025-1835, https://doi.org/10.5194/epsc-dps2025-1835, 2025.

11:48–12:00

EPSC-DPS2025-1403

ECP

On-site presentation

Detecting near-Earth objects using real-time synthetic tracking surveys

Malin Stanescu, Marcel Popescu, Lucian Curelaru, Ovidiu Vaduvescu, Daniel Berteșteanu, and Marian Predatu

Introduction

The detection of near-Earth objects (NEOs) is a critical task for planetary defense and of significant importance for astronomy and astronautics. While existing surveys have successfully identified practically all large (>1km) NEOs and have successfully progressed for those >100 m, smaller objects are still challenging to detect, being overwhelmingly detected only during close flybys.

Traditional detection methods, such as the "blink" technique, rely on large aperture telescopes to detect smaller, fainter, objects. However, doing so is increasingly expensive, with the costs becoming prohibitive. In comparison, Synthetic Tracking (ST) is a detection technique [1][2] that allows the use of (many) smaller telescopes with longer integration times, where the detections can be significantly under the noise floor of individual images, by combining the images to increase the signal to noise ratio across all possible trajectories of a faint potential moving object. Despite the advantages of Synthetic Tracking, due to the perceived need for vastly more computational power, the backbone of current surveys is still the blink method. This factor is particularly prominent for NEOs, whose apparent velocities are much higher than for other types of minor planets, leading to a large number of potential object trajectories.

Methods

We present Synthetic Tracking on Umbrella (STU), a moving object detection software which leverages the power of modern GPUs to detect NEOs in real time. STU is developed on top of the Umbrella2 library[3], which we have previously developed to implement a blink detection pipeline, including both the detection algorithm and auxiliary functionality necessary for a complete detection processing pipeline. Thanks to innovative search strategies, efficient use of hardware and a multi-step candidate rejection process, STU can perform real time detection of fast-moving Near Earth Asteroids, even on large, multi-CCD instruments. STU can do so using modest hardware resources, such as a single off-the-shelf GPU, while being robust to outlier noise.

Results

Previously, we successfully demonstrated the functionality of STU through an observation archive spanning more than 100000 images from various telescopes (Stanescu et al., 2025, accepted), using images obtained in various conditions (variable seeing, detectors with different noise levels, pixel scales, field of view, objects with different apparent magnitudes, and moving rates). Furthermore, we have further tested STU in realistic survey conditions using the 2.54m Isaac Newton Telescope (INT), the 1.6m Korean Astronomy and Space Science Institute (KASI) telescopes[4].

For the INT, real-time capability was demonstrated with the Wide Field Camera (WFC) on the 2.54m Isaac Newton Telescope (INT), which has 4 x 9 MPix CCDs, at a resolution of 0.33"/pix. We continuously acquired data in 12-minute fields of 12 exposures, in several multi-night observing runs. For each field, image processing (sensor correction and plate solving) took 8 minutes and the STU detection pipeline took <2 minutes. Moreover, the trajectory scan took only 8 seconds. This runtime allowed real-time processing on a PC with a GPU under 1000 euro. For the much larger KASI telescope we have used the Korean Microlensing Telescope Network Camera (KMTCam), of 4 x 85 Mpix, at a resolution of 0.4"/pix. The STU runtime for the KASI telescope has been 7 minutes at a maximum apparent motion of 10"/min. This runtime is the result of STU having a pixel processing rate (PPR), which is the number of image pixels co-added per unit time, of 1700 GPix/sec. In the newer versions of STU (v0.7), we have added support for repurposing "small integer dot product" instructions, originally introduced in newer GPU hardware for accelerating machine learning workloads, for co-adding pixels faster. This improved the PPR to 2400 GPix/sec.

As an example of the results in these observing runs, we present two examples of asteroids detected by our software. First, in left figure, the detection of NEAs 2023 DZ₂, formerly catalogued as Virtual Impactor, discovered and recovered by our group between February 27 and March 1 2023 (MPEC 2023-F12, https://minorplanetcenter.net/mpec/K23/K23F12.html), Second, in right figure, the detection of 2024 CW₂, at an apparent motion of 9.54"/min. This object was discovered by our team in the night of 11/12 February 2024, and paired soon after being published to 2007 EG.

Acknowledgements

This work had been supported by a grant of the Romanian National Authority for Scientific Research -- UEFISCDI, project number PN-III-P2-2.1-PED-2021-3625.

References
[1] B. Gladman, et al. (1997) Astronomy and Astrophysics 317:L35.

[2] C. Zhai, et al.(2018) Technical note: Asteroid detection demonstration from skysat-3 b612 data using synthetic tracking.

[3] M. Stănescu, et al. (2021) Astronomy and Computing 35:100453.

[4] O. Vaduvescu, et al. (2025) New Astronomy 119:102410 ISSN 1384-1076.

How to cite: Stanescu, M., Popescu, M., Curelaru, L., Vaduvescu, O., Berteșteanu, D., and Predatu, M.: Detecting near-Earth objects using real-time synthetic tracking surveys, EPSC-DPS Joint Meeting 2025, Helsinki, Finland, 7–13 Sep 2025, EPSC-DPS2025-1403, https://doi.org/10.5194/epsc-dps2025-1403, 2025.

12:00–12:12

EPSC-DPS2025-1028

ECP

On-site presentation

Modelling impacts of aluminium from ablated space debris on atmospheric chemistry

Joanna Egan, Wuhu Feng, Alexander James, Daniel Marsh, and John Plane

The increasing quantities of anthropogenic objects in Low Earth Orbit (LEO) have led to concerns over space debris and collision risks in LEO, leading to the Federal Communications Commission’s introduction of the “5-year rule”, requiring deorbit of LEO spacecraft within 5 years of mission end. To mitigate the risk of debris impacting the surface and causing damage, spacecraft are increasingly designed to ablate in the atmosphere, with most of the mass being vapourised during re-entry [1]. This causes an influx of metals into the mesosphere, where they condense and settle into the winter polar stratosphere - around 10% of Junge layer sulphuric acid droplets have been measured to contain metals from ablated space debris. Some metals – Al, Li, Cu, Ni, Mn etc. – already exceed natural background levels from cosmic dust that has ablated in the mesopause region [2]. The effect of these metals on the stratosphere is not yet known, and space debris input has been projected to increase by more than an order of magnitude in the next 15 years [3]. It is therefore vitally important to determine the level of re-entering space debris that will cause significant changes to atmospheric aerosols and stratospheric chemistry, in particular to the ozone layer.

We model the catalytic impact of ablated aluminium which recondenses in the atmosphere has “space debris particles” (SDPs), predicted to be composed mainly of aluminium hydroxide nano-particles. These particles are predicted to catalyse chlorine activation from gas phase HCl, ClONO₂, and HOCl by catalytic reactions on the SDP surface. Their impact on polar stratospheric cloud freezing and subsequent chlorine activation is also considered.

We present results of a modelling study using a sectional aerosol model within an Earth system model (Whole Atmosphere Community Climate Model with the Community Aerosol and Radiation Model for Atmospheres, WACCM-CARMA). We simulate the transport of SDPs and meteoric smoke particles (MSPs) produced by condensation of Fe and Mg silicates from ablated cosmic dust. The particles grow by coagulation and deposition of sulphuric acid through 28 size bins (0.34 nm to 1.6 µm radius, where MSPs injected at 0.34 nm radius, SDPs at 10 nm radius, and sulphuric acid is allowed to condense both heterogeneous on the MSPS and SDPs, and homogeneously. The SDPs and MSPs are initially injected in concentrations consistent with current models and observations (7.9 t d^‑1 MSPs and 0.96 t d^-1 SDPs) to assess the transport and lifetimes of the particles in the atmosphere.

The effect of increasing the mass of SDPs in line with future increases in space travel is also simulated. The maximum possible impact of SDPs on stratospheric chemistry is then estimated from the available SDP surface area and assuming upper limits for unmeasured physico-chemical parameters. Condensation of sulphuric acid onto the particles during their descent through the stratosphere reduces the surface area available for catalytic chlorine activation, but this is highly sensitive to the shape parameter assumed for the SDPs and the fraction of the surface that is coated by the condensed sulphate.

The reaction rates and physical properties of SDPs adopted in this model have not been measured or observed. Taking reasonable upper limits for values indicates that the reactions have the potential to do significant damage to the stratospheric ozone layer due to increased chlorine activation in winter and spring. The precise morphology and composition of particles must be characterised by in situ sampling of the stratosphere, and laboratory measurements of the rate constants are also required to better constrain estimates.

References

[1] Kelley 2012, https://ntrs.nasa.gov/citations/20120002794

[2] Murphy et al. 2023, https://doi.org/10.1073/pnas.2313374120.

[3] Schulz & Glassmeier 2021, https://doi.org/10.1016/j.asr.2020.10.036

How to cite: Egan, J., Feng, W., James, A., Marsh, D., and Plane, J.: Modelling impacts of aluminium from ablated space debris on atmospheric chemistry, EPSC-DPS Joint Meeting 2025, Helsinki, Finland, 7–13 Sep 2025, EPSC-DPS2025-1028, https://doi.org/10.5194/epsc-dps2025-1028, 2025.

12:12–12:24

EPSC-DPS2025-1732

On-site presentation

Space threats and their effects, SafeEarth Research Programme perspective

Eija Tanskanen, Shabnam Nikbakhsh, Reko Hynönen, Arttu Tiainen, Jouni Envall, Matias Meskanen, Kimmo Halunen, Pentti Haddington, and Emma Bruus

SafeEarth Research Programme combines research fields in space safety, cybersecurity and human security into a comprehensive and highly timely security theme, namely comprehensive security. SafeEarth examines changes in the magnetic environment and enhance our understanding on how space threats affect to the societies and nature in the arctic regions.

Arctic polar areas experience effects of space threats that are unforeseen at the more southerly latitudes. Rapid magnetic fluctuations and pulsations are observed at the auroral oval during large and small geomagnetic disturbances driven by space storms. The strongest geomagnetic disturbances are released from the complex active regions of the Sun. The harmfull effects of space storms affect technology and humans in space and on ground. This presentation will show examples on these effects and describe the methods developed to better forecast when the space threats should be expected and how to best protect the vulnerable assets agains space hazards. We will show aims and results on the European Defence Fund project (BODYGUARD) and NATO Science for Peace and Security funded project Dynamics above the epicentre of climate change (DECC).

How to cite: Tanskanen, E., Nikbakhsh, S., Hynönen, R., Tiainen, A., Envall, J., Meskanen, M., Halunen, K., Haddington, P., and Bruus, E.: Space threats and their effects, SafeEarth Research Programme perspective, EPSC-DPS Joint Meeting 2025, Helsinki, Finland, 7–13 Sep 2025, EPSC-DPS2025-1732, https://doi.org/10.5194/epsc-dps2025-1732, 2025.

Posters: Thu, 11 Sep, 18:00–19:30 | Finlandia Hall foyer

Display time: Thu, 11 Sep, 08:30–19:30

F125

EPSC-DPS2025-836

On-site presentation

Space Safety through situational awareness

Ari-Matti Harri, Hannu Koivula, Tiera Laitinen, Karri Muinonen, Kirsti Kauristie, Jouni Peltoniemi, and Antti Kestila

Space Situational Awareness (SSA) entails the detection, tracking, and comprehension of spaceborne objects and phenomena that could potentially affect Earth or space operations. It encompasses three primary components: Space Surveillance and Tracking (SST), which monitors satellites and space debris to prevent collisions; space weather monitoring, which observes solar activity—such as solar flares, coronal mass ejections, and energetic particles—that can disrupt satellite systems, communications, and power grids; and Near-Earth Object (NEO) monitoring, which tracks asteroids and comets that may approach or impact Earth.

Space weather can significantly impact planetary probes and satellites orbiting celestial bodies. These events release high-energy particles and radiation that may damage satellite electronics, disrupt communication systems, and degrade the performance of sensitive instruments on space probes. In extreme cases, these energetic particles may compromise the physical integrity of spacecraft, potentially leading to mission failure. Moreover, space weather can interfere with a satellite’s ability to accurately navigate and maintain its orbit, particularly if the spacecraft operates in or near areas like the Van Allen radiation belts or in proximity to celestial bodies such as Jupiter, whose intense magnetosphere interactions can exacerbate these effects. In sum, space weather poses a persistent risk to the longevity and operational integrity of planetary probes and satellites, necessitating continuous monitoring and mitigation to ensure mission success.

SSA is a fundamental element of space safety, which refers to the secure and sustainable use of space. Space safety not only includes SSA but also encompasses the protection of human spaceflight, Earth-based infrastructure, and space commerce from hazards such as space debris, space weather, and asteroid impacts. Additionally, it involves efforts to actively mitigate these risks through scientific and technological advancements.

As satellite operations and global reliance on space services increase, space safety and SSA have emerged as critical global priorities. The growing congestion of near-Earth orbit, with an increasing number of satellites and debris, heightens the likelihood of collisions that could generate dangerous debris fields, further disrupting global systems. Even a single collision can lead to the creation of thousands of fragments, posing long-term hazards to other spacecraft.

Simultaneously, solar activity results in geomagnetic storms and other forms of space weather that can impair satellite functionality, GPS, aviation, and even terrestrial infrastructure such as power grids. While most space debris burns up during atmospheric re-entry, larger fragments may survive and pose a risk to people or property on Earth. Planetary defense, which focuses on the detection and mitigation of NEOs, is another critical pillar of SSA, dedicated to preventing potentially catastrophic impacts from asteroids and comets.

Disruptions to satellite services also carry significant economic consequences. Precision industries, aviation, and high-frequency financial trading all depend on uninterrupted GPS and communication systems. The insurance sector has identified space weather as a high-impact, underinsured threat, yet it remains largely absent from mainstream economic risk models.

In response to these escalating risks, Finland is establishing the Finnish Space Situational Awareness Center (FSSAC) through international cooperation, such as the EU Space Surveillance and Tracking (EU SST) program, as well as national observations and modeling initiatives. This effort aims to ensure the safe and resilient operation of critical space infrastructure in collaboration with international partners, while providing SSA services to national authorities. FSSAC will be composed of two centers: a civilian command center (C-FSSAC) led by the Finnish Meteorological Institute and supported by the National Land Survey, and a military command center (M-FSSAC) operated by the Finnish Defence Forces. These centers will have complementary roles in monitoring, analysis, and response.

FSSAC will operate within an international framework, contributing to EU-led initiatives and fostering direct cooperation between Finland and the United States. It will integrate global SSA data, enhancing it with national information and analysis tailored to Finland’s security and infrastructure needs. Domestically, FSSAC will function as a center of expertise, utilizing assets such as the Metsähovi laser telescope and national space weather monitoring systems.

The civilian command center will be tasked with monitoring orbital objects, assessing collision risks, forecasting space weather effects, and evaluating the potential impact of re-entering debris or NEOs. It will provide relevant stakeholders with timely information to support informed decision-making. SST services at FSSAC will reduce collision risks, minimize disruptions, and safeguard national infrastructure. As satellite traffic continues to increase, SST is not merely a technical necessity—it is an investment in economic and societal stability.

In addition to passive monitoring, space safety places significant emphasis on preventive and mitigative measures. International efforts are increasingly focused on not only detecting hazards but also developing active responses, such as debris removal technologies and planetary defense missions designed to deflect or destroy threatening NEOs. Scientific research, technological innovation, and international cooperation are crucial to these strategies.

FSSAC will also track objects entering Earth’s atmosphere and forecast their potential impact zones. Although most large debris burns up upon re-entry and satellites are typically directed toward ocean disposal zones, this capability is vital for public safety and government preparedness.

By leveraging synergies with existing 24/7 services at the Finnish Meteorological Institute, the civilian command center will operate efficiently at a relatively low cost. Full operational readiness is anticipated by 2027. Upon its establishment, FSSAC will significantly enhance Finland’s resilience to growing space-related risks, while contributing to global space safety.

Through the integration of scientific surveillance, forecasting, and mitigation capabilities, FSSAC will support a safer and more sustainable use of space. In doing so, it will protect not only national interests but also the broader global society and economy, both of which now depend heavily on uninterrupted access to the space domain.

How to cite: Harri, A.-M., Koivula, H., Laitinen, T., Muinonen, K., Kauristie, K., Peltoniemi, J., and Kestila, A.: Space Safety through situational awareness, EPSC-DPS Joint Meeting 2025, Helsinki, Finland, 7–13 Sep 2025, EPSC-DPS2025-836, https://doi.org/10.5194/epsc-dps2025-836, 2025.

F126

EPSC-DPS2025-986

ECP

On-site presentation

Modelling the ground effects of a Carrington-class geomagnetic storm using global geospace simulation

Elena Marshalko, Ari Viljanen, Liisa Juusola, Tiera Laitinen, and Kirsti Kauristie

An extreme geomagnetic storm of a magnitude comparable to the historical Carrington event of 1859 would pose a substantial threat to ground-based technological infrastructure, particularly to electrical power transmission systems. A previously published simulation by Blake et al. (2021, Space Weather, doi:10.1029/2020SW002585) reconstructed the magnetic field observations at Colaba, India, during the Carrington storm and provided estimates of magnetic field variations around the world. Building upon this work, we employ a physics-based, first-principles modelling approach to estimate the induced geoelectric field in the Fennoscandian region. This approach uses a detailed three-dimensional model of the Earth’s subsurface electrical conductivity.

To contextualize the severity of a potential Carrington-class event, we compare the resulting geoelectric field estimates with those modelled due to the geomagnetic storm of October 2003 (commonly referred to as the Halloween storm) - one of the most intense geomagnetic disturbances recorded in the past 100 years. The Halloween event is particularly well-suited for comparison due to the availability of high-resolution, spatially dense magnetometer data across Northern Europe. Our analysis indicates that the maximum geoelectric field magnitudes modelled for a Carrington-class storm in Fennoscandia could exceed those produced during the Halloween storm by a factor of approximately 4-10. These findings underscore the elevated geoelectric hazard posed by such extreme geomagnetic activity, especially in regions with high-latitude infrastructure.

How to cite: Marshalko, E., Viljanen, A., Juusola, L., Laitinen, T., and Kauristie, K.: Modelling the ground effects of a Carrington-class geomagnetic storm using global geospace simulation, EPSC-DPS Joint Meeting 2025, Helsinki, Finland, 7–13 Sep 2025, EPSC-DPS2025-986, https://doi.org/10.5194/epsc-dps2025-986, 2025.

F127

EPSC-DPS2025-1255

ECP

On-site presentation

Probing Earth’s ionosphere with novel ionosonde and ray-tracing solutions

Elias Hirvonen, Kirsti Kauristie, Esa Kallio, Tomas Tallkvist, and Michael Fletcher

Near-real-time observations of Earth's ionosphere electron density are an important input for a variety of technologies and services that depend on HF radio or space-based communication. An ionosonde is a radar that performs ionospheric sounding and provides measurements of the current state of the ionosphere. The critical parameters (foF2, foF1, foE) derived from ionosonde data define the maximum frequencies of the ionospheric E- and F-layers for estimates of their electron densities. Users of space weather services, e.g. in aviation, use these parameters to search for optimal frequencies in over-the-horizon communication. Ray-tracing codes as combined with statistical ionospheric electron density models are often used in more accurate radio wave propagation assessments.

Finnish Meteorological Institute (FMI) started ionospheric measurements in Southern Finland with a new ionosonde in February 2025. The ionosonde, manufactured by RF-shamans Ltd, is based on a software-defined radio (SDR) implementation, which enables its small, both transmitter and receiver antenna size (≈ 1m^3), quick and flexible operation, and low transmitting power level (< 0.5W). In addition, the ionosonde antennas are easy to install and relocate when needed. The ionosonde currently produces vertical ionograms using a 10 second sweep five times per minute. The measurements are filtered, and the critical parameters are detected in real-time. The more detailed real-height analysis is currently operated manually and an automated process is under development.

In addition, FMI analyzes ionograms with a numerical ray-tracing software, AU-Ray [1]. The polynomial analysis software (POLAN) is also used in real-height analysis to verify the scaling results from the AU-Ray method. AU-Ray calculates the radio wave propagation path step-by-step based on Appleton-Hartree (AH) or Booker Quartic (BQ) Hamiltonians. AU-Ray allows ray-tracing not just with statistical ionospheric models but also with background conditions customized by the user. The code is based on freely available programming languages and open-source models. The software package includes also a mapping tool, which allows simulations with large quantities of rays, for propagation maps in user-defined ionospheric conditions. In the presentation, we describe the main specifications of the FMI ionosonde and the AU-ray code, analyze ionosonde performance during a recent space weather storm, and demonstrate the usage of AU-Ray in the real-height analysis, while comparing the simulation results to the ionosonde measurements.

[1]: E. A. O. Hirvonen, K. Kauristie and E. Kallio. “AU-Ray Program for Modelling Radio Wave Propagation in the Ionosphere“. Radio Science. Manuscript submitted May 2025.

How to cite: Hirvonen, E., Kauristie, K., Kallio, E., Tallkvist, T., and Fletcher, M.: Probing Earth’s ionosphere with novel ionosonde and ray-tracing solutions, EPSC-DPS Joint Meeting 2025, Helsinki, Finland, 7–13 Sep 2025, EPSC-DPS2025-1255, https://doi.org/10.5194/epsc-dps2025-1255, 2025.

F128

EPSC-DPS2025-1269

On-site presentation

Multi-Wire μ-Tether for Plasma Brake and Space Debris Mitigation

Petri Toivanen, Pekka Janhunen, Jarmo Kivekäs, Jouni Polkko, Maria Genzer, Jannika Vesalainen, Jari Sinkko, Tuomas Simula, Mikael Hakuri, Perttu Yli-Opas, and Iaroslav Iakubivskyi

Plasma Brake is a compact module to be mounted to a spacecraft or to a piece of space junk for orbital braking and ultimate atmospheric demise. The braking force is generated by the ionospheric plasma Coulomb drag that is analogous to the air drag in the neutral atmosphere. The dragging obstacle against the orbital plasma RAM flow is established by an electrostatic field with a high voltage difference with respect to the ambient plasma, typically -1 kV. The potential structure is supported by a long 4-wire tether with single aluminium wires with thickness of less than 50 μm. The redundant structure makes the tether resilient against μ-meteoroids. However, the hair-thin tether sets no harm to other space assets based on micro-meteoroid and space debris flux models such as MASTER-2009.

The multi-wire structure required for redundancy against the µ-meteoroid and space debris flux of the space environment is realised through a method of twist bonding traditionally used for the chicken wire. In the case of the Coulomb drag tether, the diameter of the individual wires is at most 50 µm, which introduces the main technological challenge. Preceding this method, ultrasonic bonding and cold welding were used for tether production until the twist bonding method was developed and adopted.

Until recently, the tethers have been produced running a manually operated machine while an automated tether factory has been developed alongside with manual production. Presently, we have manually produced three flight tethers each 60 metres long for three CubeSat missions of FORESAIL-1, EstCube-2, and FORESAIL-1p. One of the objectives of our CubeSat missions is to test the Plasma Brake tether deployment and high-voltage power systems, and most significantly measure the Coulomb drag effect in orbit. The first version of the automated factory was successfully run late 2024. Several samples with a length of 100+ metres were successfully produced. The second version of the tether factory is ready for set up and test runs in May, 2025. The aim is to produce tether samples with a length of 1000+ metres. This work is a central part of the Dragliner project of ESA aiming at a fully functional Plasma Brake engineering model at TRL of 5-6.

How to cite: Toivanen, P., Janhunen, P., Kivekäs, J., Polkko, J., Genzer, M., Vesalainen, J., Sinkko, J., Simula, T., Hakuri, M., Yli-Opas, P., and Iakubivskyi, I.: Multi-Wire μ-Tether for Plasma Brake and Space Debris Mitigation, EPSC-DPS Joint Meeting 2025, Helsinki, Finland, 7–13 Sep 2025, EPSC-DPS2025-1269, https://doi.org/10.5194/epsc-dps2025-1269, 2025.

F129

EPSC-DPS2025-1377

ECP

On-site presentation

Plasma Brake Payloads onboard CubeSat Missions of FS-1, EC-2, and FS-1p

Jarmo Kivekäs, Petri Toivanen, Jouni Polkko, Pekka Janhunen, and Maria Genzer

We have developed a Coulomb drag propulsion device that accommodates a 50-m electrostatic tether that can be biased to a high voltage with respect to the ambient space plasma. The high voltage electrostatic sheath around the tether serves as an obstacle that perturbates the plasma mass flow causing Coulomb drag and a net propulsive force to the tether-spacecraft system. The ultimate objective for our device is to measure the Coulomb drag effect for the first time in space and to demonstrate that the Plasma Brake can be mass-effectively and safely used for orbital braking of spacecraft and space junk to mitigate the space debris

Our device has been designed for 3-unit CubeSats or larger platforms. We have already built and integrated two of them into CubeSats of FORESAIL-1 and ESTCube-2. These were launched in 2022 and 2023, respectively. However, due to communication problems and launch pod failure, we were not able to carry out the Coulomb drag tests. Presently, the third Plasma Brake flight module has been delivered for the integration into FORESAIL-1p to be launched in Autumn 2025, thumbs up.

The key components of the device are a reeling system for the tether deployment and a high-voltage power system. The deployment is based on the centrifugal force provided by the spin-stabilised satellite platform. The reeling is driven by a stepper motor embedded and supported by a ceramic bearing inside the tether reel. The deployment system is covered and supported against launch vibrations by the tether chamber. There are three launch locks mounted to the tether chamber to secure the tether reel and tether tip mass during the launch. In addition to the high-voltage electronics, there is the electric power system and control electronics, most importantly, the stepper motor driver and microcontroller unit. The high-voltage system also monitors the tether voltage and current since these are expected to vary depending on the ionospheric density. In this presentation, we describe the device in further details and cover the basics of the Coulomb drag, and Plasma Brake operations in orbit.

How to cite: Kivekäs, J., Toivanen, P., Polkko, J., Janhunen, P., and Genzer, M.: Plasma Brake Payloads onboard CubeSat Missions of FS-1, EC-2, and FS-1p, EPSC-DPS Joint Meeting 2025, Helsinki, Finland, 7–13 Sep 2025, EPSC-DPS2025-1377, https://doi.org/10.5194/epsc-dps2025-1377, 2025.