University of Warwick leads the proposal through the Warwick Centre for Space Domain Awareness (CSDA). The CSDA has been established to tackle issues relating to the safety and sustainability of satellites, including: the timely acquisition of precise datasets to detect, track and/or characterise objects in orbit; the fusion of physical and human-based information for improved object tracking; the modelling and prediction of space weather, and the quantification of associated risk. Expertise spans related space weather, space surveillance and tracking, statistical uncertainty and engineering domains. 

Open Cosmos is a space company that delivers satellite missions from start to finish. It offers a comprehensive service that includes the design, manufacture, launch and operations of satellites tailored for companies, institutions and governments around the world. Currently Open Cosmos has flight heritage with 2U, 3U and 6U CubeSats, with 12U and 16U cubesats in development (as well as a further four 6U missions). Some of the 16U satellites in development are scientific for the NanoMagSat mission with the ESA SCOUT programme which will look at the Earth’s magnetic field. As part of NanoMagSat the Open Cosmos platform has had key subsystems magnetically characterised and adjusted to reduce the magnetic noise of the platform. 

The Consortium/Authors include leading experts from Universities, institutions and companies across Europe, North America and Africa. Expertise spans the latest scientific instruments for space applications, their utilisation for space- and ground-based observations of orbiting satellites and debris, our magnetosphere, thermosphere, ionosphere and exosphere, and space weather impacts. The team also includes leading expertise in a variety of modelling approaches embedded within this mission design. The following models are involved: the Advanced Ensemble electron density Assimilation System (AENeAS); the Centre for Geospace’s Multiscale Atmosphere Geospace Environmental Model (MAGE); the Gorgon global magnetosphere-ionosphere model, and further satellite charging and drag-based models. This team includes the following members:

University of Warwick: Ravindra Desai; Don Pollaco; Sandra Chapman; James Blake; Dimitri Veras; Jeremie Houssineau; Bogdan Gnat; Tony Arber; William Crofts  

Imperial College London, UK: Jonathan Eastwood; Martin Archer; Patrick Brown

Institut für Weltraumforschung (IWF), Austria: Martin Volwerk; Rumi Nakamura; Magnes Werner; David Fischer; Micheal Steindorfer; Helmut Lammer

University of Stuttgart, Germany: Stefan Loehle

University of Bath, UK: Cathryn Mitchell; Robert Watson

University of Oslo, Norway: Lasse Clausen

Eidsvoll Electronics AS, Norway: Tore Bekking

University of Bergen, Norway: Karl Landual; Spencer Hatch

University Centre in Svalbard, Norway: Stein Haaland

Northumbria University, UK: Eamon Scullion

University of Calgary, Canada: Johnathan Burchill; David Knudsen

University College London (UCL), UK: Graziella Branduardi-Raymond; Andrei Samsonov

Institutet För Rymdfysik (IRF), Sweden: Anders Eriksson

University of Leicester: Jennifer Carter; Andrew Cheney

South West Research Institute (SWRI), USA: Keiichi Ogasawara; Dhiren Kataria

Laboratoire Atmosphères, Observations Spatiales (LATMOS), France: Dimitra Koutroumpa; Ronan Modolo; Jean-Yves Chaufray

British Antarctic Survey (BAS), UK: Mervyn Freeman; Andrew Kavanaugh

NASA Goddard Space Flight Centre, USA: David Sibeck

Rutherford Appleton Laboratory (RAL), UK: Malcolm Dunlop

Naval Post Doctoral School (NPS), USA: Shambo Bhattercharjee 

Space Science and Geospatial Institute, Ethiopia: Daniel Chekole; Nigussie Giday 

Bahir Dar University, Ethiopia: Melesssew Nigussie  

University of Birmingham, UK: Sean Elvidge; David Themens

National Center for Atmospheric Research (NCAR), USA: Dong Lin

Johns Hopkins University/Applied Physics Laboratory (JHU/APL), USA: Kareem Sorathia

University of Manchester, UK: Katharine Smith, Nicholas Crisp

This heliospheric CubeSat mission to LEO is enabled by rapid developments in low-cost technology miniaturisation and represents an unprecedented opportunity to observe and predict the forces of satellite drag. The science strategy is to obtain a coordinated, distributed set of space- and ground-based measurements using a swarm of CubeSats, to characterise the influence of the inter-connected magnetosphere-ionosphere-thermosphere (M-I-T) system on satellite decay levels across LEO altitudes. This cross-cutting endeavour builds upon magnetospheric formation flying measurement concepts, to constrain how solar activity modulates upper atmospheric dynamics, and utilises an extensive satellite tracking ground segment that spans radar, laser ranging and optical systems and connects knowledge and technologies across disciplines. Three overarching objectives drive this multi-point mission architecture, which are designed to facilitate model validation and data assimilation:

[1] Obtain coordinated measurements of upper-atmospheric variability across multiple LEO altitudes to resolve the 3D+time structural dynamics that affect satellite drag (measure the environmental inputs)

[2] Obtain and develop the best-ever space- and ground-based measurements of the CubeSats’ orbital elements to quantify orbital perturbations arising from solar activity (measure the technological effects)

[3] Model the energy and momentum transfer through the magnetosphere-ionosphere system which modulates atmospheric variability and satellite drag to improve understanding of cause and effect (close the loop on physical understanding)

Geomagnetic storms provide an unconstrained and unpredictable disruption to the M-I-T system due to the complexities and non-linearities inherent to the magnetospheric response to solar wind conditions (1). These are superimposed upon further atmospheric dynamics caused by: the varying photon flux across the solar cycle (2); the seasonally- and diurnally-varying irradiance due to the Earth’s tilt and rotation (3); and internal forcing deriving from planetary and gravity waves, atmospheric winds and acoustic waves (4). The thermosphere consequently exhibits complex spatial dynamics that propagate across altitudes, latitudes and longitudes (5). Disentangling transfers of energy through the M-I-T system and their effects on satellites orbits represents a major scientific and technological endeavour.

Figure 1: Mass percent concentration (left) and number density (right)
across LEO from the NRLMSIS-00 model for 2011.

Studies of atmospheric variability and satellite drag resulting from geomagnetic storms are in their infancy due to the lack of coordinated space weather and satellite tracking observations. Studies typically utilised isolated spacecraft or fortuitous conjunctions, with limited spatial and temporal coverage (6-7). The atmospheric densities in LEO responsible for satellite drag, however, vary non-linearly across altitude, see Figure 1. Higher-order altitudinal constraints are therefore required to understand the response to magnetospheric driving which initially manifests at high latitudes due to non-uniform ionospheric current systems. These current systems induce travelling atmospheric waves/disturbances that propagate globally resulting in complex modulations of satellite drag across latitudes and longitudes (5,7). Simultaneous 3D observations across densities that decrease exponentially with altitude are therefore necessary to resolve the global evolution of the thermosphere and the precise spatial and temporal scales at which change occurs. ROARS will obtain high-fidelity measurements across LEO of the neutral particles that induce satellite drag.

Geomagnetic activity can be significant on timescales less than an hour (7) and small satellites in LEO can exhibit order of magnitude variations in drag during quiescent and disturbed conditions with typical small satellite orbital decay rates ranging from less than 1 metre up to 10 metres per orbit (8). Understanding how and where this orbital decay manifests is critical to predicting satellite trajectories and collision risk. Ground-based observations from the international satellite laser ranging (ILSR) network, augmented by dedicated stations at mid- and high-latitudes, will obtain sub-cm absolute positioning and inter-satellite laser ranging will be utilised with the potential to resolve mm-scale differences in the CubeSats’ positions. These provide the means to understand the role of meso-scale (10-1000 km), and potentially even micro-scale (<10 km), atmospheric dynamics (5,7). Optical tracking techniques, capable of resolving multiple targets simultaneously, will be developed and utilised in tandem (9). Further measurements from on-board GNSS receivers will also provide continuous tracking of their orbital dynamics when not in view of the ground segment. State-of-the-art satellite tracking will directly relate orbital perturbations to local and global atmospheric density variations.

Figure 2: Schematic showing how magnetospheric current systems close
through the upper atmosphere where they induce Joule heating (Credit: J. Slavin).

The three dimensionally distributed mission architecture enables global insight into M-I-T coupling to unveil the fundamental drivers behind atmospheric variability and satellite drag. Polar conjunctions at multiple altitudes, with satellites distributed along-track with small separations in local time, will enable comprehensive observations of the coupling between the magnetosphere and ionosphere, through measuring; field-aligned and auroral current intensities and topologies (10-12), ion drifts and composition and electric fields (13). In addition to this, the high-cadence measurement capabilities of modern miniaturised electronics can resolve fine-structure magnetosphere-ionosphere coupling signatures (14-15), localised electron pressure gradients (16), and the total electron content at each spacecraft (17) to constrain how micro-scale energetic particle precipitation and ion and neutral upwelling modulate meso- and macro-scale atmospheric dynamics. This novel space-based measurement set will be targeted to spatially coincide with the ground-based EISCAT_3D incoherent scatter radar field-of-view, to simultaneously constrain auroral dynamics and ionospheric conductivities lower down in the D and E regions where the ionospheric current systems close, see Figure 2, to provide the full picture of M-I-T coupling. ROARS will measure magnetic field and plasma signatures of M-I energy transfers that cause atmospheric heating and elevated satellite drag.

Payload 

Table 1. Initial Science Traceability Matrix  

Launch & Orbit  

The swarm design shall nominally be split between two orbital planes with a small RAAN offset.

The launcher shall deploy the satellites into a 500-550 km high inclination orbit.

The launcher should deploy the satellites into a polar (~90° inclination) orbit.

Primary propulsion or an in orbit transport vehicle (OTV) may be used to move from the insertion orbit to the desired orbital inclination.

In the first orbital plane, satellites shall operate in 3 coplanar orbits, see Figure 3 (one circular orbit at 500-550 km and two elliptical orbits with perigees at 400-450 km and apogees at 600-650 km and argument of periapsis at 90° and 270° respectively.)

The second orbital plane shall have a small separation in local time with two satellites in single circular orbit, phased with the first orbital plane to enable side-by-side measurements (12).  

Each satellite should have the capability to raise/lower its orbital altitude into the desired operational orbit from the insertion orbit.

Each satellite shall have the capability to maintain its orbit relative to the other satellites in the swarm (RAAN precession of all the orbits from the J2 perturbation is expected and allowed)

There shall be two satellites per orbit, with a varying along-track separation between 100-1000 km (TBC).

Satellites in each orbital plane should be phased in a way that they align when passing over each pole.

The planes and eccentricities can be phased so that polar conjunctions of the coplanar groups happen simultaneously, with all spacecraft over a single pole, or with one group at each pole.

Figure 3: Orbital plane of the CubeSat swarm orbits each containing a CubeSat pair. The upper and lower orbits have opposing arguments of perigee (blue & green) in respective hemispheres to maintain constant orbital periodicities. The intermediate orbit is circular (red). The northern perigee shown is located over EISCAT_3D in Skibotn, Norway. A further orbital plane with a small separation in local time (not shown) will contain 2 further satellites.

Space Segment

The satellites shall be sized to operate for a baseline mission of at least 24 months and ideally up to 36 months (TBC - Phase 0).

The satellites should be sized to operate for a further 24 months (TBC) without orbit maintenance. (sampling atmosphere as the orbital altitudes decrease).  

Each satellite shall have propulsion capabilities to maintain operational orbits for the baseline mission.

The satellite shall have an ADCS capability which allows for the payload face to point in the velocity direction with an accuracy of (TBD).

The satellite should have an ADCS capability to satisfy inter-satellite laser ranging requirements (TBD).  

TT&C (Telemetry, Tracking & Command) communications shall be encrypted (via AES – 128 or 256 TBC).  ​

TT&C gain pattern shall allow satellites to be contacted while in any attitude and any tumbling state. 

TT&C link budget shall allow for a minimum of 10MB downlink of housekeeping telemetry data per day. 

High data rate (HDR) downlink shall be sized to allow for all payload data to be downloaded each day. (TBD baseline X-Band).

The satellites shall be no larger than 16U CubeSats. 

The satellites shall have a mass no greater than 32 kg.

Ground Segment

The MOC (Mission operations centre) shall have the ability to command all satellites in the swarm in parallel.

The MOC (Mission operations centre) shall have the ability to receive telemetry from all satellites in the swarm in parallel.

Ground stations (GS) shall be located at high (& low) latitudes which will give the highest number of passes per day. (Baseline to use K-SAT Svalbard TBC).

The ground segment shall include laser ranging capabilities to accurately determine each satellite’s position.

Orbits: Eight CubeSats can utilise commonly available rideshare launches into the popular 10:30 / 22:30 polar LEO. Orbits with inclinations between 95° and 105° will still maintain sufficient auroral coverage, see Figure 2, and across magnetic latitudes but also enable differential precession of the local time of the orbit and provide seasonal coverage. An initial orbital configuration has been suggested as the baseline for the Phase 0, but a full analysis on the mission implications and satellite number shall be examined. In keeping with the concept of a swarm, the orbital and inter-satellite positionings are required to be maintained over timescales of days-weeks, as opposed to orbit-by-orbit. Drifting around the nominal configuration will not impinge upon the novelty of the scientific measurements or upon reaching the outlined objectives.

The satellites will occupy two circular target orbits at 500-550 km and two elliptical orbits with perigees at 400-450 km and apogees at 600-650 km, see Figure 3. Contemporaneous radial measurements at high latitudes will be achieved through the use of opposing arguments of perigee to ensure constant orbital periodicities. Each of the 3 orbits will contain a pair of satellites with a 100-1000 km along-track separation to disentangle along-track spatial and temporal dynamics, while polar conjunctions with one another. 

The spacecraft will possess propulsion capabilities in the baseline mission to maintain altitudes and will run for 24 months to resolve variations in geomagnetic activity associated with the Earth's dipole tilt (19). This will be followed by a de-orbit, to sample lower atmospheric regions not frequently observed by spacecraft. A longer mission is beneficial to reach from the planned launch in 2029 near solar minimum to the maximum of solar cycle 26 in 2035.

Launch & transfer: It is expected to fill the whole swarm from a single launch. Ideally this would be a rideshare mission with injection close to the desired operational orbits onboard a European launcher such as the Ariane 6. It is also possible that a dedicated European small satellite launcher could be used (e.g Orbex, RFA, Skyrora) to go precisely to each orbital plane. There is a range of injection parameters that could satisfy the requirements when combined with the propulsion capabilities of the satellites. An assessment will be made about the exact acceptable ranges of launch injections, as well as an assessment of orbital transfer vehicles which could be used as an option to maintain more ΔV for station-keeping by performing the orbit positioning post-launch.

Disposal: End of life disposal is part of the planned mission activities and it is hoped that science can be collected during the slow orbital degradation. The satellites will have propulsion capabilities so they can control their orbital degradation at end of life. The planned orbits are also expected to be naturally compliant with ESA debris mitigation guidelines in the case of satellite failure. 

Payloads per satellite [All payload selections to be confirmed during Phase 0 study. Table 1 shows baseline payloads]: Atmospheric neutral measurements will be required to detect particle concentrations at up to 800 km altitude. Along-track spacecraft will have separations of 100-1000 km, and high-cadence measurements are required to effectively use this spatio-temporal separation. A three dimensional ion velocity distribution will measure thermal ion drifts and associated electric field. The combination of a compositional ion and neutral mass spectrometer as well as thermal ion drift instrument will provide further information on the composition of the top-side ionosphere and the role of minor species. Polar conjunctions will enable calculations of the auroral currents through magnetic field measurements and field-aligned current signatures (12), low latitude field aligned resonances and kinetic Alfvén waves. The Langmuir Probes are also able to resolve ultra-high variations of electron densities and temperatures and, in combination with high cadence GNSS TEC and magnetic field measurements, allows the mission to target micro-scale auroral dynamics at multiple locations and examine how these control the global dynamics. The inter-satellite laser ranging will provide coincident measurements of the separation between the CubeSat pairs with precise clock synchronisation and return a single data product to the ground.

Concept of operations: The satellites shall fly with the large area (4U x 2U) face in the velocity direction to allow the atmospheric instruments to have the appropriate openings. The satellite will be controlled in a two axis control with the secondary attitude vector sun pointing (for increased power generation) or ground station pointing during ground station passes.
The different payloads will each take measurements at different frequencies throughout the orbit, with the most important regions needing high frequency measurements being those at higher latitudes. As a baseline, two ground station passes will be made per day but a data budget needs to be analysed once the payloads are finalised.

During routine operations there will be a number of orbits per week used for propulsion manoeuvres to maintain the desired orbits, during these orbits not all science will be performed. 

Ground Incoherent Scatter Radar Segment: The precise northern perigee/apogee will be optimised to enable conjunctions with EISCAT_3D at Skibotn, Norway (69.3 N, 20.4 E). EISCAT_3D will provide measurements of three scalar quantities (plasma density and electron and ion temperature) on timescales of seconds, and one vector quantity (ion drift velocity) on timescales of minutes, with a spatial resolution of 1–10 km and within a large volume of order 10⁶ km³ (i.e., a semi-rectangular box extending ~100–200 km horizontally and a few hundred km vertically), a technological achievement that is well beyond the capability of all existing incoherent scatter radar systems. From these primary measurements, estimates of the neutral wind over altitudes of ~90–125 km as well as estimates of ionospheric currents and electric fields are also planned to be derived. This will complement the ROARS data by helping to remove the ambiguity of whether observed variations are temporal or spatial in nature and provide context for the detailed satellite measurements during conjunctions.  In addition to EISCAT_3D infrastructure, an array of three all-sky imaging Fabry-Perot spectrometers in Finland and Sweden is currently being installed to provide tristatic measurements of thermospheric neutral winds and temperatures over the three EISCAT_3D sites (PI: Mark Conde, University of Fairbanks, Alaska; see letter of support).

Ground Satellite Laser Ranging Segment: The International Laser Ranging Service (ILRS), will be utilised with additional customised data return from a Ground SLR station in Graz, Austria, and a proposed SLR station at Sky Blu in Antarctica which is already being investigated by the UK Space Agency and the British Antarctic Survey. These will be utilised to track passive retro-reflectors on the bottom of the satellites to provide absolute determination of satellites positions to cm-level accuracy. At clear-sky conditions Graz SLR station is routinely tracking geodetic, scientific and GNSS satellites during day and night. Besides traditional laser ranging, Graz is also capable of measuring the diffuse reflection of space debris objects by utilizing higher powered lasers and enables synergistic studies. Due to its high precision and temporal resolution, for LEO orbits in certain geometries individual corner cube retro reflectors can be distinguished within the datasets. The Antarctic Space Surveillance and Tracking station aims is to fill a large existing gap in the global SST network over the Southern Ocean and Antarctica south of 45° latitude. 

Ground Optical Tracking Segment: Optical telescopes will contribute to the tracking of the satellites, providing angles-only measurements for orbit determination. Facilities run by the University of Warwick’s CSDA at the Roque de los Muchachos Observatory, La Palma, are well placed to observe the proposed orbits (28.7 N, 17.9 E). With multiple optical telescopes, it will be possible to provide near-continuous coverage of the eight satellites as they pass from horizon to horizon, while the laser ranging sites will be limited to observing one satellite at a time. Additional telescopes could add contingency, or track in tandem with other instruments; the accuracy of derived orbits can be improved by combining angles-only data with range measurements from parallax or laser ranging (20). Optical brightness measurements will be used to monitor the behaviour of the satellites and to aid diagnoses in the event of anomalies.

In order to effectively utilise the satellite tracking segment and effectively use the data we will determine the exact drag coefficient of the CubeSats prior to launch. 

Phase 0:

1. OSSE (Observation, Simulation, Study Experiment) modelling by the Consortium (see Background) using a range of empirical, physics-based and data assimilation approaches.

2. Instrument analysis to ensure they can meet, or be developed to meet, the mission requirements. Instrument redundancies are included.

3. Detailed analysis of the orbital emplacement strategy, management of satellite separations and effects of Earth's geopotential (e.g. semi-frozen orbits, differential drag, ground track effects, orbital control limits) 

4. Mission design phase by Open Cosmos: Overall mission budgets (Mass, Power, Link, ΔV), Orbit design and simulations to validate lifetime and debris compliance. Initial Satellite configuration design. Iteration on launch and ground segment definition. ConOps definition.

Due to the science objectives of the mission a swarm of 8 satellites has been chosen. This gives 4 pairs of satellites, meaning conjunctions in three dimensions over the poles. The OSSE modelling in Phase 0 can determine the precise relationship and trade-offs between the measurements and the mission objectives. 

The satellites will each be a 16U CubeSat. The volume budget below shows the order of magnitude of each component. All payload values are indicative and will be confirmed during the Phase 0 study. 

Table 2: Volume 

Each satellite shall have a mass below 32 kg. The below initial mass budget summarises the expected masses of different components. All payload values are indicative and will be confirmed during the Phase 0 study. 

Table 3: Mass

Power:  Up to 47 W average orbital power is available for the payloads. At this stage the payload concept of operations requires a more detailed design so for the first analysis the payloads will be constrained by the available power of a standard platform. There are methods to increase the average orbital power through increasing the size of solar arrays and adding more battery capacity. 

The proposed satellite will be compliant with European deployers. Additional protrusions could be made use of for this mission as a mitigation if the volume budget were to become at risk

The mission shall be compliant with ESA debris mitigation rules. All the planned orbits are expected to be naturally compliant (TBC) in case of satellite failure. As part of the mission extension the deorbiting will be controlled and science done during the deorbiting process to make measurements at lower altitudes. Each satellite has propulsion capabilities which can be used to control this deorbiting and avoid any dangerous conjunction events. 

Below is a first ROM estimation of the costs at a mission level including contingency.  Once more certainty is present regarding the measurement requirements and the resulting core payloads, as will be solidified early in Phase 0, a more detailed cost estimation can be made.  

Table 4: Cost

The CubeSats will be designed in pairs with laser inter-satellite links connecting the front and rear of each pair, respectively. They will otherwise carry identical scientific payloads and all the instruments will interface with the spacecraft at a single point, greatly simplifying the satellite design. The instruments described contain overlapping capabilities as will be analysed and selected within the Phase 0 study. 

Atmospheric Neutrals

UCL option: The Ion Neutral Mass Spectrometer (INMS), designed and built by the UCL Mullard Space Science Laboratory, is a cylindrical electrostatic analyser (ESA) capable of making combined ion and neutral particle measurements in the spacecraft ram direction and designed to fit into a CubeSat form factor. INMS has previously been flown as part of the EU-funded QB50 and DISCOVERER programmes. The detection principle of INMS is similar to electrostatic analysers such as the PEACE (Cluster II), CAPS-ELS (Cassini) and SWA-EAS (Solar Orbiter) instruments built by MSSL. INMS includes an ioniser to ionise incoming neutrals for detection and ion filter plates to deflect ambient ions away from the detector when measuring neutrals. By varying the voltage across an ESA, we can construct the energy-per-charge distributions of particles. Through knowledge of the spacecraft velocity and the assumption that the spacecraft speed is much larger than the bulk velocity of the particles (usually true for atmospheric particles and spacecraft in low-Earth orbit), the mass per charge can be determined and thus the individual ion/neutral species identified.

Stuttgart option: The miniaturized Flux- Φ -Probe-Experiment (FIPEX) atomic oxygen sensor system for space applications developed by the University of Stuttgart are based on very first experiments using conventional lambda probes from Bosch as in use in motor combustion to measure molecular oxygen in plasma flows. In principle, FIPEX is an electrochemical sensor. Two electrodes are separated by an ion conducting, but electrically insulating, electrolyte. In its classical potentiometric operation, a reference volume with reference density is required to measure the unknown partial pressure of the species of interest. The FIPEX sensor in amperometric operation allows a small sensor design particularly useful for spaceflight applications. FIPEX sensors are currently flight qualified for being part of the sensor package on the Swarm-Ex mission led by the University of Colorado in Boulder, Colorado, USA. 

IWF/Imperial College Magnetometer: The fluxgate magnetometer design consists of one boom mounted, triaxial sensor with dedicated sensor harness and front-end electronics which connects to the CubeSat’s processor and power supply unit. The sensor, which is provided by Imperial College London, employs fluxgate-type sense technology with ring-cores as centre element. It features low intrinsic noise, high sensitivity and high stability. The sensor will be boom-mounted to avoid interference with the stray fields from the main body of the CubeSat. The magnetometer electronics are composed around an Application Specific Integrated Circuit (ASIC) developed by the Space Research Institute of the Austrian Academy of Sciences. It is an upgrade of the ASIC based electronics which is currently operational aboard the 4-satellite NASA mission Magnetospheric Multiscale and the Korean satellite GEO-KOMPSAT-2A. The new ASIC features a larger dynamic range for full Earth’s field measurements and improved radiation hardness. It is currently under development for the Finish Foresail-2 CubeSat (launch in 2025). A Field Programmable Gate Array (FPGA) will contain data reduction, command interface and the glue logic for the digital interface.

Calgary Thermal Ion Drift Instrument (TIDI): TIDI will provide 3D bulk ion drift at 8 vectors per second per satellite. The TIDI is based on the Swarm electric field instrument thermal ion imaging technique, reduced in size, mass and power for operation on a CubeSat. In this technique, two ram-facing electrostatic ion imagers resolve the rammed ion signal into kinetic energy per unit charge and direction of arrival within their respective orthogonal 2D phase-space planes. The ion bulk drift may be determined under the assumption that the signal is composed of predominantly O+ ions or with further species identification instruments. Twelve such mini plasma imagers have flown on four scientific sounding rockets since 2018. Six of their predecessors, the thermal ion imager, have been in operation on Swarm since late 2013. Along-track electric field can be estimated from the E-cross-B ion drift velocity and either a measured or model magnetic field. From this the cross-polar-cap potential drop can be estimated once per polar pass. The TIDI consists of two identical analysers mounted on the ram face of each satellite, one with a vertically-oriented entrance slit and the other horizontally-oriented. Each analyzer has a volume of approximately 1/3 U, consumes 3.5 W of power, and has a mass of ~350 g. Each TIDI provides a pair of ion images at 8 Hz and housekeeping data at a telemetry rate of 16 kByte/s.

SWRI 3-D Ion velocity and composition Imager (3DI): 3DI can provide additional mass-resolving capability by deriving accurate composition-resolved ion number density, velocity vector, and anisotropic temperatures in all topside ionosphere conditions through delivering full three-dimensional (3D) velocity distribution functions (VDFs) for each major ion species.  The 3DI sensor assembly consists of Electro-Static Analyser (ESA) and deflector sub-system, Lens + Mass Analysis sub-system, and Microchannel Plate and anode sub-system. The ESA offers fine (>1°×3.5°) angular resolution around the ram-facing direction to measure ions precisely even with low (<15 m/s) bulk velocities. The mass-analysis sub-assembly resolves the ions in 4 mass bins: H+, He+, O+, and molecular ions continuously with 100% duty ratio, and focuses the particle to the MCP surface enabling the mass and azimuth angle imaging in a fast cadence. All key technologies have been demonstrated with a working prototype sensor in a relevant vacuum environment using ion beams. The electronics boards draw heritage from previous CubeSat based ESA developed in SwRI. The first flight model is under development in another CubeSat program funded by NASA, and will be launched for performance verifications in late 2025.

Langmuir Probes

Oslo option: A multi-Needle Langmuir Probe (m-NLP) can carry out measurements of the local plasma density at high sampling frequencies (several kHz), thereby allowing detailed investigations of small-scale ionospheric plasma structures. Additionally, it will measure the electron temperature and platform potential at lower sampling rates (several Hz). The m-NLP comprises four cylindrical probes mounted on short (several tens of cm) booms and immersed in the ambient plasma. The bias voltage applied to one probe is swept in order to extract the electron temperature and the spacecraft potential from the collected I-V characteristic curve. The other three probes are biased at different fixed voltages, which allows the collected currents to be measured at a much higher cadence. From the ratio of these currents the electron density can be calculated.  Versions of the m-NLP system are currently in orbit on board the NorSat-1 and BRIK-II satellites, and are about to be installed on the Bartolomeo platform on board the ISS.

IRF option: In its baseline configuration, the LP3 is a 1u unit intended for mounting on a corner of the ram side of the s/c, containing all electronics and sensors. Three sensors are included in this configuration, a planar probe (PLAP) mounted on the ram face of the cube and two thin cylindrical probes (CYLPs) deployed from the other two exposed sides of the cube to a total tip-to-root length of some 25 cm. The PLAP is a circular planar electrode, a few cm in diameter, on the ram side. Ram ions are collected and the resulting current is measured, providing a robust ion density measurement at any desired time resolution down to the millisecond scale, corresponding to meter-sized spatial resolution. The two CYLPs instead are biassed to slightly different positive voltages to measure electron current, providing the density and temperature of the thermal electron gas, also at very high time resolution. Occasional (with configurable intervals of a few seconds to a few minutes) voltage sweeps on the CYLPs provide electron temperature and spacecraft potential at higher accuracy for calibration. 

GNSS Receiver: The GNSS receiver will be based upon the TOPCAT (TOPside ionosphere and plasmasphere Computer Assisted Tomography) receivers developed at the University of Bath. TOPCAT I was a dual-frequency (L1/L2) GPS receiver installed on-board the UK Space Agency’s UKube-1 satellite. The TOPCAT II payload consisted of a GPS receiver (L1, L2 and L5) a payload controller board and a triple-frequency stacked-patch antenna and was installed on the CIRCE mission. The TOPCAT II antenna and low-noise amplifier (LNA) module are powered via a DC bias over a coaxial cable from the TOPCAT II receiver module. Within the antenna, bandpass filters and the LNA provide selection of signals at L1 (centre frequency 1575.42 MHz) and L2/L5 (centre frequencies of 1227.60 MHz and 1176.45 MHz, respectively). The TOPCAT II receiver module has a power consumption of 2.1 W in 10 Hz mode (0.1 s measurement epoch). The antenna is approximately 55 mm by 55 mm by 21 mm tall, with a mass of < 200 g. The interface to TOPCAT II is via a 25-way micro-D connector providing data, telemetry and control via a CCSDS compliant command set.

Northumbria Laser communications 

A compact (2U volume envelope), high-power (2-3W optical power), high-speed (>1.25 GHz bandwidth), low mass (<2kg) laser optical communications terminal, purpose built for inter-satellite ranging or communications at a range of link lengths up to 700km in LEO. The terminal has been designed, tested and prototyped at Northumbria University under the UKSA National Space Innovation Programme (NSIP) in partnership with e2E Group Limited, Durham Centre for Advanced Instrumentation (CfAI), SMS Electronics Ltd and Lockheed Martin UK. The terminal has an operational electronic power requirement of <30W, that can be reduced to <15 W for the ROARS requirements of inter-satellite positioning (no communications) through removal of the data laser amplifier.

The unit consists of: (1) a diamond machined optical head, housing a nested Galilean transmitter telescope and an inverted Newtonian receiver telescope. Both transmitter and receiver telescopes incorporate steerable (up to 400 Hz, +/-1 degree range, piezoelectric motors) mirrors for autonomously achieving a coaxial, closed-loop, duplex, transmitter and receiver channel that allows for stable pointing tracking and data acquisition.  (2) An optical interface electronic board houses two transmitter lasers that are fiber-coupled: one high power (2W) low speed (10MHz) beacon laser (830nm) for link acquisition and one erbidium-amplified (2W) high speed (1.25 GHz) data laser (1550nm) for data transmission. The board also supports the receiver detectors: a custom-built 830nm silicon-based 10x10cm Position Sensing Detector (PSD) and 200micron 1550nm Avalanche PhotoDetector (APD) receiver sensitive to 10µW optical power close to the noise level at 700km link length. (3) A laser driver board supported a Xilinx FPGA driving the high speed signals, adopting On-Off Keying (OOK) modulation, environment monitor, passing control, data and power signals to the steering actuators, managing the experiments library and interfacing with the onboard computer through a PC104 connector. 

Radar Ground Segment: EISCAT_3D: The European Incoherent Scatter Scientific Association’s (EISCAT) EISCAT_3D radar system is a ground-breaking facility that is scheduled to begin operations during Q4 2023. EISCAT_3D represents the first ever attempt to exploit ISR techniques together with phased-array radar technology and multiple receiver sites to yield 3D, volumetric estimates of both scalar and vector quantities. Initially EISCAT_3D will consist of one transceiver site in Skibotn, Norway (69.340° N, 20.315° E) transmitting at 233 MHz with maximum 5 MW Tx power, and two receiver sites located in Kaiseniemi, Sweden (68.267° N, 19.450°E) and Karesuvanto, Finland (68.403° N, 22.524° E). All EISCAT_3D Common Program data will be immediately available to all scientists in countries belonging to the EISCAT Scientific Association (Sweden, Norway, Finland, Japan, China and the United Kingdom). It will also be possible to run special programs optimised for conjunctive studies. For studies of mesoscale (10–1000 km) dynamics in which time resolutions of order 1–10 min are acceptable, it will be possible to configure EISCAT_3D to make volumetric measurements within a relatively large volume (~10⁶ km³). For studies of fine-scale (≤ 10 km) dynamics in which time resolutions of 0.1–1 s are needed, it will be possible to configure EISCAT_3D to use a small number of beams covering a diameter of order 10 km at 100-km altitude.

Satellite Laser Ranging Ground Segment: In addition to the ILRS, the Graz station utilises two different lasers to perform satellite and space debris laser ranging. A 2 kHz laser with 0.8 W and 10 ps pulse length allows to measure the range to cooperative targets up to geostationary orbits with a single shot precision of approx. 3 mm. By utilising a 16 W laser, metre sized uncooperative space debris objects can also be measured up to a maximum range of 1,500 km. A Cassegrain type receive telescope with an aperture of 50 cm collects the reflected photons and redirects them to a single photon avalanche diode detector. Within the detection package, sunlit satellites can be displayed to apply orbit corrections in real time. The high-latitude station is intended to be sited near Sky Blu, a remote blue-ice runway and fuel depot at 75 °S, 72 °W in Antarctica. A feasibility study has been completed and a design study is about to commence to adapt a commercial optical sensor for Antarctic use. Full technical specifications are to be decided but two sensor types are being considered – a satellite laser ranging sensor with cm range precision like the Graz and other ILRS instruments, and a leading SST camera-based system. Operation is expected to align with the ROARS mission schedule, and data will be freely available as required by the Antarctic Treaty.  

Optical Satellite Tracking Ground Segment: The optical brightness of the satellites will depend on several factors, such as their size, surface properties, and range, and will also vary considerably over time with observational geometry. It is likely the satellites will fall in the range 8th – 14th magnitude, well within the sensitivity limits of Warwick’s CLASP and repurposed SuperWASP instruments on La Palma. These are equipped with fast-tracking mounts and sCMOS detectors, enabling them to conduct high-cadence observations of LEO passes when visible from the site. The latter can also perform multi-colour photometry to aid satellite characterisation. The proposed along-track separations will correspond to angular separations of at least 9 degrees on-sky, thus multiple telescopes will be needed to observe contemporaneous passes (the current SuperWASP configuration achieves an 8.5-degree field of view). It may be possible to coordinate with other observatories, such as the University of Bern’s Zimmerwald (46.9 N, 7.5 E) and nodes of the Numerica network (multiple at suitable longitudes).