Professor Craig Underwood graduated from the University of York in 1982 with a BSc in Physics with Computer Science. After gaining a postgraduate certificate in education in 1983, he began a teaching career at Scarborough Sixth-Form College where he developed satellite activities. In January 1986, Craig joined the University of Surrey as a research fellow/engineer developing space education programmes and working on the UoSAT series of spacecraft, where he was responsible for the generation and maintenance of software for the UoSAT Satellite Control Ground-Station, mission analysis, thermal design and radiation environment and effects analysis and mitigation. From 1990 he has been Surrey's Principal Investigator of Space Radiation Environment and Effects, completing his PhD in this area in 1996.
In 1993, Craig became a Lecturer in Spacecraft Engineering advancing to Senior Lecturer in 1999, Reader in April 2003, and Professor in April 2012. Craig was Deputy Director of the Surrey Space Centre from 2007 to 2014 and he currently heads the Environments and Instrumentation Group developing the concepts, instruments and techniques to investigate the Earth and other planetary environments from space.
Craig is author or co-author of some 200 scientific papers and teaches or has taught undergraduate and postgraduate courses on Spacecraft Engineering, Communications Payload Engineering, Satellite Remote Sensing, Planetary Exploration and Astronomy at the University of Surrey.
- Radiation environment and effects (Cosmic Rays, Van Allan Belts, Space Weather, Atmospheric Radiation)
- Remote sensing instrumentation (UV-VIS-NIR multi- and hyper-spectral imaging, Thermal IR imaging, UV-VIS-NIR radiometry and SWIR atmospheric spectroscopy, low-power SAR Radar, GNSS Reflectometry)
- Micro-nano-satellite technologies (SNAP, PalmSat, AAReST, RDV and Docking Systems)• Planetary Exploration (Mars VTOL Aerobot, Ramon Spectroscopy)
- Space power systems (Supercapacitors, Thin-Film Solar PV, RTGs, solar-thermal power).
PhD Topics are available in any of these or related areas.
Take a look at the Group's personnel and research interests.
Craig heads the Environments and Instrumentation Group within the Surrey Space Centre, which has the remit of developing the instruments, systems and data processing techniques needed to investigate the Earth and other planetary environments from space. A particular focus of the group is on the development of low-mass, low-volume and low-power “micro-instrumentation” suitable for use on micro/nano-satellite technology platforms. Current research activities include the analysis of the space and atmospheric radiation environments and their effects on commercial-off-the-shelf (COTS) avionic technologies; the development of miniaturised instrumentation for ionising-radiation detection, UV-VIS-NIR and thermal-IR satellite remote sensing; micro-satellite-based radar imaging; a Mars VTOL unpopulated micro-air-vehicle, and micro-nano-satellite technologies.
Radiation environment and effects
over the past 25 years Craig has gained considerable expertise in understanding the space radiation environment and its effects in low, medium and high Earth orbit (out to 60,000 km). The deleterious effects of the ionising radiation environment is of particular concern when using COTS technologies in space, thus, particular emphasis has been given to a programme of monitoring “space weather” in terms of the high energy proton and heavy-ion cosmic-ray environment these spacecraft encounter, and to observing and analysing its effects - particularly with regard to single-event effects - upon the COTS devices on-board. The extended period of research has enabled a wide variety of conditions to be observed ranging across an entire solar cycle, and standard models to be verified or challenged. He was the first to show clearly the effect of the SAA trapped proton environment on COTS memories operating in LEO (UoSAT-2), and through his work with QinetiQ using the QinetiQ's CREDO (UoSAT-3) and his own CRE (KITSAT-1, PoSAT-1) instruments has shown the limitations of the AP8 and CREME models. He is now performing similar work for the MEO environment through the analysis of flight data from Surrey's CEDEX and QinetiQ's MERLIN payloads flown on GIOVE-A. It has recently delivered two miniaturised radiation monitors (MuREM, RM) for the UK's TechDemoSat-1 mission, launched in 2014. These payloads comprise solid-state (RadFET) dosimeters, ionizing dose-rate-diode detectors, and PIN-diode -based multi-channel analysers for measuring proton and heavy-ion LET spectra. The work has also been applied to the aerospace sector via the SPACERANE project.
Remote sensing instrumentation optical
Craig has had a long-term interest in remote sensing instrumentation design: He developed stratospheric ozone monitoring UV radiometers for the FASAT-Alfa (1995) and FASAT-Bravo (1998) satellites, and an ultra-compact Earth-observation CMOS video camera for the Thai Paht (1998) satellite. He also provided the pre-flight optical and radiometric calibration of the tri-band (NIR, Red, Green) imaging sensors for the Disaster Monitoring Constellation (DMC) Satellites: AlSAT-1 (2002, UK-DMC (2003) and NigeriaSat (2003).With his PhD students, Craig has developed prototype designs for a micro-bolometer array-based thermal-IR imager (B. Olerich, 2005) and for a UV spectrometer for monitoring volcanic plumes (SO2) and ozone (J. Fernandez-Saldivar, 2008). He has a strong interest in the application of Spatial Heterodyne Spectroscopy (SHS) and, through PhD studies, has applied this technique to an ultra-compact Ramon spectrometer for the analysis of Martian rocks (T. Nathanial,2011) and to the SWIR detection and measurement of atmospheric CO2 (I. Ikpaya, 2013). He is also interested in compact hyperspecteral instruments for micro-sat and UAV application.
He has proposed a bistatic Synthetic Aperture Radar (SAR) imaging concept for micro-satellites (2000) and, through PhD programmes, has developed the concept of applying low-power CWFM bi-static and mono-static SAR to micro-sat platforms (~100-150kg) (O. Mitchell, 2001; T. Wanwiwake, 2011; N. Ahmed, 2012; A. Cai, 2013). He also worked on the airborne demonstrator for the NovaSAR S-Band SAR (2010).He has also supported PhD research into GNSS Reflectometry (P. Jales, 2013; E. Simons, 2014; J. Tye); Image Data Compression (P. Hou, 1999); Machine Vision for Pose and Relative Orbit estimation (A. Cropp, 2001); and working with the National Physical Laboratory is researching Vicarious Calibration/Validation of Remote Sensing Instruments and Radiometric Uncertainty modelling (A. Bialek, J. Gorrono).Micro-Nano-Satellite Technologies.
Craig began Surrey's nano-satellite activities in 1995, through setting and supervising a series of student projects aimed at developing a "soccer ball" sized spacecraft. As Chief Architect of the SNAP concept, he played a pioneering role in developing the UK's first operational nano-satellite, SNAP-1, Surrey's 6.5 kg nano-satellite, launched in June 2000, which carried out experiments in autonomous orbital manoeuvring and remote inspection of other spacecraft. For his work on SNAP, Craig and the Surrey Space Centre achieved the award of “Finalist” in the 2001 Flight International Awards in the Space and Missiles Category. He subsequently developed the PalmSat, ~1kg pico-satellite concept in 2000, designed to play a similar rôle, and he is currently the UK PI for the AAReST multiple-mirror space telescope demonstrator concept, working with CalTech/NASA-JPL, where he is also developing a novel electro-magnetic rendezvous and docking system. AAReST is designed to demonstrate the autonomous in-orbit construction of a space telescope using multiple-mirror elements, which can change shape to form a coherent optical surface. Craig has also worked on Super-Capacitor based power systems (T. Shimizu, 2013); Thin-film solar PV systems; data-handling and RF systems (V.Asenek, 1998; S. Maqbool, 2006).Planetary Exploration
Away from orbit, Craig is working on a vertical take-off and landing (VTOL) “flying wing” aerobot concept for the exploration of Mars (J. Fielding, 2004; H. Song, 2008; W. Zhao, 2013; N. Collins, current). He has also worked on studies for Lunar microsatellite missions, spaceborne radio astronomy, Mars sample returns and NEO investigations.
Over the last 30 years, Craig has played a key role in developing and teaching Surrey's spacecraft engineering postgraduate, undergraduate and industrial-training courses. He was the recipient of the Department of Electrical and Electronic Engineerings Tony Jeans Inspirational Teaching Prize, 2013. Currently he teaches:
- Level 1 (FHEQ 4) Mathematics
- Level 2 (FHEQ 5) Space Engineering and Mission Design
- Level 3 (FHEQ 6) Space Systems Design
- Level M (FHEQ 7) Spacecraft Systems Design
- Level M (FHEQ 7) Launch Vehicles and Propulsion
- Level M (FHEQ 7) Space Environment and Protection
- Short Course: Spacecraft Systems Design.
The Surrey Training Research and Nanosatellite Demonstrator (STRaND) programme has been success in identifying and creating a leading low-cost nanosatellite programme with advanced attitude and orbit control system (AOCS) and experimental computing platforms based on smart-phone technologies. The next demonstration capabilities, that provide a challenging mission to the existing STRaND platform, is to perform visual inspection, proximity operations and nanosatellite docking. Visual inspection is to be performed using a COTS LIDAR system to estimate range and pose under 100 m. Proximity operations are controlled using a comprehensive guidance, navigation and control (GNC) loop in a polar form of the Hills Clohessy Wiltshire (HCW) frame including J2 perturbations. And finally, nanosatellite docking is performed at under 30 cm using a series of tuned magnetic coils. This paper will document the initial experiments and calculations used to qualify LIDAR components, size the mission thrust and tank requirements, and air cushion table demonstrations of the docking mechanism.
Optical earth observation (EO) satellite sensors generally suffer from drifts and biases relative to their pre-launch calibration, caused by launch and/or time in the space environment. This places a severe limitation on the fundamental reliability and accuracy that can be assigned to satellite derived information, and is particularly critical for long time base studies for climate change and enabling interoperability and Analysis Ready Data. The proposed TRUTHS (Traceable Radiometry Underpinning Terrestrial and Helio-Studies) mission is explicitly designed to address this issue through re-calibrating itself directly to a primary standard of the international system of units (SI) in-orbit and then through the extension of this SI-traceability to other sensors through in-flight cross-calibration using a selection of Committee on Earth Observation Satellites (CEOS) recommended test sites. Where the characteristics of the sensor under test allows, this will result in a significant improvement in accuracy. This paper describes a set of tools, algorithms and methodologies that have been developed and used in order to estimate the radiometric uncertainty achievable for an indicative target sensor through in-flight cross-calibration using a well-calibrated hyperspectral SI-traceable reference sensor with observational characteristics such as TRUTHS. In this study, Multi-Spectral Imager (MSI) of Sentinel-2 and Landsat-8 Operational Land Imager (OLI) is evaluated as an example, however the analysis is readily translatable to larger-footprint sensors such as Sentinel-3 Ocean and Land Colour Instrument (OLCI) and Visible Infrared Imaging Radiometer Suite (VIIRS). This study considers the criticality of the instrumental and observational characteristics on pixel level reflectance factors, within a defined spatial region of interest (ROI) within the target site. It quantifies the main uncertainty contributors in the spectral, spatial, and temporal domains. The resultant tool will support existing sensor-to-sensor cross-calibration activities carried out under the auspices of CEOS, and is also being used to inform the design specifications for TRUTHS.
In the framework of the European Copernicus programme, the European Space Agency (ESA) has launched the Sentinel-2 (S2) Earth Observation (EO) mission which provides optical high spatial -resolution imagery over land and coastal areas. As part of this mission, a tool (named S2-RUT, from Sentinel-2 Radiometric Uncertainty Tool) estimates the radiometric uncertainties associated to each pixel using as input the top-of-atmosphere (TOA) reflectance factor images provided by ESA. The initial version of the tool has been implemented — code and user guide available1 — and integrated as part of the Sentinel Toolbox. The tool required the study of several radiometric uncertainty sources as well as the calculation and validation of the combined standard uncertainty in order to estimate the TOA reflectance factor uncertainty per pixel. Here we describe the recent research in order to accommodate novel uncertainty contributions to the TOA reflectance uncertainty estimates in future versions of the tool. The two contributions that we explore are the radiometric impact of the spectral knowledge and the uncertainty propagation of the resampling associated to the orthorectification process. The former is produced by the uncertainty associated to the spectral calibration as well as the spectral variations across the instrument focal plane and the instrument degradation. The latter results of the focal plane image propagation into the provided orthoimage. The uncertainty propagation depends on the radiance levels on the pixel neighbourhood and the pixel correlation in the temporal and spatial dimensions. Special effort has been made studying non-stable scenarios and the comparison with different interpolation methods
Millimeter wave (mm-wave) bands are becoming potentially attractive candidates for next generation communication systems. It is envisioned that high gain smart antennas will be one of the key enabling technologies for such systems. At mm-wave bands, where electrical size of an individual antenna becomes very small, the inclusion of a reconfigurable mechanism in the antenna becomes a great challenge due to real estate constraints. In these scenarios a designer has to decide on the number of bits in a phase shifter for antenna beam steering which will result in an optimum design. This contribution addresses the issue of phase quantization in mm-wave high gain reflectarray smart antennas to achieve an optimum performance. Implementing coarse phase quantization greatly reduces the complexity at mm-wave bands. A case study is presented to highlight the effects of coarse phase quantization using various numbers of bits.
In the framework of the European Copernicus programme, the European Space Agency (ESA) has launched the Sentinel-2 (S2) Earth Observation (EO) mission which provides optical high spatial resolution imagery over land and coastal areas. As part of this mission, a tool (named S2-RUT, from Sentinel-2 Radiometric Uncertainty Tool) has been developed. The tool estimates the radiometric uncertainty associated with each pixel in the top-of-atmosphere (TOA) reflectance factor images provided by ESA. This paper describes the design and development process of the initial version of the S2-RUT tool. The initial design step describes the S2 radiometric model where a set of uncertainty contributors are identified. Each of the uncertainty contributors is specified by reviewing the preand post-launch characterisation. The identified uncertainty contributors are combined following the guidelines in the ‘Guide to Expression of Uncertainty in Measurement’ (GUM) model and this combination model is further validated by comparing the results to a multivariate Monte Carlo Method (MCM). In addition, the correlation between the different uncertainty contributions and the impact of simplifications in the combination model have been studied. The software design of the tool prioritises an efficient strategy to read the TOA reflectance factor images, extract the auxiliary information from the metadata in the satellite products and the codification of the resulting uncertainty image. This initial version of the tool has been implemented and integrated as part of the Sentinels Application Platform (SNAP).
This paper studies the distribution associated with the measurement of the satellite derived Top-Of-Atmosphere (TOA) reflectance on a pixel-to-pixel level, within a defined spatial region of interest (ROI) within a vicarious calibration target site. The study analyses the effects of the atmosphere and surface reflectance distribution spatial shape. The analysis shows that some of the contributing effects are inherently non-linear, so produce non-normal distributions. For these non-normal distributions, the use of the mean and standard deviation alone does not allow sufficient parameterisation of the distribution to capture all the information associated with the ROI reflectance measurement. Therefore, additional information concerning the distribution is required to provide a full site reflectance characterisation. This additional information can be useful in establishing the sources of change in the distribution and ultimately improve the site radiometric characterisation, particularly for long term monitoring. In this study LandSat-8 (L8) Operational Land Imager (OLI) measurements over the CEOS Libya-4 Pseudo Invariant Calibration Site (PICS) are used as a demonstration.
Future space telescopes with diameter over 20 m will require new approaches: either high-precision formation flying or in-orbit assembly. We believe the latter holds promise at a potentially lower cost and more practical solution in the near term, provided much of the assembly can be carried out autonomously. To gain experience, and to provide risk reduction, we propose a combined micro/nano-satellite demonstration mission that will focus on the required optical technology (adaptive mirrors, phase-sensitive detectors) and autonomous rendezvous and docking technology (inter-satellite links, relative position sensing, automated docking mechanisms). The mission will involve two "3U" CubeSat-like nanosatellites ("MirrorSats") each carrying an electrically actuated adaptive mirror, and each capable of autonomous un-docking and re-docking with a small central "15U" class micro/nano-satellite core, which houses two fixed mirrors and a boom-deployed focal plane assembly. All three spacecrafts will be launched as a single ~40 kg micro-satellite package. The spacecraft busses are based on heritage from Surrey's SNAP-1 and STRaND-1 missions (launched in 2000 and 2013 respectively), whilst the optics, imaging sensors and shape adjusting adaptive mirrors (with their associated adjustment mechanisms) are provided by CalTech/JPL. The spacecraft busses provide precise orbit and attitude control, with inter-satellite links and optical navigation to mediate the docking process. The docking system itself is based on the electromagnetic docking system being developed at the Surrey Space Centre (SSC), together with rendezvous sensing technology developed for STRaND-2. On orbit, the mission profile will firstly establish the imaging capability of the compound spacecraft before undocking, and then autonomously re-docking a single MirrorSat. This will test the docking system, autonomous navigation and system identification technology. If successful, the next stage will see the two MirrorSat spacecraft undock and re-dock to the core spacecraft in a linear formation to represent a large (but sparse) aperture for high resolution imaging. The imaging of stars is the primary objective, but other celestial and terrestrial targets are being considered. Teams at CalTech and SSC are currently working on the mission planning and development of space hardware. The autonomous rendezvous and docking system is currently under test on a 2D air-bearing table at SSC, and the propulsion and precision attitude control system is currently in development. Launch is planned for 2016. This paper details the mission concept; technology involved and progress to date, focussing on the spacecraft buses.
This paper presents the design and development of a dual-band switched-beam microstrip array for Global Navigation Satellite System (GNSS) applications such as ocean reflectometry and remote sensing. In contrast to the traditional Butler matrix, a simple, low cost, broadband and low insertion loss beam switching feed network is proposed, designed and integrated with a dual band antenna array to achieve continuous beam coverage of ±25° around the boresight at the L1 (1.575 GHz) and L2 (1.227 GHz) bands. To reduce the cost, microstrip lines and PIN diode based switches are employed. The proposed switched beam network is then integrated with dual-band step-shorted annular ring (S-SAR) antenna elements in order to produce a fully integrated compact-sized switched beam array. Antenna simulation results show that the switched beam array achieves a maximum gain of 12 dBic at the L1 band and 10 dBic at the L2 band. In order to validate the concept, a scaled down prototype of the simulated design is fabricated and measured. The prototype operates at twice of the original design frequency i.e. 3.15 GHz and 2.454 GHz and the measured results confirm that the integrated array achieves beam switching and good performance at both bands.
Two new miniaturized scientific radiation monitoring payloads are presented prior to their first flight on the TechDemoSat-1 Spacecraft. They are capable of monitoring the space radiation environment and its effects on radiation-sensitive devices. Micro radaion environment monitor (MuREM) and Surrey Satellite Technology radiation monitor (SSTL RM) carry RADFET dosimeters, dose-rate-sensitive photodiodes, and p-i-n diode particle detectors. SSTL RM is also connected to external RADFET sensors placed around the spacecraft, while MuREM carries a radiation effects payload consisting of COTS devices that will be monitored while exposed to the space radiation environment. © 2012 IEEE.
Reflectarrays are becoming a potentially attractive replacement of parabolic reflectors for high gain requirements. A large reflectarray consists of thousands of elements. To predict their performance a simulation model is required which is very cumbersome to build manually due to a large number of elements. It takes exhaustive efforts, keen attention to details and significant amount of time to build such a simulation model. When several iterations of modelling are required it worsens the issue even further. We have presented here an algorithm as an automated solution to this problem by interfacing Matlab® with an electromagnetic simulation software. It is very generic, time efficient and makes the modelling easy with least intervention of the designer.
As part of the Sentinel-2 mission, a Radiometric Uncertainty Tool (RUT) has been recently released to the community. This tool estimates the Sentinel-2 radiometric uncertainty associated with each pixel in the top-of-atmosphere (TOA) reflectance factor images provided by the European Space Agency (ESA). The use of such information enables users to assess the “fitness for purpose” of the data to their specific application. The work described here summarises the efforts and results of integrating the RUT for radiometric validation activities for the Sentinel-2 mission. Starting from the results provided by the RUT, the focus will be on providing a methodology to calculate the uncertainty associated with the mean TOA reflectance factor in a Region of Interest (ROI). Two different methods – one simple method directly using the RUT and a more rigorous one based on Monte Carlo method (MCM) propagation – are proposed and compared. These two methods focus on the effect of the spectral, spatial and temporal correlation of the errors in different ROI pixels and the impact of correlation on the uncertainty associated with the mean TOA reflectance factor. The study has also considered the impact of uncertainty contributions not included in the first version of the RUT.
Proximity flight systems for rendezvous-and-docking, are traditionally the domain of large, costly institutional manned missions, which require extremely robust and expensive Guidance Navigation and Control (GNC) solutions. By developing a low-cost and safety compliant GNC architecture and design methodology, low cost GNC solutions needed for future missions with proximity flight phases will have reduced development risk, and more rapid development schedules. This will enable a plethora of on-orbit services to be realised using low cost satellite technologies, and lower the cost of the services to a point where they can be offered to commercial as well as institutional entities and thereby dramatically grow the market for on-orbit construction, in-orbit servicing and active debris removal. It will enable organisations such as SSTL to compete in an area previously exclusive to large institutional players. The AAReST mission (to be launched in 2018), will demonstrate some key aspects of low cost close proximity “co-operative” rendezvous and docking (along with reconfiguration/control of multiple mirror elements) for future modular telescopes. However this is only a very small scale academic mission demonstration using cubesat technology, and is limited to very close range demonstrations. This UK National Space Technology Programme (NSTP-2) project, which is being carried out by SSTL and SSC, is due to be completed by the end of November 2017 and is co-funded by the UK Space Agency and company R&D. It is aiming to build on the AAReST ("Autonomous Assembly of a Reconfigurable Space Telescope") mission (where appropriate), and industrialise existing research, which will culminate in a representative model that can be used to develop low-cost GNC solutions for many different mission applications that involve proximity activities, such as formation flying, and rendezvous and docking. The main objectives and scope of this project are the following: Definition of a reference mission design (based on a scenario that SSTL considers credible as a realistic scenario) and mission/system GNC requirements. Develop a GNC architectural design for low cost missions applications that involve close proximity formation flying, rendezvous and docking (RDV&D) - i.e. “proximity activities” Develop a low cost sensor suite suitable for use on proximity missions Consider possible regulatory constraints that may apply to the mission The SSTL/SSC reference mission concept is a “co-operative” two-spacecraft rendezvous and docking mission demonstrator using microsatellites (an active Chaser and a passive Target), however the GNC model is generic and can be utilized for other “non-co-operative” rendezvous and docking missions. This paper presents the latest results from the study, particularly the mission analysis, GNC simulation and modelling, sensors, and key mission and spacecraft systems aspects. The results so far show that such a GNC model and mission demonstrator is feasible, and in line with anticipated UK regulatory constraints that may apply to the mission.
The InflateSail (QB50-UK06) CubeSat, designed and built at the Surrey Space Centre (SSC) for the Von Karman Institute (VKI), Belgium, was one of the technology demonstrators for the European Commission’s QB50 programme. The 3.2 kg 3U CubeSat was equipped with a 1 metre long inflatable mast and a 10m2 deployable drag sail. InflateSail's primary mission was to demonstrate the effectiveness of using a drag sail in Low Earth Orbit (LEO) to dramatically increase the rate at which satellites lose altitude and re-enter the Earth's atmosphere and it was one of 31 satellites that were launched simultaneously on the PSLV (polar satellite launch vehicle) C-38 from Sriharikota, India on 23rd June 2017 into a 505km, 97.44o Sun-synchronous orbit. Shortly after safe deployment in orbit, InflateSail automatically activated its payload. Firstly, it inflated its metrelong metal-polymer laminate tubular mast, and then activated a stepper motor to extend four lightweight bi-stable rigid composite (BRC) booms from the end of the mast, so as to draw out the 3.1m x 3.1m square, 12m thick polyethylene naphthalate (PEN) drag-sail. As intended, the satellite immediately began to lose altitude, causing it to re-enter the atmosphere just 72 days later – thus successfully demonstrating for the first time the de-orbiting of a spacecraft using European inflatable and drag-sail technologies. The InflateSail project was funded by two European Commission Framework Program Seven (FP7) projects: DEPLOYTECH and QB50. DEPLOYTECH had eight European partners including DLR, Airbus France, RolaTube, Cambridge University, and was assisted by NASA Marshall Space Flight Center. DEPLOYTECH’s objectives were to advance the technological capabilities of three different space deployable technologies by qualifying their concepts for space use. QB50 was a programme, led by VKI, for launching a network of 50 CubeSats built mainly by university teams all over the world to perform first-class science in the largely unexplored lower thermosphere. The boom/drag-sail technology developed by SSC will next be used on a third FP7 Project: RemoveDebris, launched in 2018, which will demonstrate the capturing and de-orbiting of artificial space debris targets using a net and harpoon system. This paper describes the results of the InflateSail mission, including the observed effects of atmospheric density and solar activity on its trajectory and body dynamics. It also describes the application of the technology to RemoveDebris and its potential as a commercial de-orbiting add-on package for future space missions.
Most satellites, currently in use, don't incorporate any post mission disposal system and therefore will end up as debris once they reach the end of their life. Without any intervention, these objects could endanger future space missions, once their density is high enough. In this paper the authors propose a new concept for an uncontrolled removal system based on electro dynamical principles. Instead of long flexible tethers (which have proven problematic to deploy), we consider using relatively short (~150m-300m) rigid electro-dynamic booms. The main advantage of such a structure is that, for satellites in polar orbits, it leads to a larger Lorentz force. Also, the deployment is more reliable and the attitude control is greatly simplified (because the booms are rigid). A ground demonstrator is under development based around a 6U CubeSat structure. We also look at different techniques which could be used for electron emission into the surrounding plasma because currently this is what limits the generated currents in the proposed system. This work is conducted as a part of the European Commission funded Horizon-2020 TeSeR (Technology for Self-Removal) project, which aims to demonstrate the feasibility of a scalable post mission removal system which should be able to be connected to different satellites via a standard interface.
One major source of new space debris are spacecraft (S/C) that are not removed from orbit after the end of their operational lifetime. Many regulations (e.g. ISO 24113) require the removal of S/C at the end of operation - known as Post-Mission-Disposal (PMD) - with a compliance rate of 90% to ensure that S/C do not become a new source of space debris. An analysis performed by ESA shows that the success rate of PMD in 2013 was in the range of about 50%-60%. The goal of TeSeR (Technology for Self-Removal) is to take the first step towards the development of a costefficient, but highly reliable PMD module. This PMD module is to be attached to the S/C on ground and it shall ensure the PMD of the S/C at the end of the operational lifetime. This PMD module shall be scalable and flexible, thus, enabling the PMD of any future S/C in an Earth orbit. Ultimately, the gap between the required 90% PMD success rate and the current success rate can be closed. The technological enhancements and developments required for successful PMD are addressed and analysed in TeSeR. The project’s primary aims are to develop, manufacture and test an on-ground prototype of the PMD module, to develop three different removal subsystems (solid propulsion, electro-dynamical systems and deployable structures) for easy plug-in/plug-out implementation to the PMD module. This is the first step to demonstrate the main aspects of such a PMD module and the required main technologies. The technical activities are supported by non-technical tasks, e.g. investigation of legal issues relating to a PMD module, execution of a market study and consideration of this technology as a leverage to advance ISO norms. This double tracked approach ensures that the technological developments are embedded into the needs of the space community right from the start. Up to now the prototypes of the three removal subsystems have been developed, manufactured and tested with a common interface for implementation into the PMD module prototype. The PMD module prototype will be manufactured until summer 2018. Afterwards the removal subsystems will be integrated via the same interface. Airbus is the coordinator (and potential launch customer) of TeSeR. The project is conducted together with 10 notable institutes and companies from all across Europe with experts who have been working in the space debris issue for many years.
Climate change is of increasing concern and efforts to mitigate its effects are targeted on reducing fossil CO2 emissions. Satellite observations play a key role in the understanding and management of the problem. Whilst detecting CO2 optically is relatively straight-forward, and has been achieved with small satellites, accurate quantitative mapping of CO2 requires very high precision (300:1). This normally requires high performance, large and complex instruments whose high cost, mass, volume, and power requirements preclude their use on small satellites. This paper presents the developmental stage of a single channel (1.6 μm) compact precision Spatial Heterodyne Atmospheric Carbon-Dioxide Spectrometer (SHACS), which utilises the Spatial Heterodyne Spectrometer (SHS) technique to form a robust, compact, no-moving-part Fourier Transform Spectrometer (FTS). This instrument achieves a high spectral resolution of 0.25cm-1 at a high SNR of >700:1 and can fit into a micro-satellite platform. With this performance, high quality measurements of atmospheric CO2 concentration with measurement precision of
The implementation of a viable Synthetic Aperture Radar (SAR) mission using small satellites faces significant technological and financial challenges, and this paper evaluates how small such a spacecraft could be made whilst still fulfilling a useful mission. SAR offers a range of complementary capabilities alongside other Earth Observation systems with various unique features, but developing such spacecraft has traditionally been expensive and technologically challenging. It is only in the most recent years that small satellite SAR missions have been implemented and operated, and this paper examines the state of the art and the challenges. Furthermore the opportunities of how small SAR satellites can help realise new Earth Observation capabilities not available on existing traditional SAR satellites are described using examples of missions under development or reference design missions.
Persistent monitoring of large areas using spaceborne Synthetic Aperture Radar (SAR) is a challenging problem for various defence and civil applications. The PASSAT project was proposed and undertaken by the University of Birmingham, under the sponsorship of the UK Defence Science and Technology Laboratory, to analyse the concept of a fully passive (receive only) spaceborne SAR system based on a constellation of CubeSats. By making use of terrestrial transmitters (e.g. Digital Video Broadcasting – Terrestrial (DVB-T) or similar transmitters of opportunity), the problem of having to carry a high power pulsed radar transmitter on the satellite is eliminated. Instead, the satellite only need carry a suitable receiver, antenna and signal storage facility. It is expected that such a system would provide imaging of populated areas with a potential resolution of ~2-3 m. In this paper, we describe progress towards the design of such a system, including the results of a series of ground-based and airborne trials which make use of DVB-T transmissions from the Sutton Coldfield transmitter. In the processed images, roads, wind turbines, electricity pylons, hedgerows and trees are all clearly identified.
This article reports two contributions related to reflectarray antenna design at millimeter waves (mm-waves). First, a closed form analytical formulation is provided for the prediction of reflection properties of square/rectangular mm-waves reflectarray unit cells based on various quality factors and the theory of waveguide coupled resonators. To ensure a high accuracy at mm-waves, the effects of fringing fields, surface waves, metal conductivity, and metal surface roughness are included in the analysis. This analysis program greatly facilitates the parametric studies of a unit cell's constituting parameters to converge on an optimum design solution. Secondly, the concept of phase quantization is proposed for a cost effective realization of mm-waves reflectarrays. The developed formulation in the first contribution was used to design two 3 bit phase quantized, single layer, 19 wavelength, passive reflectarrays at 60 GHz. The test results are compared with simulations and a very good agreement was observed. These findings are potentially useful for the realization of high gain antennas for mm-wave inter-satellite links in small satellite platforms.
In this paper, we propose the use of a novel fixed-wing vertical take-off and landing (VTOL) aerobot. A mission profile to investigate the Isidis Planitia region of Mars is proposed based on the knowledge of the planet's geophysical characteristics, its atmosphere and terrain. The aerobot design is described from the aspects of vehicle selection, its propulsion system, power system, payload, thermal management, structure, mass budget, and control strategy and sensor suite. The aerobot proposed in this paper is believed to be a practical and realistic solution to the problem of investigating the Martian surface. A six-degree-of-freedom flight simulator has been created to support the aerobot design process by providing performance evaluations. The nonlinear dynamics is then linearized to a state-space formulation at a certain trimmed equilibrium point Basic autopilot modes are developed for the aerobot based on the linearized state-space model. The results of the simulation show the aerobot is stable and controllable.
The InflateSail (QB50-UK06) CubeSat, designed and built at the Surrey Space Centre (SSC) for the Von Karman Institute (VKI), Belgium, was a technology demonstrator built under the European Commission’s QB50 programme. The 3.2 kilogram 3U CubeSat was equipped with a 1 metre long inflatable mast and a 10m2 deployable drag sail and was one of 31 satellites that were launched simultaneously on the PSLV (polar satellite launch vehicle) C-38 from Sriharikota, India on 23rd June 2017 into a 505km, 97.44o Sun-synchronous orbit. Shortly after insertion into orbit, InflateSail automatically activated its drag-sail payload, and, as planned, began to lose altitude, causing it to re-enter the atmosphere just 72 days later – successfully demonstrating for the first time the de-orbiting of a spacecraft using European inflatable and drag-sail technologies. This paper discusses the dynamics we observed during the descent, including the sensitivity of the craft to atmospheric density changes. The InflateSail project was funded by two European Commission Framework Program Seven (FP7) projects: DEPLOYTECH and QB50. QB50 was a programme, led by VKI, for launching a network of 50 CubeSats built mainly by university teams all over the world to perform first-class science in the largely unexplored lower thermosphere.
Understanding the lunar near-surface distribution of relevant in-situ resources, such as ilmenite (FeTiO3), and volatiles, such as water/ice, is vital to future sustained manned bases. VMMO is a highly-capable, low-cost 12U Cubesat designed for operation in a lunar frozen orbit. It accomodates the LVMM Lunar Volatile and Mineralogy Mapper and the CLAIRE Compact LunAr Ionising Radiation Environment payloads. LVMM is a multi-wavelength Chemical Lidar using fiber lasers emitting at 532nm and 1560nm, with an optional 1064nm channel, for stand-off mapping of the lunar ice distribution using active laser illumination, with a focus on the permanently-shadowed craters in the lunar south pole. This combination of spectral channels can provide sensitive discrimination of water/ice in various regolith. The fiber-laser technology has heritage in the ongoing Fiber Sensor Demonstrator flying on ESA's Proba-2. LVMM can also be used in a low-power passive mode with an added 280nm UV channel to map the lunar mineralogy and ilmenite distribution during the lunar day using the reflected solar illumination. CLAIRE is designed to provide a highly miniaturized radiation environment and effect monitor. CLAIRE draws on heritage from the MuREM and RM payloads, flown on the UK’s TDS-1 spacecraft. The payload includes PIN-diode sensors to measure ionizing particle fluxes (protons and heavy-ions) and to record the resulting linear energy transfer (LET) energy-deposition spectra. It also includes solid-state RADFET dosimeters to measure accumulated ionizing dose, and dose-rate diode detectors, designed to respond to a Coronal Mass Ejection (CME) or Solar Particle Event (SPE). CLAIRE also includes an electronic component test board, capable of measuring SEEs and TID effects in a selected set of candidate electronics, allowing direct correlations between effects and the real measured environment.
Persistent monitoring of large areas using spaceborne Synthetic Aperture Radar (SAR) is a challenging problem for various defence and civil applications. Despite the fact that spaceborne SAR from low Earth orbit (LEO) is a welldeveloped technology, in practice it cannot provide persistent monitoring of any particular geographical region, as any single satellite has a rather long revisit time. Geostationary Earth Orbit (GEO) SAR missions have been proposed, but here there are major engineering issues due the severe path loss across the distances involved. Indeed, path loss is even more severe in radar systems than it is in radio communications. To provide persistent (or near persistent) monitoring from LEO, a very large number of satellites (~100) would be required to detect short-lived events. However, even though such a solution may be technically possible, a satellite constellation development of this scale may not be economically viable. The PASSAT project was proposed and undertaken by the University of Birmingham, under the sponsorship of the UK Defence Science and Technology Laboratory, to analyse the concept of a fully passive (receive only) spaceborne SAR system based on a constellation of microsatellites. By making use of terrestrial transmitters (we propose to use ground-based broadcasting systems, i.e. DVB-T, DAB, FM radio and similar as transmitters of opportunity), the problem of having to carry a high power pulsed radar transmitter on a microsatellite is eliminated. Instead, the satellite only need carry a suitable receiver, antenna and signal storage facility. It is expected that such a system will: (i) provide imaging of a monitored area with a potentially achievable resolution of 2-3 m in either direction; (ii) cover mainly populated parts of the Earth and, partly, littoral waters; (iii) its costs will be orders of magnitude less in comparison to an equivalent active spaceborne SAR constellation. In addition we may expect more information-rich images, as we are dealing with a multi-static, multi-frequency (VHF/UHF) system which effectively has no equivalent at present. In this paper, we report the results of a series of ground-based and airborne trials of the system, around Birmingham, Coventry and Bruntingthorpe Airfield, which make use of DVB-T transmissions from the Sutton Coldfield transmitter at ranges up to 46km. In the processed images, roads, wind turbines, hedgerows and trees are all clearly identified. We also discuss a proposed spaceborne demonstrator, based on a 12U CubeSat platform with a deployable high-gain UHF helical antenna
This work describes progress towards achieving a flexible, high specific power and low-cost photovoltaic (PV) for emerging large area space applications. The study reports the highest conversion efficiency of 15.3% AM1.5G for a CdTe device on ultra-thin cerium-doped cover glass, the standard protective material for extra-terrestrial PVs. The deposition technique used for all of the semiconductor layers comprising the device structure was atmospheric pressure metal organic chemical vapour deposition. Improvements to the device structure over those previously reported led to a Voc of 788 mV and a relatively low series resistance of 3.3 Ω·cm2. These were largely achieved by the introduction of a post-growth air anneal and a refinement of the front contact bus bars, respectively. The aluminium-doped zinc oxide transparent conductive oxide, being the first layer applied to the cover glass, was subject to thermal shock cycling +80 to (-) 196°C to test the adhesion under the extreme conditions likely to be encountered for space application. Scotch Tape testing and sheet resistance measurements before and after the thermal shock testing demonstrated that the aluminium-doped zinc oxide remained well adhered to the cover glass and its electrical performance unchanged.
Persistent monitoring of large areas using spaceborne Synthetic Aperture Radar (SAR) is a challenging problem for various defence and civil applications. Despite the fact that spaceborne SAR from low Earth orbit (LEO) is a well-developed technology, in practice it cannot provide persistent monitoring of any particular geographical region, as any single satellite has a rather long revisit time. Geostationary Earth Orbit (GEO) SAR missions have been proposed, but here there are major engineering issues due the severe path loss across the distances involved. Indeed, path loss is even more severe in radar systems than it is in radio communications. To provide persistent (or near persistent) monitoring from LEO, a very large number of satellites (~100) would be required to detect short-lived events. However, even though such a solution may be technically possible, a satellite constellation development of this scale may not be economically viable. The PASSAT project was proposed and undertaken by the University of Birmingham, under the sponsorship of the UK Defence Science and Technology Laboratory, to analyse the concept of a fully passive (receive only) spaceborne SAR system based on a constellation of microsatellites. By making use of terrestrial transmitters (we propose to use ground-based broadcasting systems, i.e. DVB-T, DAB, FM radio and similar as transmitters of opportunity), the problem of having to carry a high power pulsed radar transmitter on a microsatellite is eliminated. Instead, the satellite only need carry a suitable receiver, antenna and signal storage facility. It is expected that such a system will: (i) provide imaging of a monitored area with a potentially achievable resolution of 2-3 m in either direction; (ii) cover mainly populated parts of the Earth and, partly, littoral waters; (iii) its costs will be orders of magnitude less in comparison to an equivalent active spaceborne SAR constellation. In addition we may expect more information-rich images, as we are dealing with a multi-static, multi-frequency (VHF/UHF) system which effectively has no equivalent at present. In this paper, the emphasis is on the PASSAT concept, the space segment investigation and the experimental results of passive SAR imaging with DVB-T transmissions undertaken at the University of Birmingham using a local DVB-T transmitter.
Global navigation satellite system reflectometry (GNSS-R) has found many applications in the field of Earth observation including ocean wind-speed detection, ice altimetry, soil moisture monitoring, and more. The main focus of GNSS-R research to date has been on forward-scattered reflections, but theoretical work has proposed a backscattering regime and associated new application opportunities, including marine target detection. This article discusses the methods and results of processing the U.K. TechDemoSat-1 raw data collections in a backscattering regime for the first time, with initial results from sea ice datasets presented. The research has also identified a key problem with the backscatter method-for certain geometries the power from the specular point (forward scattered) may contaminate the data. The theory behind this and a method for predicting such occurrences is also discussed.
GNSS Reflectometry (GNSS-R), a method of remote sensing using the reflections from satellite navigation systems, was initially envisaged for ocean wind speed sensing. In recent times there has been significant interest in the use of GNSS-R for sensing land parameters such as soil moisture, which has been identified as an Essential Climate Variable (ECV). Monitoring objectives for ECVs set by the Global Climate Observing System (GCOS) organisation include a reduction in data gaps from spaceborne sources. GNSS-R can be implemented on small, relatively cheap platforms and can enable the launch of constellations, thus reducing such data gaps in these important datasets. However in order to realise operational land sensing with GNSS-R, adaptations are required to existing instrumentation. Spaceborne GNSS-R requires the reflection points to be predicted in advance, and for land sensing this means the effect of topography must be considered. This paper presents an algorithm for on-board prediction of reflection points over the land, allowing generation of DDMs on-board as well as compression and calibration. The algorithm is tested using real satellite data from TechDemoSat-1 in a software receiver with on-board constraints being considered. Three different resolutions of Digital Elevation Model are compared. The algorithm is shown to perform better against the operational requirements of sensing land parameters than existing methods and is ready to proceed to flight testing.
The UK’s Defence Science and Technology Laboratory (Dstl) is partnering with the US Naval Research Laboratory (NRL) on a joint mission to launch miniature sensors that will advance space weather measurement and modelling capabilities. The Coordinated Ionospheric Reconstruction Cubesat Experiment (CIRCE) comprises two 6U cube-satellites that will be launched into a near-polar low earth orbit (LEO), targeting 500 km altitude, in 2021. The UK contribution to CIRCE is the In situ and Remote Ionospheric Sensing (IRIS) suite, complementary to NRL sensors, and comprising three highly miniaturised payloads provided to Dstl by University College London (UCL), University of Bath, and University of Surrey/Surrey Satellite Technology Ltd (SSTL). One IRIS suite will be flown on each satellite, and incorporates an ion/neutral mass spectrometer, a tri-band global positioning system (GPS) receiver for ionospheric remote sensing, and a radiation environment monitor. From the US, NRL have provided two 1U Triple Tiny Ionospheric Photometers (Tri-TIPs) on each satellite (Nicholas et al., 2019), observing the ultraviolet 135.6 nm emission of atomic oxygen at night-time to characterize the two-dimensional distribution of electrons.
This paper details 3-years of cadmium telluride (CdTe) photovoltaic performance onboard the AlSat- 1N CubeSat in low earth orbit. These are the first CdTe solar cells to yield I-V measurements from space and help to strengthen the argument for further development of this technology for space application. The data has been collected over some 17,000 orbits by the CubeSat with the cells showing no signs of delamination, no deterioration in short circuit current or series resistance. The latter indicating that the aluminium-doped zinc oxide transparent front electrode performance remained stable over the duration. Effects of temperature on open circuit voltage (Voc) were observed with a calculated temperature coefficient for Voc of -0.19 %/⁰C. Light soaking effects were also observed to increase the Voc. The fill factor decreased over the duration of the mission with a major contribution being a decrease in shunt resistance of all 4 of the cells. The decrease in shunt resistance is speculated to result from gold diffusion from the rear contacts into the absorber and through to the front interface. This has likely resulted in the formation of a deep trap state within the CdTe and micro-shunts formed between the rear and front contact. Further development of this technology should therefore utilise more stable back contacting methodologies more commonly employed for terrestrial CdTe modules.
Space photovoltaics is dominated by multi-junction (III-V) technology. However, emerging applications will require solar arrays with; high specific power (kW/kg), flexibility in stowage and deployment and a significantly lower cost than the current III-V technology offers. This research demonstrates direct deposition of thin film CdTe onto the radiation-hard cover glass that is normally laminated to any solar cell deployed in space. Four CdTe samples, with 9 defined contact device areas of 0.25 cm2, were irradiated with protons of 0.5 MeV energy and varying fluences. At the lowest fluence, 1×1012 cm-2, the relative efficiency of the solar cells was 95%. Increasing the proton fluence to 1×1013 cm-2 and then 1×1014 cm-2 decreased the solar cell efficiency to 82% and 4% respectively. At the fluence of 1×1013 cm-2, carrier concentration was reduced by an order of magnitude. Solar Cell Capacitance Simulator (SCAPS) modelling obtained a good fit from a reduction in shallow acceptor concentration with no change in the deep trap defect concentration. The more highly irradiated devices resulted in a buried junction characteristic of the external quantum efficiency, indicating further deterioration of the acceptor doping. This is explained by compensation from interstitial H+ formed by the proton absorption. An anneal of the 1×1014 cm-2 fluence devices gave an efficiency increase from 4% to 73% of the pre-irradiated levels, indicating that the compensation was reversible. CdTe with its rapid recovery through annealing, demonstrates a radiation hardness to protons that is far superior to conventional multi-junction III-V solar cells.
The AlSat-Nano mission is a joint endeavour by the UK and Algeria to build and operate a 3U CubeSat. The project was designed to provide training to Algerian students, making use of UK engineering and experience. The CubeSat was designed and built by the Surrey Space Centre (SSC) of the University of Surrey and hosts three UK payloads with operations run by the Algerian Space Agency (ASAL). The educational and CubeSat development were funded by the UK Space Agency (UKSA), whilst the UK payloads were self-funded. Launch and operations are funded by ASAL. This paper illustrates the development of the programme, the engineering of the satellite and the development of collaborative operations between the SSC and ASAL.
Space researchers at the University of Surrey and Surrey Satellite Technology Limited (SSTL) have developed 'STRaND-1', a satellite containing a smartphone payload that will be launched into orbit around the Earth later this year. STRaND-1 (Surrey Training, Research and Nanosatellite Demonstrator) is being developed by the Surrey team to demonstrate the advanced capabilities of a satellite built quickly using advanced commercial off-the-shelf components. The satellite will be launched into orbit around the Earth in 2011. The phone will run on Android's powerful open-source operating system. A powerful computer, built at the Surrey Space Centre, will test the vital statistics of the phone once in space. The computer will check which components of the phone are working normally and will relay images and messages back to Earth via a radio system. Once all the tests are complete, the plan is to switch off the micro computer and the smartphone will be used to operate parts of the satellite. The smartphone avionics suite is only one of the many technological advances packed into this 4kg satellite. To precisely point and manoeuvre, the satellite also incorporates advanced guidance, navigation and control systems. © 2011 IEEE.
The upper atmosphere is a transition region between the neutron-dominated aviation environment and satellite environment where primary protons and ions dominate. We report high altitude balloon measurements and model results characterising this radiation environment for single event effects (SEE) in avionics. Our data, from the RaySure solid-state radiation monitor, reveal markedly different altitude profiles for low linear energy transfer (LET) and high LET energy depositions. We use models to show that the difference is caused by the influence of primary cosmic ray particles, which induce counts in RaySure via both direct and indirect ionization. Using the new Model of Atmospheric Ionizing Radiation Effects (MAIRE), we use particle fluxes and LET spectra to calculate single event upset (SEU) rates as a function of altitude from ground level to the edge of space at 100 km altitude. The results have implications for a variety of applications including high altitude space tourism flights, UAVs and missions to the Martian surface.
Future space telescopes with diameter over 20 m will require in-space assembly. High-precision formation flying has very high cost and may not be able to maintain stable alignment over long periods of time. We believe autonomous assembly is a key enabler for a lower cost approach to large space telescopes. To gain experience, and to provide risk reduction, we propose a demonstration mission to demonstrate all key aspects of autonomous assembly and reconfiguration of a space telescope based on multiple mirror elements. The mission will involve two 3U CubeSat-like nanosatellites (“MirrorSats”) each carrying an electrically actuated adaptive mirror, and each capable of autonomous un-docking and re-docking with a small central “9U” class nanosatellite core, which houses two fixed mirrors and a boom-deployed focal plane assembly. All three spacecraft will be launched as a single ~40kg microsatellite package.
Space telescopes are our ‘eyes in the sky’ that enable unprecedented astronomy missions and also permit Earth observation integral to science and national security. On account of the increased spatial resolution, spectral coverage, and signal-to-noise ratio, there is a constant clamour for larger aperture telescopes by the science and surveillance communities. This paper addresses a 25 m modular telescope operating in the visible wavelengths of the electromagnetic spectrum; such a telescope located at geostationary Earth orbit would permit 1 m spatial resolution of a location on Earth. Specifically, it discusses the requirements and architectural options for a robotic assembly system, called Robotic Agent for Space Telescope Assembly (RASTA). Aspects of a first-order design and initial laboratory test-bed developments are also presented.
The measured and projected growth of space debris makes it clear that technology for the removal of spacecraft at the end-of-life is an absolute necessity if we are to prevent the Kessler syndrome of catastrophic collisional cascading. Electro-dynamic tethers (EDTs) have been proposed as an effective means of deorbiting spacecraft – particularly from low Earth orbit (LEO). Such systems rely on the Lorentz force developed by a long conductive tether cutting through the Earth’s magnetic field due to the host spacecraft’s orbital motion. The electro-motive force generated drives a current through the tether, which is returned through the local space plasma by some form of active or passive plasma-contacting electrode. This removes (or adds) energy from the spacecraft’s motion, causing it to lose (or gain) altitude. As such, EDTs have the advantage of been self-powered, and propellantless, however, to be effective, the tethers typically have to be several km long, and be very thin to save mass. They are therefore flexible and derive their stability through the gravity gradient effect. This leads to such systems being most effective in low-Earth equatorial orbits, and unfortunately, much less effective in near polar orbits (e.g. Sun-synchronous orbit) or for orbits beyond LEO. To this end, we have developed a novel concept for an uncontrolled removal system based on electro dynamical principles. Instead of a long flexible tether (which have proven problematic to deploy), we propose the use of long (~150m-300m) rigid electro-dynamic booms in a “bar” or “cross” formation, actively powered, and coated with an electron emissive material. The main advantage of such a structure is that, for satellites in polar orbits, it leads to a larger Lorentz force. Also, the deployment is more reliable and the attitude control is greatly simplified (compared to the use of a flexible tether). To complete the circuit, electrons will be passively collected by a conductive deployable “sail”, which will also act as a drag sail at low altitudes. A ground demonstrator is under development based around a 6U CubeSat structure, which could form the basis for a later in-orbit demonstrator. This work is conducted as a part of the European Commission funded Horizon-2020 TeSeR (Technology for Self-Removal) project, which aims to demonstrate the feasibility of a scalable post mission removal system which should be able to be connected to different satellites via a standard interface.
This paper details the AM0 conversion efficiency of a metal-organic chemical vapor phase deposition thin-film cadmium telluride (CdTe) solar cell deposited onto a cerium-doped cover glass (100 μm). An AM0 best cell conversion efficiency of 12.4% (0.25-cm2 contact area) is reported. An AM0 mean efficiency of 12.1% over eight cells demonstrated good spatial uniformity. Excellent adhesion of the cell structure to the cover glass was observed with an adhesive strength of 38 MPa being measured before cohesive failure of the test adhesive. The device structure on cover glass was also subject to severe thermal shock cycling of +80 °C to -196 °C, showing no signs of delamination and no deterioration of the photovoltaic (PV) performance.
The InflateSail CubeSat, designed and built at the Surrey Space Centre (SSC) at the University of Surrey, UK, for the Von Karman Institute (VKI), Belgium, is one of the technology demonstrators for the QB50 programme. The 3.2 kilogram InflateSail is “3U” in size and is equipped with a 1 metre long inflatable boom and a 10 square metre deployable drag sail. InflateSail's primary goal is to demonstrate the effectiveness of using a drag sail in Low Earth Orbit (LEO) to dramatically increase the rate at which satellites lose altitude and re-enter the Earth's atmosphere. InflateSail was launched on Friday 23rd June 2017 into a 505km Sun-synchronous orbit. Shortly after the satellite was inserted into its orbit, the satellite booted up and automatically started its successful deployment sequence and quickly started its decent. The spacecraft exhibited varying dynamic modes, capturing in-situ attitude data throughout the mission lifetime. The InflateSail spacecraft re-entered 72 days after launch. This paper describes the spacecraft and payload, and analyses the effect of payload deployment on its orbital trajectory. The boom/drag-sail technology developed by SSC will next be used on the RemoveDebris mission, which will demonstrate the applicability of the system to microsat deorbiting.
Human spaceflight to/on/from the Moon will benefit from exploitation of various in-situ resources such as water volatile and mineral. Evidence for water ice in Permanently Shadowed Regions (PSRs) on the Moon is both direct and indirect, and derives from multiple past missions including Lunar Prospector, Chandrayaan-1 and LCROSS. Recent lunar CubeSats missions proposed through the Space Launch Systems (SLS) such as Lunar Flashlight, LunaH-Map and Lunar Ice-Cube, will help improve our understanding of the spatial distribution of water ice in those lunar cold traps. However, the spatial resolution of the observations from these SLS missions is on the order of one to many kilometres. In other words, they can miss smaller (sub-km) surficial deposits or near-surface deposits of water ice. Given that future lunar landers or rovers destined for PSRs will likely have limited mobility (but improved landing precision), there is a need to improve the spatial accuracy of maps of water ice in PSRs. The VMMO (Volatiles and Mineralogy Mapping Orbiter) is a semiautonomous, low-cost 12U lunar Cubesat being developed by a multi-national team funded through European Space Agency (ESA) for mapping lunar volatiles and mineralogy at relatively high spatial resolutions. It has a potential launch in 2023 as part of the ESA/SSTL lunar communications pathfinder orbiter mission. This paper presents the work carried out so far on VMMO concept design and development including objectives, profile, operations and spacecraft payload and bus.
Observations of highly red-shifted 21-cm hydrogen signals have been suggested as the only means to probe the early Universe from recombination to reionization. During this era, called the Dark Ages, the Universe consisted of neutral hydrogen gas and was opaque to light. It did not become transparent, as we see it today, until reionization was completed. The Dark Ages was the time period when matter clumped together, the very first stars and black holes were born, and, eventually, the first galaxies were formed. To enable observations of the Dark Ages is therefore one of the top priorities in cosmology and astrophysics. Today, the cosmological 21-cm signals are highly red-shifted and should peak in the FM radio band. Observing the Dark Ages from Earth is therefore next to impossible, due to man-made radio frequency interference (RFI) and ionospheric disturbances. To efficiently block the RFI, which would otherwise overwhelm the weak cosmological signal; it has been proposed to use the Moon as a radio shield and either place a satellite equipped with an ultra-sensitive radio instrument in lunar orbit or to deploy a large low-frequency radio array on the far-side of the Moon. Such missions are technically challenging and expensive and have so far failed to gain support from any national or international space program. Our goal is therefore to use a constellation of small inexpensive satellites in lunar orbit to collect pathfinder data, which would demonstrate EPSC Abstracts Vol. 9, EPSC2014-798, 2014 European Planetary Science Congress 2014 c Author(s) 2014 EPSC European Planetary Science Congress the feasibility of using the Moon as a radio shield, and map out the spatial extent of this RF quiescent zone to support future missions to explore the cosmos. This paper examines the design and radio payload of this mission. Alternative orbits, constellation and payload designs are analyzed to optimize the mission for performance and cost.
CubeSats are being increasingly specified for demanding Earth Observation and Astronomical applications where precise pointing, agility and stability are critical requirements. Such precision is difficult in the case of CubeSats as, firstly, their small moments of inertia mean that even small disturbance torques, such as those due to a residual magnetic moment, have a significant effect. Secondly, there are hardware limitations in terms of power, weight and size, which make the task more challenging. Recently, a research programme has been undertaken at Surrey Space Centre, to study the source of the residual magnetic moment in CubeSats, and to characterise the effect of the resulting disturbance on the attitude of the spacecraft. It has been found that, although the disturbances may be minimised by good engineering practice, in terms of minimising the use of permeable materials, and minimising current-loop areas, these disturbances can still be an issue when a high degree of stability is required. The dynamic nature of the disturbances requires an active mitigation strategy. We therefore propose a new technique using a network of magnetometers to dynamically characterize and then compensate the residual magnetic moment in real time. This paper reports on our findings to date.
A new Interplanetary electron environment model based on statistical analyses of historical datasets is presented. The model reports generates confidence limits for solar electron fluences in a similar fashion to existing Solar proton models, as well as peak event fluxes and fluences. Electrons of Jovian origin are also modeled based on simplified diffusive transport equations to provide predicted fluxes for locations within the ecliptic plane.
This paper presents the development and execution of an airborne experimental campaign as part of the continuing investigation of a passive Synthetic Aperture Radar using digital television broadcasting stations as illuminators of opportunity, and micro-/nano-satellite receivers in Low Earth Orbit. For the flight experiments, a hardware demonstrator was developed that utilised two receiving antennas, allowing both cross-correlation and auto-correlation range compression schemes, and was mounted to an airborne platform to image stationary rural areas up to 50 km from the transmitter. This paper presents the first image results of these experiments as well as initial analysis of image formation aspects including, range compression scheme and effects on the image quality of the signal to noise on the reference channel.
The Giove-A spacecraft carries two radiation monitors, CEDEX, built by the University of Surrey and Merlin, built by QinetiQ, to study the radiation environment encountered in the Galileo orbit. The two monitors have been functioning since the beginning of the mission and have measured protons, heavy ions and electrons. The electron environment has been found to be highly variable and driven by solar interactions. Comparisons with AE-8 indicate that the electron energy spectrum for the period measured was somewhat harder than that expected from the model. A series of large Solar proton events were detected in December 2006, registering as enhanced fluxes of protons, heavy ions and also triggering a large enhancement in the outer electron belt. Comparisons with POLE and INTEGRAL/IREM show an improved spectral match over AE-8.
CubeSats are being increasingly specified for demanding Earth Observation and Astronomical applications where precise pointing, agility and stability are critical requirements. Such precision is difficult in the case of CubeSats, mainly because their small moment of inertia means that even small disturbance torques, such as those due to a residual magnetic moment, have a significant effect. In addition, hardware limitations in terms of power, weight and size, make the task more challenging. The effect of magnetic disturbances has shown itself by the problem of high tumbling rates observed on several CubeSat missions. Post-flight analysis indicates this is often due to un-modelled magnetic moments mainly caused by the current flowing in the spacecraft. Some CubeSats also carry permanent magnets. However, by contrast, the other typical attitude disturbance sources for spacecraft (gravity gradient, aerodynamic, and solar radiation pressure torques) decreases significantly when the satellites become small. Recently, a research programme has been undertaken at Surrey Space Centre at the University of Surrey, to study the source of the residual magnetic field in CubeSats, and to characterise the effect of the resulting disturbance on the attitude of the spacecraft. It has been found that, although the disturbances may be minimised by good engineering practice, in terms of minimising current-loop areas, and minimising the use of permeable materials, these disturbances can still be an issue when a high degree of stability is required. The dynamic nature of the disturbances requires an active mitigation strategy. We therefore propose a new technique using a network of magnetometers to dynamically characterize and then compensate the calculated residual magnetic moment – in flight and in real time. This can be done by implementing a network of eight miniature 3-axis magnetometers on the spacecraft, with an additional one mounted on a deployable boom. These are used to determine the strength and the centre of the magnetic dipole of the spacecraft dynamically. The information will be used by the Attitude Determination and Control System (ADCS) control loops to compensate the measured residual magnetic moment. This technique will contribute to achieving more precise pointing, agility and stability of CubeSats. A hardware prototype using eight HMC1053 3-axis magnetometers monitored and controlled via a Raspberry Pi, was developed and successfully tested with the engineering model of the Alsat-1N CubeSat in a Helmholtz Coil arrangement at the Surrey Space Centre. This demonstrated the real-time dynamic measurement aspect of the technique proposed in this paper. This paper reports on our findings to date.
A steady increasing trend towards millimetre waves (mm-waves) for next generation communication has initiated an intensive research in the field of mm-wave antenna technologies. Reflectarray antennas being one of the potential candidates offer significant advantages over parabolic and phased array antennas at mm-wave bands. In a well-designed reflectarray, the overall performance is mainly determined by its comprising unit cell(s). Most of the recent reflectarray designs are based on printed microstrip technology. It is well known that surface waves get generated in printed microstrip technology and contribute to loss in the radiated signal power in the intended direction. This paper analyses the effect of surface waves in the reflection properties of a printed microstrip millimetre wave reflectarray unit cell. The analytical results are compared with measured data at 32 GHz and an excellent agreement was observed. It was observed that surface waves, though generally considered to have malign effects in antennas, play a significant positive role in the reduction of reflection loss magnitude at unit cell level.
Modern small satellites (MSS) are revolutionizing the space industry. They can drastically reduce the mission cost, and can make access to space more affordable. The relationship between a modern small satellite and a "conventional" large satellite is similar to that between a modern compact laptop and a "conventional" work-station computer. This paper gives an overview of antenna technologies for applications in modern small satellites. First, an introduction to modern small satellites and their structures is presented. This is followed by a description of technical challenges in the antenna designs for modern small satellites, and the interactions between the antenna and modern small satellites. Specific antennas developed for modern small-satellite applications are then explained and discussed. The future development and a conclusion are presented.
Space photovoltaics is dominated by multi‐junction (III‐V) technology. However, emerging applications will require solar arrays with high specific power (kW/kg), flexibility in stowage and deployment, and a significantly lower cost than the current III‐V technology offers. This research demonstrates direct deposition of thin film CdTe onto the radiation‐hard cover glass that is normally laminated to any solar cell deployed in space. Four CdTe samples, with 9 defined contact device areas of 0.25 cm2, were irradiated with protons of 0.5‐MeV energy and varying fluences. At the lowest fluence, 1 × 1012 cm−2, the relative efficiency of the solar cells was 95%. Increasing the proton fluence to 1 × 1013 cm−2 and then 1 × 1014 cm−2 decreased the solar cell efficiency to 82% and 4%, respectively. At the fluence of 1 × 1013 cm−2, carrier concentration was reduced by an order of magnitude. Solar Cell Capacitance Simulator (SCAPS) modelling obtained a good fit from a reduction in shallow acceptor concentration with no change in the deep trap defect concentration. The more highly irradiated devices resulted in a buried junction characteristic of the external quantum efficiency, indicating further deterioration of the acceptor doping. This is explained by compensation from interstitial H+ formed by the proton absorption. An anneal of the 1 × 1014 cm−2 fluence devices gave an efficiency increase from 4% to 73% of the pre‐irradiated levels, indicating that the compensation was reversible. CdTe with its rapid recovery through annealing demonstrates a radiation hardness to protons that is far superior to conventional multijunction III‐V solar cells.
Over the next two decades, unprecedented astronomy missions could be enabled by space telescopes larger than the James Webb Space Telescope. Commercially, large aperture space-based imaging systems will enable a new generation of Earth Observation missions for both science and surveillance programs. However, launching and operating such large telescopes in the extreme space environment poses practical challenges. One of the key design challenges is that very large mirrors (i.e. apertures larger than 3m) cannot be monolithically manufactured and, instead, a segmented design must be utilized to achieve primary mirror sizes of up to 100m. Even if such large primary mirrors could be made, it is impossible to stow them in the fairings of current and planned launch vehicles, e.g., SpaceX’s Starship reportedly has a 9m fairing diameter. Though deployment of a segmented telescope via a folded-wing design (as done with the James Webb Space Telescope) is one approach to overcoming this volumetric challenge, it is considered unfeasible for large apertures such as the 25m telescope considered in this study. Parallel studies conducted by NASA indicate that robotic on-orbit assembly (OOA) of these observatories offers the possibility, surprisingly, of reduced cost and risk for smaller telescopes rather than deploying them from single launch vehicles but this is not proven. Thus, OOA of large aperture astronomical and Earth Observation telescopes is of particular interest to various space agencies and commercial entities. In a new partnership with Surrey Satellite Technology Limited and Airbus Defence and Space, the Surrey Space Centre is developing the capability for autonomous robotic OOA of large aperture segmented telescopes. This paper presents the concept of operation and mission analysis for OOA of a 25m aperture telescope operating in the visible waveband of the electromagnetic spectrum; telescopes of this size will be of much value as it would permit 1m spatial resolution of a location on Earth from geostationary orbit. Further, the conceptual evaluation of robotically assembling 2m and 5m telescopes will be addressed; these missions are envisaged as essential technology demonstration precursors to the 25m imaging system.
Microstrip printed reflectarrays are becoming a potential replacement of parabolic reflector and phased array antennas due to their simple design, low cost and ease of manufacture to attain high gain and wide angle beam pointing at millimeter waves (mm-waves). Significant challenges are faced while implementing continuous phase reflectarrays at mm-waves. However, discretizing the required reflection phase provides a practically implementable solution. This contribution addresses the selection of phase states and its scattering in a phase discretized mm-wave reflectarray. The performance of two 1.5 bit phase quantized reflectarrays having closely spaced geometrical features is analyzed at 60 GHz. This study provides a better understanding to achieve a wider bandwidth response in practically implementable mm-wave reflectarrays.
This paper describes the radiation susceptibility testing and analysis of a miniaturised space Global Positioning System (GPS) receiver for small satellite applications. Tests on commercial-off-the-shelf (COTS) parts included total ionising dose (TID), single-event effect (SEE) testing and receiver operational effects under heavy ion exposure.
The increasing power demands of spacecraft payloads, and the future prospect of space based solar photovoltaic (PV) power stations, mean that there is an emerging requirement for large area solar arrays that will provide far greater power (kWpeak) than is currently available. To be practical, such arrays will need to use solar cells which have a much higher specific power (i.e. power per unit mass) and a much lower cost per watt than current space-rated solar PV technologies. To this end, the Centre for Solar Energy Research (CSER) at Swansea University, the Surrey Space Centre (SSC) and the Department for Mechanical Sciences at the University of Surrey, have been working on a new solar cell technology, based on thin film cadmium telluride (CdTe), deposited directly onto ultra-thin space qualified cover glass. This offers a potentially high specific power, low-cost technology with the added benefit of allowing a high degree of solar array flexibility for improved stowage volume and novel deployment strategies. Cells based on this innovative solar cell architecture have been manufactured and tested under a three year UK Engineering and Physical Science Research Council (EPSRC) funded project, with the result that highly efficient (for their class) cells were produced, which passed mechanical, thermal and ionising radiation tests with great success. Whilst this work was in progress, an opportunity to fly test cells on the joint Algerian Space Agency – UK Space Agency AlSat-1N Technology Demonstration CubeSat arose, and a successful bid was made to fly a payload capable of characterising the cells in orbit, via an automatic Current-Voltage (I-V) measurement circuit. The resulting Thin Film Solar Cell (TFSC) payload, comprising four test cells, was integrated onto AlSat-1N at Surrey, and launched from India into a 661 km × 700 km, 98.20° Sun Synchronous orbit on 26th September 2016. This paper describes the new cell technology, the pre-flight ground testing, the flight payload, and the first flight results of thin film CdTe solar cells flying on an international 3U CubeSat technology demonstrator. Keywords: (Thin-Film Solar Cells, CubeSat, Technology Demonstration)
Three spacecraft for the UK, Turkey and Nigeria were launched together in September 2003, to join Algeria's satellite, AlSat-1, in the Disaster Monitoring Constellation (DMC). Surrey Satellite Technology Ltd. has designed, built and launched the world's first constellation to provide daily global Earth observation coverage at moderate resolution in three spectral bands. This international initiative will provide daily images for global disaster monitoring, as well as supporting each partner nation's indigenous remote sensing requirements. The DMC programme establishes a novel model for international collaboration, and demonstrates how small satellite missions can be employed for a wide range of applications. This paper shows the first in-orbit mission results from DMC satellites including examples of unique EO data products comprising up to 600 x 600 km images gathered at 32-metres GSD in 3 spectral bands.
A novel dual-polarized broadband antenna array for S-band is presented. This antenna is composed of 6 × 2 microstrip antenna elements with a hybrid feed-line network providing an isolation ≥ 18.6 dB between the H- and V-ports. The operative bandwidth is from 3.15 to 3.25 GHz, and the peak measured gain is approximately 19 dBi. The array is suitable for spacecraft operation because of the selected materials, its low profile (~8 mm thickness), and light weight. It has potential applications in synthetic aperture radar (SAR), remote sensing, and wireless communications.
Reflectarray antennas are a potential candidate solution to realize high gains at millimetre waves (mm-waves). A reflectarray contains a large number of spatially illuminated unit cells. The performance of a good reflectarray design is manifested by the behaviour of its comprising unit cells. An established technique to characterise a unit cell is by placing it inside a waveguide to achieve periodic boundary conditions. This usually requires custom waveguide products; making the tests difficult and expensive. Additionally, when the unit cells are reconfigurable as in a smart reflectarray it is hard to take the DC bias lines out of the waveguide without using custom made waveguide parts. This contribution address the issue of unit cell placement inside the waveguide and proposes simple unit cell structures to avoid custom made waveguide parts. The idea was verified by measuring a series of unit cells at mm-waves in various configurations and a practically acceptable agreement was found. The proposed structures greatly simplify the reconfigurable unit cell testing.
The size of any single spacecraft is ultimately limited by the volume and mass constraints of currently available launchers, even if elaborate deployment techniques are employed. Costs of a single large spacecraft may also be unfeasible for some applications such as space telescopes, due to the increasing cost and complexity of very large monolithic components such as polished mirrors. The capability to assemble in-orbit will be required to address missions with large infrastructures or large instruments/apertures for the purposes of increased resolution or sensitivity. This can be achieved by launching multiple smaller spacecraft elements with innovative technologies to assemble (or self-assemble) once in space and build a larger much fractionated spacecraft than the individual modules launched. Up until now, in-orbit assembly has been restricted to the domain of very large and expensive missions such as space stations. However, we are now entering into a new and exciting era of space exploitation, where new mission applications/markets are on the horizon which will require the ability to assemble large spacecraft in orbit. These missions will need to be commercially viable and use both innovative technologies and small/micro satellite approaches, in order to be commercially successful, whilst still being safety compliant. This will enable organisations such as SSTL, to compete in an area previously exclusive to large commercial players. However, inorbit assembly brings its own challenges in terms of guidance, navigation and control, robotics, sensors, docking mechanisms, system control, data handling, optical alignment and stability, lighting, as well as many other elements including non-technical issues such as regulatory and safety constraints. Nevertheless, small satellites can also be used to demonstrate and de-risk these technologies. In line with these future mission trends and challenges, and to prepare for future commercial mission demands, SSTL has recently been making strides towards developing its overall capability in “in-orbit assembly in space” using small satellites and low-cost commercial approaches. This includes studies and collaborations with Surrey Space Centre (SSC) to investigate the three main potential approaches for in-orbit assembly, i.e. deployable structures, robotic assembly and modular rendezvous and docking. Furthermore, SSTL is currently developing an innovative small ~20kg nanosatellite (the “Target”) as part of the ELSA-d mission which will include various rendezvous and docking demonstrations. This paper provides an overview and latest results/status of all these exciting recent in-orbit assembly related activities.
The InflateSail (QB50-UK06) CubeSat, designed and built at the Surrey Space Centre (SSC) at the University of Surrey, UK, for the Von Karman Institute (VKI), Belgium – was one of the technology demonstrators for the QB50 pro-gramme. The 3.2 kilogram 3U CubeSat was equipped with a 1 metre long inflat-able boom and a 10m2 deployable drag sail. InflateSail's primary mission was to demonstrate the effectiveness of using a drag sail in Low Earth Orbit (LEO) to dramatically increase the rate at which satellites lose altitude and re-enter the Earth's atmosphere and it was one of 31 satellites that were launched simultane-ously on the PSLV (polar satellite launch vehicle) C-38 from Sriharikota, India on 23rd June 2017 into a 505km, 97.44o Sun-synchronous orbit (SSO). Shortly after orbital insertion, InflateSail booted-up, and, once safely clear of the other satellites on the launch, it automatically activated its payload – firstly, deploying a 1 metre long inflatable boom comprising a metal-polymer laminate tube, using a cool gas generator (CGG) to provide the inflation gas, and secondly, using a brushless DC motor at the end of the boom to extend four lightweight bistable rigid composite (BRC) booms to draw out the 3.1m x 3.1m square, 12 micron thick polymer drag-sail. As intended, the satellite immediately began to lose alti-tude, and re-entered the atmosphere just 72 days later – thus demonstrating for the first time the de-orbiting of a spacecraft using European inflatable and drag-sail technologies. The boom/drag-sail technology developed by SSC will next be used on the RemoveDebris mission, due for launch in 2018, which will demon-strate the capturing and de-orbiting of artificial space debris targets using a net and harpoon system.
Surrey Space Centre has been working on an autonomous fixed-wing all-electric vertical take-off and landing (VTOL) aerobot for the exploration of Mars for several years. The current design is a novel “Y-4” configured tilt-rotor, comprising two large fixed co-axial lift rotors embedded in a blended wing/body, with a pair of smaller tractor tilt-rotors mounted just forward of the wing. Thus, there are 4 rotors configured in a “Y” shape. During take-off and landing, all four rotors operate in the vertical direction, with the bulk of the lifting force being provided by the thrust of the co-axial lift rotors. During transition to horizontal flight, the pair of tilt-rotors rotate to the horizontal position and the co-axial rotors are slowed as the wings begin to provide aerodynamic lift. Once sufficient speed has been built up, these rotors are stopped, and a set of clam-shell doors close to enclose them to provide a smooth lifting surface over the body. Thus, in forward flight, only the pair of tractor tilt-rotors operate, thereby considerably reducing the electrical power demands of the aircraft compared to, for example, a conventional quad-copter or helicopter design. The baseline mission of the aerobot is to investigate the Isidis Planitia region on Mars over a month long period using optical sensors during flight and a surface science package when landed. During flight operations the aerobot will take off and land vertically, transitioning to and from horizontal flight. The flight time is around an hour, with the flights taking place close to local noon to maximize the power production of the wing/body mounted solar cells. A nonlinear six degree of freedom (6DoF) dynamic model incorporating aerodynamic models of the aerobot’s body and rotors has been developed to model the vertical, transition, and horizontal phases of flight. A nonlinear State-Dependent Riccati Equation (SDRE) controller has been developed for each of these flight phases. The nonlinear dynamic model was transformed into a pseudo-linear form based on the states and implemented in the SDRE controller. During transition flight the aerobot is over actuated and the weighted least squares (WLS) method is used for allocation of control effectors. Simulations of the aerobot flying in different configurations were performed to verify the performance of the SDRE controllers, including hover, transition, horizontal flight, altitude changes, and landing scenarios. Results from the simulations show the SDRE controller is a viable option for controlling this novel Martian Aerobot.
In recent years, there has been a desire to develop space-based optical telescopes with large primary apertures. Current monolithic large telescopes, as exemplified by 6.5m aperture James Webb Space Telescope, are limited by the diameter of the launch vehicle – despite their ability to unfold and deploy mirror elements. One method to overcome this obstacle is to autonomously assemble small independent spacecraft, each with their own mirror, while in orbit. In doing so, a telescope with a large, segmented primary mirror can be constructed. Furthermore, if each of these mirrors is manufactured to have an identical initial shape and then adjusted upon assembly, a substantial reduction in manufacturing costs can be realized. In order to prove the feasibility of such a concept, a collaborative effort between the California Institute of Technology, the University of Surrey, and the Indian Institute of Space Science and Technology has been formed to produce and fly the "Autonomous Assembly of a Reconfigurable Space Telescope" (AAReST) mission. AAReST comprises two 3U Cubesat-like nanosatellites (“MirrorSats”) each carrying an electrically actuated adaptive mirror, and each capable of autonomous un-docking and re-docking with a central “9U” class nanosatellite (“CoreSat”), which houses two fixed mirrors and a boom-deployed focal plane assembly (camera). All three spacecraft will be launched as a single ~30kg microsatellite package. The central premise is that the satellite components can manoeuvre and dock in different configurations and the mirrors can change shape and move to form focused images on the camera focal plane. The autonomous manoeuvres and docking will be under the control of the Surrey developed electro-magnetic docking system and near infra-red lidar/machine-vision based relative navigation sensors. On orbit, the mission profile will firstly establish the imaging capability of the compound spacecraft before undocking, and then autonomously re-docking a single MirrorSat. This will test the docking system, autonomous navigation and system identification technology. If successful, the next stage will see the second MirrorSat spacecraft undock and re-dock to the core spacecraft to form a wide linear formation which represents a large (but sparse) aperture for high resolution imaging. Celestial targets will be imaged. Currently, the flight hardware is under construction and launch is planned for ~2019-2020. This paper details the mission concept, technology involved and its testing and progress on the production of the flight hardware.
Presented here is CHAFF (CubeSat Hyperspectral Application For Farming), a design concept for a CubeSat-based Hyperspectral Imager (CHSI) intended to supply high quality hyperspectral image data cubes to the agricultural community. CHAFF has been designed holistically as a system, considering all design and operational characteristics of a CHSI instrument and platform together: including the re-stricted payload mass and volume associated with CubeSats, the platform pointing stability/accuracy limitations, and the restricted downlink data budget. To this end, CHAFF will employ optically aided geometric co-registration methods, which will allow on-board construction of the hyperspectral data cube. This allows the use of powerful lossless data compression schemes to mitigate the downlink data budget limitations. In addition, a calibration methodology using a tuneable laser source at NPL, will be employed pre-flight to achieve rapid and accurate spectral and radiometric calibration, essential for the production of science-grade data sets from the proposed CubeSat constellations. A benchtop prototype has been constructed and a promising spectral resolution of 3nm at around 625nm has been achieved. In addition the auxiliary imager for the optically-aided geometric co-registration has been demonstrated.
CubeSats to date have shown an excellent potential in providing quality science and demonstrating the use of COTS technology in space applications. A next stage in the evolution of CubeSat technology is the ability to demonstrate useable on board propulsion. A propulsion flight module has been designed and developed that will be able to provide pitch, roll and yaw around a central axis and translational movement in two axes. The module comprises of eight parallel bar micro Pulsed Plasma Thrusters (μPPTs). To aid in the miniaturisation there have been two significant changes. The first was a replacement of the typical sparkplug with a contact trigger mechanism that initiates a discharge. The second was the removal of the typical TeflonTM propellant bar used in PPTs between the discharging electrodes. In studies at the Surrey Space Centre it was shown that discharges without the presence of TeflonTM produced 60-75% the impulsebit (based on integral calculations of the averaged current discharge profiles) compared to discharges which had the presence of TeflonTM between the electrodes, at the parameters that were tested. The mass eroded for the plasma production was theorised to originate from the electrodes, which is similar in the mechanisms of operation to the Vacuum Arc Thruster (VAT). The module is split into three PC104 boards, two boards house four μPPTs and the third board is the power unit. The power unit uses award winning minature voltage multipliers that take the 5V CubeSat bus voltage and transforms this to 800V for the PPT high voltage capacitors. This paper focuses on the developmental work that has been conducted to construct a propulsion module for the Surrey Training Research and Nano-satellite Demonstration (STRaND) 3U CubeSat.
The radiation monitors on board the Galileo Giove-A satellite, CEDEX and Merlin, and their data are presented. The instruments include energetic proton and ion detectors, an internal charging monitor, RADFETs and experimental dose-rate photodiodes. A comparison of the data with existing monitors and models is presented.