Prof. Craig I. Underwood graduated from the University of York in 1982 with a B.Sc. in Physics with Computer Science. After gaining a Post Graduate 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 full 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 & 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 (Supercapcitors, Thin-Film Solar PV, RTGs, solar-thermal power)PhD Topics are available in any of these or related areas.
The Group's personnel and their research interests can be seen here.
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 & EffectsOver 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 InstrumentationOptical: 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.
Radar: 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 post-graduate, undergraduate and industrial-training courses. He was the recipient of the Department of Electrical and Electronic Engineering's 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
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.
Switched-Beam Array for Global Navigation
Satellite System, IEEE Transactions on Antennas and Propagation 62 (4) pp. 1-8
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.
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.
cerium?doped space glass, Progress in Photovoltaics 25 (12) pp. 1059-1067 Wiley
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.
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
This research provides a potentially competing novel high gain electronic beamsteering antenna solution for mm-waves in the form of a phase quantized smart reflectarray consisting of high performance reconfigurable unit cells. Novel contributions of this research are: (a) Analysis of mm-wave reflectarray unit cells including the effects of fringing fields, surface waves, finite metal conductivity and metal surface roughness. (b) New measurement techniques for mm-wave reflectarray unit cells to ease the alignment, orientation, and DC biasing issues. (c) Characterization of PIN diodes at 10 GHz and 60 GHz for their ON/OFF state models extraction from measurements. (d) Design of three state implicit phase shifter reflectarray unit cell at 60 GHz, reduction in its DC bias lines, and an optimization technique to improve polarization purity of a multi-state reconfigurable unit cell. (e) A fast algorithm to prepare the electromagnetic simulation model of large reflectarrays. (f) Conception and measurement based validation of phase quantized reflecarrays and their performance matrix. (g) Conception and measurement based analytical solution of low DC power consuming smart reflectarrays.
The resulting solution is agile, simple to implement, do not necessarily require multiple RF chains, enables wide angle electronic beamsteering (+-78 degree), is scalable for any gain/frequency requirements, can be made foldable for smaller satellite platforms, is very reliable, and consumes low DC power. This smart reflectarray platform can implement any phase only synthesis technique for radiation pattern control including single/multiple pencil beams, contoured beams, and their scanning over wider angles. Findings of this research would potentially benefit next generation terrestrial/air/space communication systems and radars.
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.
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.
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.
Keywords: (Thin-Film Solar Cells, CubeSat, Technology Demonstration)
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.
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
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
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, 12mm 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.
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.
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.
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
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
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.
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
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.
This work describes a full end-to-end analysis of the uncertainty at a pixel level of the Top-Of-Atmosphere (TOA) radiance/reflectance factor products. It develops a methodology framework that can be adapted and reproduced by several EO missions to provide TOA radiometric uncertainty. The method is not only described but implemented as a software tool named Radiometric Uncertainty Tool (RUT) using as an example the Sentinel-2 (S2) mission.
The uncertainty methodology starts from a radiometric model, where a set of uncertainty contributors are identified and specified at a pixel level, by reviewing the pre- and post-launch sensor radiometric characterisations. These contributors are assessed using the metadata and quality information associated to the satellite products where possible. As a consequence, the uncertainty contributions are specified for the specific satellite acquisition time, scene and processing. Some of the uncertainty contributions required the use of novel estimation methods that have been specifically applied to the assessment of the uncertainty propagation produced by the image orthorectification and the radiometric impact of the spectral knowledge. The study proposes an uncertainty combination model with an important effort in using the best metrological practices as described in the ?Guide to Expression of Uncertainty in Measurement? (GUM) model. The assumptions in the model have been validated by comparing the results to a Monte Carlo Method (MCM), the correlation among the different uncertainty contributions has been studied, and the impact of simplifications in the combination model has been assessed. As an extension of the work towards its larger application, a methodology has been proposed and implemented to estimate the uncertainty associated to the mean of the pixels in a Region of Interest (ROI). The study considers the correlation of the pixels in the spatial, temporal and spectral dimension. As a result, the TOA radiometric uncertainty estimates can be of direct use for applications as the radiometric validation activities or product spatial binning. Further extension of the uncertainty concepts has resulted in a set of tools, algorithms and methodologies that have been 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 (Traceable Radiometry Underpinning Terrestrial and Helio-Studies) mission. This study considers the criticality of the instrumental and observational characteristics on pixel level reflectance factors, within a defined spatial ROI within the target site. It quantifies the main uncertainty contributors in the spectral, spatial, and temporal dimension.