Surrey Space Centre Research Articles

IAC-22-C3.3.8 Six years of spaceflight results from the AlSat-1N Thin-Film Solar Cell (TFSC) experiment

The increasing power demands of spacecraft payloads and the realistic prospect of space based solar power (SBSP) stations as a means of providing zero carbon electricity in the 2030s, means that there is an emerging requirement for large area, yet lightweight, solar photovoltaic (PV) 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 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 material. This offers a potentially high specific power and when adopting the conventional CdTe manufacturing process, a low-cost technology. The ultra-thin glass can produce a solar cell which is sufficiently flexible to allow “roll-out” deployment strategies. Four prototype cells were flown as part of the Thin-Film Solar Cell (TFSC) experimental payload, developed by CSER and the Surrey Space Centre (SSC), on the joint Algerian Space Agency (ASAL) – UK Space Agency AlSAT-1N Technology Demonstration CubeSat, launched into a 661 km × 700 km, 98.20° Sun Synchronous orbit, on September 26, 2016. The experiment has provided the first in-orbit current/voltage (I/V) measurements of this novel technology, and more than five years of flight results have now yielded new insights into its longer-term performance and inherent radiation hardness, which makes them particularly attractive for maintaining high end-of-life (EOL) performance for long duration space missions. The results help to strengthen the argument for further development of this technology for space application. The data, collected over ∼30,000 orbits, show no signs of cell delamination (a potential risk for such technologies), no deterioration in short circuit current or in series resistance. However, all four cell's fill factors were observed to decrease over the duration of the mission, caused primarily by a decrease in their shunt resistance. This has been attributed to the diffusion of gold atoms from the back electrical contacts. We conclude therefore that further development of this technology should utilize more stable back contacting methodologies more commonly employed for terrestrial CdTe modules. However, this flight has proven the basic soundness of the technology for use in space.

Validation of a torsional balance for thrust measurements of Hall effect and microwave-based space propulsion systems.

A torsional thrust balance has been designed and validated by Surrey Space Centre and Added Value Solutions UK Ltd. in collaboration with the UK Space Agency. The thrust stand has been tested with two electric propulsion (EP) systems operating with xenon: the Halo thruster and the XJET thruster. The first consists of a low-power (<1 kW) Hall effect-based thruster, whose thrust level is between 3 and 20mN, depending on the power of the system. The second is an electron cyclotron resonance thruster whose operative point is in the 0.3-1.5mN thrust range. The thruster is mounted on a titanium rotating beam, whose movement is measured by an optical fiber displacement sensor. The thrusters' direct current electrical connections are routed through room temperature liquid metal pots and microwave power is transmitted via a wireless transfer system, minimizing friction effects. To reduce thermal issues during long thruster operations, the torsional thrust balance is designed with a water-cooling hub around the flex pivot. Noise from the laboratory environment is lessened by using four vibration-dampening spring systems as thrust balance feet. The tests on the two EP systems have shown accurate and repeatable results, demonstrating that the balance can be used to characterize different EP systems in the μN-mN thrust range.

RemoveDEBRIS: An in-orbit demonstration of technologies for the removal of space debris

ABSTRACTThe RemoveDEBRIS mission has been the first mission to successfully demonstrate, in-orbit, a series of technologies that can be used for the active removal of space debris. The mission started late in 2014 and was sponsored by a grant from the EC that saw a consortium led by the Surrey Space Centre to develop the mission, from concept to in-orbit demonstrations, that terminated in March 2019. Technologies for the capture of large space debris, like a net and a harpoon, have been successfully tested together with hardware and software to retrieve data on non-cooperative target debris kinematics from observations carried out with on board cameras. The final demonstration consisted of the deployment of a drag-sail to increase the drag of the satellite to accelerate its demise.

InflateSail de-orbit flight demonstration results and follow-on drag-sail applications

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 m long inflatable mast and a 10 m2 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 505 km, 97.44° Sun-synchronous orbit.Shortly after safe deployment in orbit, InflateSail automatically activated its payload. Firstly, it inflated its metre-long 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 3.1 m × 3.1 m square, 12 μm 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 Centre. 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 mast/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.

Modular Testbed for Spinning Spacecraft

The ground verification of a spacecraft control algorithm is commonly done via air bearing facility. Air bearing testbeds are frequently developed for testing a three-axis stabilized spacecraft control algorithm but hardly for a spin-stabilized spacecraft. A modular testbed for testing a spinning spacecraft has been developed at Surrey Space Centre initially for the real-time verification of a prolate spinner slew control algorithm. This testbed is made from commercial off-the-shelf components with a modular system design approach through rapid control prototyping using Matlab xPC Target and is extendable to other rapid control prototyping techniques. It is equipped with a novel low-cost monocular vision system for attitude determination with accuracy of 0.06 deg and angular velocity accuracy of . For the current specification, a cold-gas propulsion system is fixed to the testbed with a two-degree-of-freedom thruster set that can deliver up to 0.25 N of thrust and an air bearing capability that gives three degrees of freedom with a maximum tilt angle of 30 deg. In this paper, the testbed implementation is described, and the test platform is verified.

Using CubeSat/micro-satellite technology to demonstrate the Autonomous Assembly of a Reconfigurable Space Telescope (AAReST)

Future space telescopes with diameter over 20m 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 ~40kg 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.

Peircing the Extraterrestrial Surface: Integrated Robotic Drill for Planetary Exploration

As we explore our solar system and other extraterrestrial bodies, the subsurface plays a vital role in allowing us to peer back into the history of a particular body, looking for life or signs that it may have been habitable. This can be achieved by using a form of drill or penetrator, although traditional technologies require large masses to produce an overhead force (OHF) that pushes the drill into the subsurface. Dual reciprocating drilling (DRD) is a new biologically inspired technology based on the wood wasp ovipositor. It consists of two reciprocating backward-facing teethed halves that generate a drilling force that reduces the required overhead penetration force and mass requirements. The Surrey Space Centre (SSC) has overseen the design, development, and testing of a proof-of-concept model with funding from European Space Agency. The system is now evolving to include a drive mechanism within the drill head and bays for scientific instrumentation.

Editoral: Special Issue on Photonic and RF Communications Systems

Editoral: Special Issue on Photonic and RF Communications Systems

Analysis of state-of-the-art single-thruster attitude control techniques for spinning penetrator

The attitude dynamics and manoeuvre survey in this paper is performed for a mission scenario involving a penetrator-type spacecraft: an axisymmetric prolate spacecraft spinning around its minor axis of inertia performing a 90° spin axis reorientation manoeuvre. In contrast to most existing spacecraft only one attitude control thruster is available, providing a control torque perpendicular to the spin axis. Having only one attitude thruster on a spinning spacecraft could be preferred for spacecraft simplicity (lower mass, lower power consumption etc.), or it could be imposed in the context of redundancy/contingency operations. This constraint does yield restrictions on the thruster timings, depending on the ratio of minor to major moments of inertia among other parameters. The Japanese Lunar-A penetrator spacecraft proposal is a good example of such a single-thruster spin-stabilised prolate spacecraft.The attitude dynamics of a spinning rigid body are first investigated analytically, then expanded for the specific case of a prolate and axisymmetric rigid body and finally a cursory exploration of non-rigid body dynamics is made. Next two well-known techniques for manoeuvring a spin-stabilised spacecraft, the Half-cone/Multiple Half-cone and the Rhumb line slew, are compared with two new techniques, the Sector-Arc Slew developed by Astrium Satellites and the Dual-cone developed at Surrey Space Centre. Each technique is introduced and characterised by means of simulation results and illustrations based on the penetrator mission scenario and a brief robustness analysis is performed against errors in moments of inertia and spin rate. Also, the relative benefits of each slew algorithm are discussed in terms of slew accuracy, energy (propellant) efficiency and time efficiency. For example, a sequence of half-cone manoeuvres (a Multi-half-cone manoeuvre) tends to be more energy-efficient than one half-cone for the same final slew angle, but more time-consuming. As another example, the new Sector-Arc Slew and Dual-cone techniques are designed to overcome a specific restriction on attainable slew angle that is associated with the half-cone manoeuvre, giving one additional degree of freedom for designers to fine-tune.

Non-stop rejection

An ultra-wideband(UWB) antenna for short-range communications over 3.1–10.6 GHz that can reject interference from WiMAX and WLAN without needing extra bandstop filters has been presented by a team at the Surrey Space Centre at the University of Surrey in the UK.

CubeSail: A low cost CubeSat based solar sail demonstration mission

CubeSail is a nano-solar sail mission based on the 3U CubeSat standard, which is currently being designed and built at the Surrey Space Centre, University of Surrey. CubeSail will have a total mass of around 3 kg and will deploy a 5 × 5 m sail in low Earth orbit. The primary aim of the mission is to demonstrate the concept of solar sailing and end-of-life de-orbiting using the sail membrane as a drag-sail. The spacecraft will have a compact 3-axis stabilised attitude control system, which uses three magnetic torquers aligned with the spacecraft principle axis as well as a novel two-dimensional translation stage separating the spacecraft bus from the sail. CubeSail’s deployment mechanism consists of four novel booms and four-quadrant sail membranes. The proposed booms are made from tape-spring blades and will deploy the sail membrane from a 2U CubeSat standard structure. This paper presents a systems level overview of the CubeSat mission, focusing on the mission orbit and de-orbiting, in addition to the deployment, attitude control and the satellite bus.

Completely stripped solar sail concept using bi-stable reeled composite booms

The Surrey Space Centre is currently working on CubeSail, a 5×5 m 2 nano-solar sail spacecraft to be fitted inside a 3U Cubesat platform. The CubeSail mission has a two-fold aim: first, to demonstrate the concept of solar sailing using a 3-axis stabilized sail and second, to demonstrate the potential use of its large membrane as a drag augmentation device for de-orbiting satellites and launch vehicles' upper stages in low Earth orbit (LEO). The novel solar sail concept presented here is a result of the highly restrictive area-to-mass ratio and volume constraints imposed by the mission requirements. Therefore, this concept can be readily scaled up to produce future very high performance sailcrafts. Research on ultra-lightweight deployable structures leads to the invention of a new bi-stable composite boom that entails a light and simple deployment subsystem. This simplicity is the key feature that enables the sail to be attached along the length of the booms, thus reducing the structural mass requirement of these support structures. In the completely stripped sail concept the previously considered square or quadrant membrane is split into a set of parallel strips, in order to achieve a uniaxial tension state in the film. Also, a test-bed created for the successful deployment under gravity conditions of the 5 m 2 prototype sail is presented.

Lunar science with affordable small spacecraft technologies: MoonLITE and Moonraker

Returning to the Moon has been advocated by a large number of international planetary scientists in order to answer several key scientific questions. The UK also has an active lunar science community keen to support (robotic) lunar exploration missions. However, for several years these interests have been eclipsed by the drive to Mars. Recently there is a renewed global interest in the Moon demonstrated by the Vision for Space Exploration in the USA, the evolving Global Exploration Partnership, and new lunar missions from Europe, Japan, China, India and the USA. The ESA Aurora programme may also broaden its focus to embrace the Moon as well as Mars—realizing that the risks associated with many of the major technical challenges that are faced by Mars missions could be reduced by relatively inexpensive and timely lunar technology tests. Surrey Satellite Technology Ltd. (SSTL) and Surrey Space Centre (SSC) have been preparing a ‘smallsat’ approach [ Sweeting, M.N., Underwood, C.I., 2003. Small-satellite engineering and applications. In: Fortescue, P., Stark, J., Swinerd, G., (Eds.), Spacecraft Systems Engineering, third edition. Wiley, New York, pp. 581–612] to achieving a low-cost lunar mission for more than a decade—including various activities, such as the earlier LunarSat study funded by ESA and a current hardware contribution to the Chandrayaan-1 mission. With the recent successes in GIOVE-A, TOPSAT and BEIJING-1, 1 1 Recent Earth satellite missions built and launched by SSTL. alongside participation in Aurora and Chandrayaan-1, Surrey have developed capabilities for providing affordable engineering solutions to space exploration. Recently, SSTL/SSC was funded by the UK Particle Physics and Astronomy Research Council (PPARC) (now subsumed into the UK Science and Technology Facilities Council) to undertake a study on low-cost lunar mission concepts that could address key scientific questions. This paper presents some major results from this study [Phipps and Gao, 2006. Lunar mission options study. UK Particle Physics and Astronomy Research Council Report Reference No. 118537, pp. 1–104] and provides preliminary definitions of two mission proposals.

“You can get there from here”: Advanced low cost propulsion concepts for small satellites beyond LEO

Small satellites have historically been forced to use low cost propulsion, or to do without in order to maintain low cost. Since 1999 an increasing number of SSTL's customers have demanded the capability to precisely position and subsequently manoeuvre their satellites, driven largely by the current attraction of small satellite constellations such as Disaster Monitoring (DMC), which require propulsion for launcher injection error correction, drag compensation, constellation phasing and proximity manoeuvring and rendezvous. SSTL has successfully flight qualified a simple, low cost propulsion system based on a low power (15–100 W) resistojet employing green propellants such as butane and xenon, and demonstrated key constellation manoeuvres. The system is capable of up to 60 m/s deltaV and will be described here. The SSTL low power resistojet is however limited by a low Isp ( ∼ 50 s for Xenon in the present design, and ∼ 100 s with nitrogen and butane) and a slow reaction time ( 10 min warm-up required). An increasing desire to apply small satellite technology to high deltaV missions while retaining the low cost aspect demands new solutions. ‘Industry standard’ solutions based on cryogenic propulsion, or toxic, carcinogenic storable propellants such as hydrazine/nitrogen oxides combination are not favourable for small satellite missions developed within SSTL's low cost engineering environment. This paper describes a number of strawman missions with high deltaV and/or precision manoeuvring requirements and some low cost propulsion solutions which have been explored at the Surrey Space Centre to meet future needs: (1) Deployment of a complex constellation of nano- or pico-satellites from a secondary launch to a new orbit. The S3TV concept has been developed to allow deployment up to 12 payloads from an ‘off-the-shelf’ thrust tube, using a restartable nitrous oxide hybrid engine, operating in a dual mode with resistojets for attitude control. (2) Orbit transfer of an enhanced microsatellite from a typical 700 km sun-synchronous orbit to a lower or higher orbit using a low cost 40 N thrust concentrated hydrogen peroxide/kerosene bipropellant engine. A spin stabilized ‘tug’ concept capable of providing between 130 and 300 m/s of deltaV to the payload is described. (3) Transfer of an enhanced microsatellite from LEO to lunar orbit using a novel, storable propellant solar thermal propulsion system under development at the Surrey Space Centre. The solar thermal propulsion unit is designed for low cost small satellite support and will be compared with a more traditional approach using and industry standard storable bipropellant chemical engine. (4) Nanosatellite manoeuvring for formation flying using advanced low power electric propulsion. A colloid thruster system concept is planned for development jointly between SSTL, Queen Mary University London and Rutherford Appleton Laboratory, UK. The colloid thruster system is designed to complement an existing butane resistojet to give full 3-axis manoeuvrability to an upgraded SNAP nanosatellite platform which could be reflown in 2007 alongside ESA's Proba 2 technology demonstrator microsatellite. A comparison between low power resistojets, a colloid thruster system, and pulsed plasma thrusters for orbit manoeuvring of microsatellites will be made. This paper's final section will briefly describe some of the interplanetary missions which have been considered at the Surrey Space Centre, and will highlight the few as yet practical solutions for sending small spacecraft on high deltaV missions without the use of a costly upper stage.

CERISE microsatellite recovery from first detected collision in low Earth orbit

The first ever collision between two catalogued space objects was detected in low Earth orbit in July 1996. After one year of perfect operation in space, controllers at the Space Mission Control Groundstation, Surrey Space Centre, UK observed a sudden change to the attitude dynamics of the CERISE microsatellite as it tumbled rapidly end-over-end in its 700 km polar Earth orbit. After initial investigations, a collision with a piece of space debris was suspected. The paper describes the analysis of the collision dynamics and the subsequent re-programming of the on-board computers with novel magnetic control algorithms to re-stabilise the otherwise undamaged microsatellite and regain almost full operational mission capability.

Cost effective propulsion systems for small satellites using butane propellant

This paper will describe the work performed at the Surrey Space Centre to produce cost effective propulsion systems for small spacecraft with relatively low deltaV (ΔV) requirements. Traditionally, cold gas nitrogen systems have been used for this type of application, however they have high storage volume requirements. This can be a problem on small spacecraft, which are typically volume limited. An alternative solution is to use liquefied gases, which store as liquids, hence have reasonable density levels, and can be used in a cold gas thruster. At the Surrey Space Centre, butane has been selected as the propellant of choice. Although it has slightly lower specific impulse performance than nitrogen, it has a significantly higher storage density and it stores at a very low pressure, hence no regulation system is required. On 28 th June 2000 Surrey Satellite Technology Ltd (SSTL) launched it first nanosatellite SNAP-1. This 6.5kg spacecraft was equipped with a small cold gas propulsion system utilising 32.6 grams of butane propellant. During the propulsion system operation phase the spacecraft's semi major axis was raised by nearly 4 kilometers using the propulsion system. The design of the propulsion system will be described and the low cost features highlighted. Telemetry data will be used to describe the propulsion operations and an overall mission specific impulse will be derived. SSTL are currently under contract to build three Earth observation spacecraft for a Disaster Monitoring Constellation (DMC). Each spacecraft will weigh approx 100 kg and have a ΔV requirement of 10 m/sec. A butane system has been designed and manufactured to meet the requirements of these spacecraft. The system is based very much on the flight heritage of the SNAP-1 system, with the addition of greater propellant storage capacity. The lessons learnt from the SNAP-1 operation will be reviewed and the resulting design improvements on the DMC propulsion systems will be detailed.

Low cost planetary exploration: surrey lunar minisatellite and interplanetary platform missions

In order to meet the growing global requirement for affordable missions beyond Low Earth Orbit, two types of platform are under design at the Surrey Space Centre. The first platform is a derivative of Surrey's UoSAT-12 minisatellite, launched in April 1999 and operating successfully in-orbit. The minisatellite has been modified to accommodate a propulsion system capable of delivering up to 1700 m/s delta-V, enabling it to support a wide range of very low cost missions to LaGrange points, Near-Earth Objects, and the Moon. A mission to the Moon - dubbed “MoonShine” - is proposed as the first demonstration of the modified minisatellite beyond LEO. The second platform - Surrey's Interplanetary Platform - has been designed to support missions with delta-V requirements up to 3200 m/s, making it ideal for low cost missions to Mars and Venus, as well as Near Earth Objects (NEOs) and other interplanetary trajectories. Analysis has proved mission feasibility, identifying key challenges in both missions for developing cost-effective techniques for: spacecraft propulsion; navigation; autonomous operations; and a reliable safe mode strategy. To reduce mission risk, inherently failure resistant lunar and interplanetary trajectories are under study. In order to significantly reduce cost and increase reliability, both platforms can communicate with low-cost ground stations and exploit Surrey's experience in autonomous operations. The lunar minisatellite can provide up to 70 kg payload margin in lunar orbit for a total mission cost US$16–25 M. The interplanetary platform can deliver 20 kg of scientific payload to Mars or Venus orbit for a mission cost US$25–50 M. Together, the platforms will enable regular flight of payloads to the Moon and interplanetary space at unprecedented low cost. This paper outlines key systems engineering issues for the proposed Lunar Minisatellite and interplanetary Platform Missions, and describes the accommodation and performance offered to planetary payloads.

Intelligent optimisation of microsatellite on-board power systems

The Surrey Space Centre has pioneered the research and development of modern microsatellites technologies over the last twenty years. Obviously, the small volume of these satellites places severe constrains on power available for applications payloads on board the satellite therefore many research projects at Surrey have been carried out to reduce power consumption of both satellite platform and payloads. This paper describes one of these projects which is the use of an automatic power scheduling algorithm for the PICOSAT mission, a micro-satellite developed by SSTL under contract to the USAF SSP (Small Satellite Programme). The purpose of this algorithm is to predict the battery and memory levels on a short term basis and automatically schedule the on-board experiments activities in order to optimise the power and memory usage over the selected period, meeting the constraints and requirements for the mission. The algorithm makes use of a recursive feedback loop to reach the optimum output. An initial prototype of the algorithm has been implemented using matlab and, once fully tested, it is intended to port and run it on the satellite On Board Computer in orbit. The experiment priorities and payload characteristics are specified in separate modules, allowing the easy re-use and upgrade of the algorithm for different payload configurations under other SSTL satellites.

Observations on the reliability of COTS-device-based solid state data recorders operating in low-Earth orbit

This paper presents the results of Surrey Space Centre's experience in using different coding schemes and hardware configurations to protect data and software stored in COTS-device based memories on-board operational spacecraft in low Earth orbit.

Low-Cost Orbit Manoeuvres for Minisatellites using Novel Resistojet Thrusters

Electric propulsion using resistojets operating at low power (~100 W) and using liquid propellants have re-emerged as attractive propulsion options for small satellites. Compared to low-power pulsed plasma thrusters (PPT), the resistojet achieves 50 times the increase in thrust (~1–50 mN), which is necessary to overcome drag at solar maximum for small satellites in high drag obits (~200 km). Two resistojet thrusters have been developed at the Surrey Space Centre (SSC), which utilize a packed bed of silicon carbide (SiC) particles for the heat exchanger. A thermodynamic model has been developed to study and optimize the thruster design and a series of practical performance tests with both nitrous oxide and water as propellants has been completed at the USAF Research Lab (Edwards AFB) using the NASAJPL inverted pendulum thrust stand. Endurance tests (~300 h duration) were conducted to determine possible lifetime limitations or failure modes. The results are very encouraging and resistojet thrusters are now proposed as options for future USAF and SSC minisatellite missions. This paper describes the complete resistojet system, test results for the USAF Mighty Sat II. I and its flight performance, its performance and integration on the UoSAT-12 minisatellite and its launch in April 1999 where this novel technique was demonstrated for the first time in low Earth orbit on 26 July 1999.