Dr Andrew Viquerat

A bistable mechanism has two stable states with energy input required to move from one stable state to another. This energy barrier allows for energy storage and release which can be used to improve systems characteristics. Bistability has been used to increase the frequency range over which a kinetic energy harvester is effective, and it has been proposed that bistability can increase the efficiency of biomimetic swimming robots. However, experiments involving bistable swimming robots have typically used bistability as a means of overcoming limitations inherent to soft actuators, rather than to increase overall performance. This article implements bistability into a swimming robotic and compares performance with and without bistable action. The static thrust generation and power consumption for bistable and nonbistable configurations for five different tail morphologies are compared. Bistability is generally found to increase the system efficiency, particularly at lower frequencies where increases are observed up to 250%. The untethered swimming speed of the robot in open water is also found to increase by approximately 30%. The results show that bistability can offer direct performance benefits for biomimetic swimming, but that the bistable transmission must be well tuned to the dynamics of the rest of the system. This study asks whether bistability is a fundamentally useful feature of swimming mechanisms, or simply a tool to overcome actuation limitations. It quantifies the effects of introducing bistability into a simple bioinspired propulsion system with variable morphology. The results show efficiency gains up to 250% and speed increases of 30%, but also a need for mechanical tuning.image (c) 2024 WILEY-VCH GmbH

Deployable structures provide many advantages for lightweight and compact space applications. One of the greatest challenges in using flexible deployable structures for high-precision applications is achieving sufficient positional accuracy, including high dimensional stability. Composite material properties feature a large number of uncertainties, compromising the predictability of the dimensional stability of the structure. The harsh thermal environment of LEO can cause excessive distortions for some high precision applications, including Earth obsevation optics. In this paper, a deployable Cassegrain telescope for nano-satellites is presented. The design is based on six bistable composite booms to separate the mirrors, with a stowed volume of four CubeSat units. A sensitivity analysis based on the Morris method is carried out to assess the most influential CFRP parameters on the dimensional stability, discarding those with negligible effect. An experimental setup for a 3 degree of freedom interferometer is proposed, with the goal of determining relative tilts as well as displacements from the interferogram recorded on the detector.

Space launch vehicles provide a very limited and expensive allowance for new satellites to be put into orbit, so that spacecraft manufacturers are subjected to stringent constraints in terms of volume. Moreover, the mass budget and the overall complexity of subsystems play a signifi- cant role, especially in the design of small platforms. Deployable structures address such issues as they require smaller volumes and allow for less complicated and lighter weight mechanisms. Extendable appendages such as composite STEM booms have thermal-mechanical behaviour that could be detrimental for in-orbit operations, as they can undergo deflection accompanied by possibly unstable thermal vibrations. For high accuracy applications, a prediction of the structural performance under space environment conditions is crucial. In this paper the interaction between composite slit tubes and Solar heat flux is studied through an analytical model and finite element simulations. The main motivation is to examine the feasibility of a support structure for a telescope secondary mirror featuring coilable booms. The possibility of scaling the subsystem from nanosatellites to bigger platforms could be appealing in the Earth Observation market. The effects of changing geometrical and material parameters will be explored, especially the impact of the number of plies, stacking sequence and the uncertainties related to the thermal properties of composites. Finally, a finite element model of the telescope assembly under Solar heat flux will be analysed.

This paper presents novel ultra-light booms for solar sails and other large deployable space structures. These CFRP booms have a unique property: bistability over the whole length (BOWL), which enables simple and compact deployment mechanism designs that can reduce overall system mass. They were produced to solve some of the previously encountered problems with bistable composite tubular booms that reduced their optimal length and scalability due to local buckling phenomena when the diameter of the coil increased. A new low-cost manufacturing technique, which consists of using braids with a variable angle change over the boom length, was found to have a positive effect in reducing that tendency. An analytical model is used to explain this behavior and predict the secondary stable state properties and natural diameter of the coiled/packed boom. A 3.6 m tape spring version of these bistable CFRP booms has been designed for a 25 m2 Gossamer Sail Deorbiter of future space assets and is being considered for an upcoming solar sail demonstration mission called CubeSail. Larger booms are being designed for a new scalable roll-up solar array concept.

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.

The Surrey Space Centre (SSC) and Surrey Satellite Technology (SSTL) have collaborated in a Centre for Earth Observation Instrumentation (CEOI) project to design a deployable cassegrain telescope. Given the success of deployable structures in the past decades, attention has naturally turned to incorporating deployable structures into more complex space systems. Optical telescopes are currently flown with a fixed focal length, which limits their use to larger satellites. Incorporating a deployable system could open the door for smaller satellites to carry larger optical telescopes, making Earth observation more accessible for research and businesses. The aim of this project was to increase the capability of small satellites for earth observation by demonstrating a proof of concept breadboard model and development model of a deployable cassegrain telescope. The design uses three concentric carbon fibre barrels deployed into lockout stops by three leadscrew sets equi-spaced around the barrels. This paper will provide an overview of the design, analysis and test campaign for this deployable optical system.

Composite materials properties are affected by uncertainties that cannot be overlooked for accurate modelling predictions. In the present study, a novel implementation of statistical screening methods for sensitivity analysis on composites is proposed. The effect of uncertainties on the behaviour of the model is assessed rapidly and reliably. Despite their efficiency when models with several input factors are employed, screening approaches are rarely used in engineering. Two sampling strategies are explored, and the results for several case studies are shown and compared with statistical estimators from regression-based methods. It is shown that screening techniques manage to provide subsets of influential parameters for a variety of applications, including analytical and finite element models, with low computational cost.

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 altitude, 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 demonstrate the capturing and de-orbiting of artificial space debris targets using a net and harpoon system.

© 2014 by ASME.Two types of foldable rings are designed using polynomial continuation. The first type of ring, when deployed, forms regular polygons with an even number of sides and is designed by specifying a sequence of orientations which each bar must attain at various stages throughout deployment. A design criterion is that these foldable rings must fold with all bars parallel in the stowed position. At first, all three Euler angles are used to specify bar orientations, but elimination is also used to reduce the number of specified Euler angles to two, allowing greater freedom in the design process. The second type of ring, when deployed, forms doubly plane-symmetric (irregular) polygons. The doubly symmetric rings are designed using polynomial continuation, but in this example a series of bar end locations (in the stowed position) is used as the design criterion with focus restricted to those rings possessing eight bars.

This paper describes progress towards developing design guidelines for a number of composite bonded joints in aerospace applications. The premise of a universal failure criterion is impractical given the number of adherend-adhesive configurations and layups available. However, for a finite number of joint configurations, design rules can be developed based on experimental test data and detailed finite element (FE) modelling. By using these techniques rather than the traditional overly conservative knock down factors, more of the performance of composite bonded joints can be accessed. The work presented here experimentally studied the effect of the substrate layup, adhesive type and adhesive thickness on double-lap joint (DLJ) strength. The corresponding failure surfaces were analysed and failure modes identified. Following this, detailed FE models were developed to identify the trends associated with altering joint parameters. Finally, the stresses and strains within the adhesive and substrate were analysed at the joints respective failure loads to identify critical parameters. These parameters can provide an insight as to the stress state of the joint at failure or near failure loads, and hence its true performance.

Large deployable space structures are an integral part of reflectors, earth observation satellite antennas and radars, observation and radar targets, radiators, sun shields, solar sails and solar arrays. Launch vehicle faring sizes have not increased in the last three decades, meaning ever more efficient ways of packaging large space structures must be sought. Deployable structures come with the promise and capability of reducing payload mass substantially and allowing for very compact storage of systems during the launch phase. Gossamer structures hold particular promise for systems involving large apertures, solar panels, thermal shields and solar/deorbiting sails. The Technology Readiness Level (TRL) of a great part of these technologies is still very low (in the order of 2-3). The objective of DEPLOYTECH is to develop three specific, useful, robust, and innovative large deployable space structures to a TRL of 6-8 in the next three years. These include: a 10 m^2 (3.6 m diameter) sail structure that uses inflatable technology for deployment and support; a 5x5 m roll-out flexible solar array that utilizes bistable composite booms; and 14 m solar sail CFRP booms with a novel deployment mechanism for extension control.

The Surrey Space Centre (SSC) has performed several TRL raising projects for maturing deorbiting technology. Since 2017, the SSC has launched four deorbit sail systems hosted on a range of platforms. The InflateSail CubeSat demonstrated deployment of a dragsail and a novel inflateable boom system to provide passive spacecraft stability. The debris removal demonstration payload deployed a 10m2 transparent polymer drag sail, supported by four carbon fibre reinforced polymer (CFRP) booms. Launched on 23rd June 2017, the mission achieved rapid success through a successful deorbit from a 500km altitude orbit in only 72 days. InflateSail was the first successfully deployed Sail from a European spacecraft and the first successful use of inflatable structures on a CubeSat. The same sail deployment mechanism has been implemented on the 100kg class microsat for the RemoveDebris mission, which has successfully demonstrated Active Debris Removal with the sail deployment planned for March 2019 with re-entry expected by August 2019. Building on the experience from these missions, the Surrey Space Centre is developing a range of large area deployable sails for application to a variety of mission profiles, platform masses and orbits in order to deorbit satellites quickly and thus reduce collision risks. The SSC has delivered two such systems, designed as a self-powered 16m2 deorbiting sail installed on 250kg and 1,000kg class deployment structures. The sails were set deploy on 4th December 2018 with confirmation pending from optical observations. This paper will report the flight performance for the four recently flown sails before addressing drivers for appropriate dragsail design and operations. The paper will present a comparative study of the flight results of the sail systems utilising drag measurements, optical measurements and in-situ attitude measurements.

Thin carbon fibre reinforced polymer (CFRP) tape-springs are attractive structures for use in space-based optical instruments because of their compact stowed form, and their high dimensional stability when deployed. In this paper we present, with examples, two inexpensive methods to assess the thermal expansion properties of tape-spring structures: one based on strain gauges to obtain coupon level values, and another based on laser interferometry for structure level measurements. The strain gauge technique is a versatile approach that exploits the thermal output characteristics of the sensors. The thermal expansion characterisation of thin-composite samples measured a longitudinal expansion of 4.44 ppm/C and transverse expansion 5.95 of ppm/C. The interferometry system is designed with a view to capturing the displacements and tilts that occur when a structure with a low thermal mass, like a tape-spring, experiences a rapid change in flux, as occurs in the space environment. The homodyne interferometer is developed for three degree-of-freedom (DoF) measurements with a resolution of 10^-8 m for distances and 10^-6 rad for angles. The interferometric setup is based on the classical Michelson architecture and consists of few inexpensive commercial optical components. The source is a 0.8 mW Helium-Neon laser with a wavelength of 632.8 nm. The other elements include two spherical singlets, a right-angle prism, a cubic beamsplitter and a CMOS camera. The recorded interference fringes are analysed by using an algorithm based on Discrete Fourier Transform (DFT). Spectral information on the light intensity signals can be used to determine relative displacements and tilts. The dimensional stability of an optical payload based on high-strain composites was tested. The telescope has a deployable Cassegrain design, which uses six extendable members for the separation of its secondary mirror. Axial deformations between 20-30 microns along with angle variations of the order of 0.1 mrad were recorded with good repeatability.

The space industry has long benefited from the engineering innovations of deployable structures. Deployable structures are incrementally being used and proposed as the main structure for higher risk missions owing to the potential cost savings made in launching smaller spacecraft. To increase the revisit time of Earth observation satellites it is proposed that multiple small satellites with on-board deployable telescopes can be used in a constellation. The structure separating the primary and secondary mirrors of any deployable telescope must be sufficiently stiff so as to avoid distorted images from on-board micro-vibrations. In this paper the stiffness characteristics of three types of tape spring deployer are investigated using analytical, finite element and experimental methods. A deployer that clamps the root of a tape spring performs well in terms of the stiffness, but suffers due to larger stowed volume necessary. Partially restraining the root allows a smaller stowed volume to be used, but compromises the stiffness. An analytical resilient beam model is employed to predict the stiffness of tape springs deployed in this manner, and is shown to be in good agreement with experiment. A novel deployer is presented that aims to combine the benefits of the two previous deployers. The partially restrained deployer is identified as the most promising design via a trade matrix.

This paper presents preliminary results of an experimental study on the characterisation of the nonlinear dynamics of bistable composite shell structures. The property of bistability and snap-through motion of bistable composite structures gives them enormous potential in numerous aerospace applications including deployable spacecraft structures and vibration energy harvesting. The dynamic response of a square bistable composite unsymmetric laminate plate with two approximately cylindrical stable states supported at its centre is experimentally characterised by means of controlled amplitude and frequency harmonic base excitation. Primary resonance excitation of the first bending mode of the plate is performed using amplitude sweeps. For both stable states, the response begins with periodic single-well oscillations at low excitation amplitudes. Increasing the excitation amplitude beyond a critical value, cross-well oscillations in the form intermittent subharmonic-chaotic snap-through are initiated. At higher excitation levels, period-5, period-4, and period-3 subharmonic continuous cross-well oscillations are observed, in addition to chaotic snap-through. Future work will extend the experimental study to characterize the nonlinear dynamics of a bistable composite cylindrical shell and a bistable composite doubly curved shell such that the responses of the various types of bistable composite structure can be compared and used to inform the design of bistable composite shell structures for applications involving highly dynamic environments.

Composite storable tubular extendable members (STEMs) have been attached to cylindrical drums to deploy deorbiting sails, sensors and other instruments in space. Attaching STEMs in this way causes the cross section to flatten at the root reducing the structural performance. This reduction in properties needs to be quantified before future high risk applications can be considered. A beam model and finite element model (FEM) are presented here to quantify the detrimental effects of attaching a STEM to a cylindrical drum by way of the rotational stiffness. The beam model is derived using a finite difference approach to numerically solve for the curvature field in the transition from the root to the free end. The FEM produced in ABAQUS uses a contact simulation to deform a composite STEM. The beam model is unable to capture the crucial partially clamped root condition, but does give a qualitative understanding of the layup effects on the stiffness. The FEM shows a significant reduction in the rotational stiffness, from 88.1~Nm/rad to 14.3~Nm/rad for a [±45/0/±45] layup once the STEM is attached to the drum. The braid angle of a [±theta/0/±theta] layup is shown to have a lesser effect on the stiffness with larger braid angles increasing the rotational stiffness by 2~Nm/rad due to the shorter transition length.

Partially flat tape springs have been proposed as deployment booms that allow flat conductive elements to be embedded in the cross section. This paper presents an energy model that can predict the moment-rotation behaviour and second stable geometry of partially flat tape springs/booms. The energy model is implemented through Matlab using a constrained optimiser. The model is validated using a combination of analytical approaches and a Finite Element (FE) model. The energy model is in good agreement with previous literature for tapes with a uniform cross section radius. It is shown that the introduction of the flat section decreases the maximum snap-through moment of tape springs and increases the second stable coiled radius of booms. For a tape spring of uniform radius the snap-through moment is approximately 353 Nmm and 400 Nmm, predicted by the energy and FE model, respectively. The introduction of a 5 mm flat section, whilst maintaining the total arc length, results in a snap-through moment of 220 Nmm and 175 Nmm from the energy and FE model, respectively. For small deformations the energy model agrees well with the finite elements results.

The Surrey Space Centre (SSC) and Surrey Satellite Technology (SSTL) are collaborating in a Centre for Earth Observation Instrumentation (CEOI) funded programme to develop a deployable optical telescope system for small satellites. Current Earth observation optical telescopes are launched with a fixed focal length that dictates the size of the satellite it is launched on. A deployable telescope offers the possibility of launching on smaller satellite platforms, reducing the cost, with no reduction in resolution because the focal length in orbit is the same. A trade-off study showed that a structure of a large diameter concentric tubes would provide a suitable structure. A novel deployable telescope mechanism is designed that utilises a gear and leadscrew arrangement to deploy. A breadboard model, consisting of flight materials and 3D printed parts, shows a high level of deployment repeatability. Future work includes more extensive testing of a development model.

The Surrey Space Centre (SSC) and Surrey Satellite Technology Ltd (SSTL) have collaborated in a Centre for Earth Observation Instrumentation (CEOI) sponsored project to design a deployable Cassegrain telescope. Optical telescopes are currently flown with a fixed focal length, set by the desired magnification, which limits their use to larger satellites. Incorporating a deployable system, to extend to the fixed length from a smaller volume, could open the door for smaller satellites to carry larger optical telescopes, making Earth observation more accessible for research and businesses. The aim of this project was to increase the capability of small satellites for Earth observation by demonstrating a proof of concept breadboard and development model of a deployable Cassegrain telescope. The design shows excellent repeatability when deployed and meets the stiffness requirements to avoid on board micro-vibrations distorting images. The design uses three concentric composite barrels that deploy using a leadscrew-gear train. The design, analysis and test campaign is documented in this paper, as well as the lessons learnt on completion of the project.

The preparation of ultra-thin CFRP laminates, which incorporate a cycloaliphatic epoxy resin reinforced with polyhedral oligomeric silsesquioxane (POSS) reagent nanofiller, using out-of-autoclave procedure is reported. The influence of the amount of POSS within the laminate on the mechanical properties and surface roughness of the laminates is analysed before and after exposure to atomic oxygen (AO) to simulate the effects of low Earth orbit (LEO). The addition of 5 wt% POSS to the base epoxy leads to an increase in both flexural strength and modulus, but these values begin to fall as the POSS content rises, possibly due to issues with agglomeration. The addition of POSS offers improved resistance against AO degradation with the laminates containing 20 wt% POSS demonstrating the lowest erosion yield (1.67 x 10-24 cm2/atom) after the equivalent of a period of 12 months in a simulated LEO environment. Exposure to AO promotes the formation of a silicon-rich coating layer on the surface of the laminate, which in turn reduces roughness and increases stiffness, as evidenced by measurements of flexural properties and spectral data after exposure.

In this study, novel nanocomposites were created by incorporation of Silsesquioxane containing eight glycidylether groups (octa-POSS) into a cycloaliphatic epoxy cured by an anhydride. The developed resin system, with different nanoparticle concentrations, was used on the outer layers of an ultra-thin CFRP structure in order to provide better environmental resistance to the environment of low Earth orbit (LEO) which was tested in a ground-simulation facility. The developed resins were subjected to space-like degrading factors and their response to corrosion, radiation and elevated temperatures was monitored by mass loss, together with measuring changes in surface chemistry (ATR-FTIR), functionality development (contact angle measurement and XPS), roughness (scanning laser microscopy) and morphology (SEM). The influence of increasing octa-POSS content on thermo-mechanical properties was measured with DMTA and the strength and modulus of elasticity were determined by flexural test. The addition of octa-POSS in any loading improves the environmental resistance, however, the most significant retention of mass and mechanical and surface properties after space-like exposure was observed in the 20 wt% octa-POSS reinforced cycloaliphatic epoxy. The results presented here may contribute to the development of novel class of nanocomposites which can offer an extended service life in LEO.

Bistability in doubly curved and twisted (helical) composite slit tubes is investigated for the first time. This work establishes a natural extension in this area which has been focused on straight and until more recently, doubly curved (toroidal) tubes with positive Gaussian curvature. The model developed introduces longitudinal and transverse curvature, and twist into strips of laminated composite material. The composite is engineered to be bistable and the second stable state determined via strain energy minimisation using the Rayleigh-Ritz method. The strain energy is formulated as a function of curvature strains, longitudinal stretching and a variable middle ply fibre angle of the laminate. The second stable state forms a compact and untwisted cylindrical coil with the latter engineered by tailoring the middle ply fibre angle. A new manufacturing process capable of producing helically curved tubes using glass-fibre/polypropylene-matrix composite is presented to verify the hypothesis of this work. An untwisted coil enables the efficient stowage and deployment of new forms of bistable composite tube which adhere to similar form factors as straight and toroidal ones. By embedding electrical conductors, helical bistable composites enable new lightweight, compact and multifunctional structures for communication and sensing applications.

Aluminium inserts are frequently embedded within the composite sandwich structures used in modern sports cars. Since the inserts are used for attaching safety-critical components to the structure, the adhesion between the metal and composite needs to be strong and durable. The bond is achieved through an epoxy resin system. However, when the resin includes an internal mould release agent (IMR) to facilitate the demoulding process of the structure, the adhesion between the inserts and the composite structure can be compromised. In addition, inserts can be exposed to high treatment temperatures. As such, to ensure that high adhesion performance can still be achieved, an appropriate surface treatment should be applied on the inserts. In this paper, four commercially available surface treatments for aluminium inserts were assessed using single lap joint (SLJ) and double cantilever beam (DCB) tests. Moreover, to investigate the influence of the IMR and the high-temperature process on the joint properties, four sample manufacturing methods were used to prepare the SLJ samples: without IMR and without high-temperature process (Method 1), with IMR (Method 2), with high-temperature process (Method 3) and with both IMR and high-temperature process (Method 4). The DCB samples were only prepared using Method 4. The locus of failure for the SLJ samples prepared with Method 4 was evaluated by using X-ray photoelectron spectroscopy (XPS) and optical microscopy. It was found that a silane-based primer was sensitive to the use of both the IMR and high temperature (31% drop in the lap joint strength when both applied). A cataphoretic electrocoat was also investigated and only deteriorated when exposed to high temperature (up to 40% decrease in the lap joint strength). An epoxy-based primer did not show a significant sensitivity to the IMR, however, when exposed to high-temperatures, the joint became even stronger (up to 15% increase in the lap joint strength).

Thin metal-polymer laminates make excellent materials for use in inflatable space structures. By inflating a stowed envelope using pressurized gas, and by increasing the internal pressure slightly beyond the yield point of the metal films, the shell rigidizes in the deployed shape. Structures constructed with such materials retain the deployed geometry once the inflation gas has either leaked away, or it has been intentionally vented. For flight, these structures must be initially folded and stowed. This paper presents a numerical method for predicting the force required to achieve a given fold radius in a three-ply metal-polymer-metal laminate and to obtain the resultant springback. A coupon of the laminate is modeled as a cantilever subject to an increasing tip force. Fully elastic, elastic-plastic, relaxation and springback stages are included in the model. The results show good agreement when compared with experimental data at large curvatures.

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, 12m 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.

This paper describes a novel deployment method for rollable spacecraft booms that is suitable for high-force and precision deployment applications where reliability and low mass are critical. The piezoelectric linear shear motor controls and constrains the boom at the deployer exit with longitudinal and shear actuators through a 'pinch and push' process. This design eliminates the 'blossoming' failure mode and produces a less massive device that does not generate a magnetic field. Simulation and testing are used to evaluate the major limitation of rollable boom slippage through the device. The force at which this scenario occurs is dependent upon the friction force exerted between the tape-spring and linear motor, and hence can be characterised by the coefficient of static friction and the normal force acting between the linear motor and the rollable boom. The piezoelectric linear shear motor provides a high-force, high-precision, easily scalable deployment method.

Inflatable structures offer the potential of compactly stowing lightweight structures, which assume a fully deployed state in space. An important category of space inflatables are cylindrical booms, which may form the structural members of trusses or the support structure for solar sails. Two critical and interdependent aspects of designing inflatable cylindrical booms for space applications are i) packaging methods that enable compact stowage and ensure reliable deployment, and ii) rigidization techniques that provide long-term structural ridigity after deployment. The vast literature in these two fields is summarized to establish the state of the art.

An investigation into the bistability of positively curved laminated composite slit tubes is presented, establishing a natural extension in this area that has previously been focused on straight tubes. Curved slit tubes are modeled as the surface segments of a torus. The design space is explored through a parametric study to investigate the effect on the second stable state, representing a small coil. This includes the effects of longitudinal curvature, cross-section subtending angles, nonuniform transverse curvature, and spatially varying laminate properties. The second equilibrium state is determined through strain energy minimization using the Rayleigh–Ritz method. To verify the model, samples are manufactured from glass-fiber braid and polypropylene resin. This investigation finds 1) the initial curvature along the length of the tube has little effect on coil radius, however, the coil has a distinct barrel shape; 2) highly enclosed and 3) highly curved cross-sections result in higher edge strains of the second equilibrium, enabling identification of practical bistable tubes; and 4) conversely, the greater the initial curvature along the length of the tube, the lower the second equilibrium strain.

An inflatable-rigidisable cylindrical mast was developed as part of the InflateSail technology demonstration mission. The light-weight deployable mast is inflated using a Cool Gas Generator (CGG). To ensure long-term structural performance after deployment, the boom is rigidised by removing the residual creases in the aluminium-laminate skin material through strainrigidisation. The 1 m long and 90 mm diameter mast is folded using an origami pattern, and in its stowed configuration takes up 63 mm of height in the InflateSail Cube-Sat structure. The benefits of this folding method include minimal material deformation during deployment, a compact stowed configuration, and an open cross-section to accommodate the rapid release of inflation gas. Deployment tests showed a repeatable deployment, with minimal deviation from the intended straight path. Post deployment vibration experiments established the efficacy of strain-rigidisation in recovering the stiffness of the deployed boom. Experiments were also performed on fully rigidised booms to determine their bending and compression strengths.

An ultra-compact deployable helical antenna is presented, designed to enhance space-based reception of Automatic Identification System signals for maritime surveillance. The radio frequency performance (i.e. peak gain and directionality) is simulated at 162 MHz using ANSYS High Frequency Structure Simulator and evaluated over a range [0.5–8] of helical turns. Established and commercially available omnidirectional antennas suffer interference caused by the large number of incoming signals. A 7-turn helix with planar ground plane is proposed as a compact directional-antenna solution, which produces a peak gain of 11.21±0.14 dBi and half-power beam width of 46.5±0.5 degrees. Manufacturing the helical structure using bistable composite enables uniquely high packaging efficiencies. The helix has a deployed axial length of 3.22 m, a diameter of 58 cm, and a stowed (i.e. coiled) height and diameter of 5 cm — the stowed-to-deployed volume ratio is approximately 1:9,800 (0.01%). The use of ultra-thin and lightweight composite results in an estimated mass of 163 grams. The structural stability (i.e. natural vibration frequency) is also investigated to evaluate the risk an unstable deployed antenna may have on the radio frequency performance. The first vibration mode of the 7-turn helix is at 0.032 Hz indicating the need for additional stiffening.

A deployment solution for a parabolic sail structure for solar photon thrusters (SPTs) is presented. SPTs decouple the function of collection and reflection of light, achieving many advantages over flat solar sails. Although recent and increasingly realistic studies have concluded SPTs an unattractive option, the motivation behind this work is to progress the novel SPT concepts by resolving two problems identified: presenting a feasible solution for deployment and maintaining tight control over the collector shape; and addressing the space durability of carbon-fibre reinforced epoxy-resin composites for long duration solar sailing missions. Laterally curved bistable reeled composites were manufactured in such a way that their beneficial structural properties and bistable behaviour have been complimented with improved environmental resistance. This was achieved by implementing a cycloaliphatic based coating system reinforced with silicon nano-additive. The effect of curvature and additive on the natural frequency were investigated. In addition, response to vacuum outgassing, UV resistance, surface degradation due to atmospheric oxygen and thermal stability were investigated and improved.

The bistability of a toroidal slit tube is modeled using the Rayleigh-Ritz method. Approximate explicit expressions for the original stable deployed geometry, and the deformed stowed geometry are used to derive forms for the bending and stretching strain energy. The surface of a torus has varying Gaussian curvature, requiring a new approach to the modeling and analysis of the stable configurations. A comparative study with established straight-BRC models was conducted from which the doubly curved-BRC model presented here predicts second stable state coil radii with 96.25% agreement.

The design of a deployable structure which deploys from a compact bundle of six parallel bars to a rectangular ring is considered. The structure is a plane symmetric Bricard linkage. The internal mechanism is described in terms of its Denavit-Hartenberg parameters; the nature of its single degree of freedom is examined in detail by determining the exact structure of the system of equations governing its movement; a range of design parameters for building feasible mechanisms is determined numerically; and polynomial continuation is used to design rings with certain specified desirable properties. © 2013 Elsevier Ltd.

Bistable composite shells patented as Bistable Reeled Composite (BRC) booms have the potential to be used as lightweight structural elements for a number of space applications. This paper details an approach to increase the natural frequency and stiffness of BRCs. The motivation for this research is the desire to increase the scalability of a flexible "roll-up" solar array which, in its deployed state, consists of two cantilevered BRCs supporting a flexible Photo Voltaic (PV) cell covered blanket between them. A Finite Element (FE) numerical model is combined with a nonlinear constrained optimization to maximize the natural frequency of BRC booms with respect to the fiber orientation angles and ply discontinuity locations. The results demonstrate that careful selection of the fiber orientation angles and the location of step thickness variations can significantly optimize the natural frequency. Experimental verification of the vibration characteristics of optimized BRC booms has also been conducted. Finally, stability analysis of the optimized BRC booms under bending has been carried out using FE simulation to quantify the Maximum Rotational Acceleration (MRA) that they can take before failure.

The purpose of this study is to demonstrate the properties of novel nanocomposites, based on cycloaliphatic epoxy resin additionally reinforced with silicon-containing nanostructures (mono- or octa-functional POSS or nanosilica). The changes in properties are discussed for the varied combinations of cycloaliphatic epoxy with a curing agent (cycloaliphatic amine or anhydride) and the nanomodifier. The in uence of modification on thermal stability, curing behaviour, morphology, surface chemistry, and topography were studied with TGA, DSC, ATR-FTIR, XPS and LCM. The results show that when POSS and/or nanosilica are incorporated to the cycloaliphatic matrix they in uence curing behaviour and glass transition temperatures (Tg), where mono-POSS increases Tg and octa-POSS decreases it with respect to nanosilica. Mono-POSS produces silicon-rich surfaces but tends to agglomerate and increase surface roughness. Octa-POSS and nanosilica penetrate the polymer matrix more deeply and disperse more easily. From the selected modifiers, octa-POSS shows the highest thermal stability.

This paper presents an overview of the different gossamer sail flight projects being undertaken at the Surrey Space Centre. The missions consist of a 25 m2 solar sail to be launched in Q1 2014 (CubeSail), a gossamer deorbiter for future European space assets (DGOSS), a scalable sailcraft that will demonstrate satellite deorbiting in Low Earth Orbit (DeorbitSail), and a drag sail that uses inflatable and rigidizable technology to be flown as part of the QB50 mission (InflateSail). The key technologies currently being developed for each project will be summarized and the most relevant scientific results presented.

Space data services provide the largest market value to the global space industry. With the increasing use of small satellites that lower costs and lead times, the entrepreneurial space age has begun. However, advances in technology miniaturization are required to improve small satellite capabilities, which are limited by small volumes and low power consumption. This paper presents a deployable antenna for small satellites capable of achieving high-gain radiation performance despite being ultra-compact. The antenna is a helically curved boom that is deployed from a coil. The boom is an open slit tube. A ground plane comprised of four metallic booms supporting a sparse mesh is deployed by stored strain energy. A prototype of the antenna system has been built to test and validate the deployer mechanism, deployment strategy, and dimensional stability of the helical antenna and ground plane. The architecture builds on proven space technology, specifically the deployer mechanism of the InflateSail de-orbiting drag sail that successfully demonstrated low-Earth orbit space-debris removal in 2017. In this work, the deployer unrolls the helical boom whilst the sail itself is repurposed to boost the radiation performance of the helical antenna.

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. •Description of the InflateSail QB-50 Spacecraft and mission and its results.•First demonstration of cool gas generator inflated inflatable structures in Europe.•First successful European demonstration of using a drag sail to cause re-entry of a spacecraft.•First detailed observations of orbit and body descent using such technology.•Discussion of future application to tackling the space debris problem.

Deployable coilable booms are extendible structures that have been used in a number of space and terrestrial applications. Fiber reinforced polymer (FRP) composite variants of these deployable booms have advantages over the metallic versions in the form of higher specific stiffness and greater design flexibility. A deployment failure mode called ‘blossoming’, in which the boom unwinds and extends within the deployment mechanism, can occur if an excessive load is applied to the boom tip. Blossoming can be mitigated by using compression rollers which radially constrain the coiled boom. An energy method is used to model the composite boom during deployment, and to predict the tip force a boom can withstand before blossoming occurs. The analytical results are compared with experimental results. The effects of the boom material properties and geometric parameters are investigated to provide more guidance in the design of deployable coilable boom systems.

The advances of carbon usage for Carbon Fibre Reinforce Polymer (CFRP) structures led to multiple applications in a large number of industries. This chapter presents methods for CFRP material characterization and usage for aeronautic, automotive and satellite applications. The major CFRP components used for antennas and microwave applications within these industries are presented. The accelerated adoption of carbon-based composites, current challenges and future directions are also reported.

The non-linear flexural behaviour of a tape spring — whose cross section has a uniform curvature — is produced by the presence of an elastic instability. However, these are not the only structures that possess non-linear flexural behaviour. Panel-springs and bridged tape springs are examples of structures that can also produce this behaviour. Predicting the flexural behaviour of these structures is more challenging because their cross section is non-uniform. This work presents an analytical model that can predict the opposite-sense flexural behaviour of thin strips that have a non-uniform cross section. The novel attribute of this model is in the utilisation of a univariate polynomial series to describe the initial cross section of the strip. The model is applied to a range of bridged tape springs and panel-springs, with the results compared with solutions obtained from the Finite Element (FE) method. A consistent error of less than 5.5% is achieved between the analytical model and the FE model when the snap-through moment and propagation moment are compared. The analytical model is also compared with independent experimental data of the flexural behaviour of a traditional carpenter’s tape measure. An error of less than 1% and 7% of the snap-through moment and propagation moment, respectively, is achieved.

A 1 m long inflatable-rigidizable mast was developed as a payload for InflateSail: a 3U CubeSat technology demonstration mission. The thin-walled cylindrical mast consists of an aluminum-polymer laminate, and long-term structural performance is ensured through strain-rigidization: the packaging creases are removed through plastic deformation of the aluminum plies. During ground tests it was observed that after rigidization the internal pressure dropped more rapidly than could be accounted for by leakage of inflation gas alone. It was hypothesized that viscoelastic behaviour of the laminate material causes a further, time-dependent (order of seconds), increase in cylinder diameter, with a corresponding drop in internal pressure. Additional experiments revealed an increase in diameter, including large visco-elastic shear in the adhesive of the lap joint. This was not found to be sufficient to fully account for the observed reduction in pressure. An increase in temperature of the gas during inflation, with subsequent cooling down to ambient is thought to cause the additional pressure drop.

The natural frequency of cantilevered bistable carbon/epoxy reeled composite (BRC) slit tubes constructed from combinations of braided and unidirectional (UD) plies is optimized with respect to ber orientation angles and laminate stacking sequences. BRC tubes have the same geometry as a carpenter's tape; however, they also have a second stable con guration in the coiled state, and it is considered likely that the coiled state diameter will be xed by the geometry of the deployment mechanism or its housing. The optimization process uses the BRC coiled diameter as a constraint, and the maximum and minimum physically achievable braid angles as bounds. Both individual tubes, and a simple deployable solar array concept are analyzed. It is observed that the braid angle, rather than ply location in the stack is of greater importance when optimizing long slender or shallow BRCs, whereas both factors must be considered in shorter BRCs. The sensitivity of natural frequency and coiled diameter to braid angle perturbations indicates the importance of precision during manufacture.

Coiled deployable booms have seen a wide variety of uses both in space and terrestrial applications. CubeSail, a solar sailing mission at the Surrey Space Centre, uses coiled deployable booms to extend and give structure to its thin film sail. During deployment testing a problem was found where the coiled boom would unwrap or "blossom" within the deployer instead of deploying the sails. During an investigation into this blossoming problem it was found that a coiled tape spring could act as a spring in tension or compression; this paper aims to describe this phenomenon.

Deployable booms are an essential part of the deployable structures family used in space. They can be stowed in a coiled form and extended into a rod like structure in an action similar to that of a carpenter’s tape measure. “Blossoming” is a failure mode that some boom deployers experience where the booms uncoil within the deployer instead of extending. This paper develops a method to predict the force that a boom can exert before blossoming occurs by using the strain energy stored in the coiled boom and in the compression springs. An experimental apparatus is used to gain practical results to compare to the theory.

A self-contained inflatable and rigidizable truss based substructure, its constraining mechanism, and stowage enclosure were developed for the RemoveDEBRIS technology demonstrator. RemoveDebris is a European Commission FP7 funded mission due for launch in late 2016. The hardware discussed in this paper will be integrated with the DebrisSat-1 microsatellite. During the course of the mission, active debris removal will be achieved by capturing DebrisSat-1 with the aid of a net fired from the primary platform. The inflatable module is key to this experiment as it allows the simulation of a much larger piece of debris than would be possible with a CubeSat alone. Following its capture, the inflatable structure will continue with its second objective as an end of life removal solution by passively drag augmenting DebrisSat-1's orbit to re-entry. The inflatable structure is constructed with six aluminum-polymer laminate cylindrical booms. These are connected in an axial manner to form a regular octahedron with a cross sectional area of 0.5 m2. A set of eight triangular polyester film segments or sails enclose the structure. The segments serve a dual purpose: firstly to increase the aerodynamic drag of the spacecraft, and secondly to distribute impact loads between the compressive inflatable members. A single cool gas generator (CGG) is utilised to deploy and rigidize the structure. This paper examines the development of the inflatable module from the early conceptual stages to the pre-qualification test level.

Large deployable space structures are an integral part of reflectors, earth observation satellite antennas and radars, observation and radar targets, radiators, sun shields, solar sails and solar arrays. Launch vehicle faring sizes have not increased in the last three decades, meaning ever more efficient ways of packaging large space structures must be sought. Deployable structures come with the promise and capability of reducing payload mass substantially and allowing for very compact storage of systems during the launch phase. Gossamer structures hold particular promise for systems involving large apertures, solar panels, thermal shields and solar/deorbiting sails. The Technology Readiness Level (TRL) of a great part of these technologies is still very low (in the order of 2-3). The objective of DEPLOYTECH is to develop three specific, useful, robust, and innovative large deployable space structures to a TRL of 6-8 in the next three years. These include: a 10 m2 (3.6 m diameter) sail structure that uses in atable technology for deployment and support; a 5 × 1 m roll-out exible solar array that utilizes bistable composite booms; and 14 m solar sail CFRP booms with a novel deployment mechanism for extension control. © 2013 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

A method of recovering laminate ply stacking sequences from a set of up to twelve lamination parameters using polynomial homotopy continuation techniques is presented. The ply angles are treated as continuous variables, and are allowed to take any value between -90 and +90 degrees. The individual plies are assumed to be orthotropic and have constant stiffness. The method is fully deterministic, and does not rely on an optimisation process to establish the stacking sequence. Polyhedral continuation methods are used to limit the solution space in which the stacking sequences are sought. The method can reliably find every stacking sequence solution that exists to achieve a precisely specified set of lamination parameter "targets", with the number of real solutions to a feasible combination of target properties found to vary from 1 to over 100. The same method is also demonstrated to be able to find stacking sequences to satisfy a set of specified ABD stiffness matrix terms, as might be required following a direct-stiffness modelling design process.

Communications present a major bottleneck for small-satellite functionality given their extremely small volumes and low power. This work addresses this gap by presenting an ultra-compact, high-gain deployable helical antenna designed for space-based reception of Automatic Identification System signals at 162 MHz for maritime surveillance. The radio frequency characteristics of helically curved ribbons are investigated and optimized through a parametric study of the helical and ground plane geometry. Square, planar ground planes of various size and thickness, and a range of helical ribbon widths are studied. Both are modeled as perfect electrical conductors using ANSYS High Frequency Structure Simulator. Simulation results indicate that the addition of a ground plane centered and positioned at the base of the helical antenna element: 1) reduces back lobe radiation and 2) enables optimization of the radiative performance through adjusting the antenna geometry i.e. the peak gain may be increased by 3.5% (on average) for each additional helical turn — 1-8 helical turns are simulated. The half-power beam width may also be improved indefinitely by adding more helical turns. The most focused beam presented, 40 deg, is produced by an 8-turn helix, which is 58 cm in diameter and has an axial length of 3.68 m. Two ground plane sizes are considered, with the largest, which is four times larger in area, producing 5% higher peak gain. Conversely, the ground plane size had negligible effect on the half-power beam width in long helices (i.e. >3 helical turns). Increasing the helical ribbon width in steps of 10 mm was found to improve the peak gain by 8% on average in long helices.

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.

Bistable reeled composite booms (BRCs) constructed from braided carbon/epoxy plies are suitable candidates for use as extendible booms or as elements of large deployable space structures. However, without modification, BRCs have an open section which limits their torsional stiffness, and makes them prone to collapse under low bending moments. In this study a “roll-up” deployable photovoltaic (PV) solar array with two side-by-side extendible BRCs is used as a case study to numerically analyse the dynamic behaviours of BRCs on spacecraft undergoing rotational manoeuvres. The BRCs have rotational accelerations applied to their roots to simulate the effect of being attached to a manoeuvring spacecraft. Budiansky-Hutchinson criterion is used to define an instability failure point based on a change in cross-sectional shape. This was used to estimate the maximum angular acceleration. While it is extremely difficult to replicate the behaviour of a large flexible lightweight structure in microgravity on the ground, an experiment to determine the point of collapse of BRCs under gravity were used to verify the simulation results.

Thin-walled tape springs are extended from cylindrical deployment drums to support instruments and sensors in a cantilevered manner on spacecraft. Attaching tape springs onto the deployment drums results in a partially flattened and partially restrained cross section, which is far from the ideal case of a fixed-free beam. The consequence is a more compliant root condition that has the potential to couple and amplify on-board microvibrations with the natural frequency of the extended instrumentation. In this paper it is shown that an Euler-Bernoulli beam model can be used to calculate the natural frequency of drum-deployed tape springs using elastic boundary conditions to represent the root condition. A Finite Element (FE) model and experimental data are used to validate the beam model's correctly predicted relationships between the natural frequency and tape spring length, f 1,Res. ∝ L −1.5 , and its rotational stiffness, f 1,Res. ∝ k 0.5 rot. For the investigated beryllium copper tape springs the FE model and beam model are in excellent agreement with experiment, with the error < 10%. For carbon fibre epoxy tape springs there is also strong agreement between the FE model and experiment, and approximately 10-20% error with the beam model. On a scale from a hinged beam to a fixed-free beam, the non-* dimensionalised beam equation reveals that the drum-deployed tape springs are close to the hinged beam end of the scale. Two stiffening methods are proposed to increase the stiffness of the tape springs, and hence move the tape springs towards being a fixed-free beam.

Research into reactive collision avoidance for unmanned aerial vehicles has been conducted on unmanned terrestrial and mini aerial vehicles utilising active Doppler radar obstacle detection sensors. Flight tests conducted by flying a mini UAV at an obstacle have confirmed that a simple reactive collision avoidance algorithm enables aerial vehicles to autonomously avoid obstacles. This builds upon simulation work and results obtained using a terrestrial vehicle that had already confirmed that active sensors and a reactive collision avoidance algorithm are able to successfully find a collision free path through an obstacle field.

This paper describes the scalability analysis of bistable Carbon Fibre Reinforced Plastic (CFRP) tubes for space applications, with the aim of attaining a better understanding of the scaling laws of Bistable Reeled Composite (BRC) tubes. BRCs with substantially higher natural frequency are designed. The application for this work is a deployable solar array, which uses two BRC tubes to support a membrane containing flexible photovoltaic cells. Novel types of bistable tubes with stepped thickness changes, tapered diameter and reduced included angle are proposed to improve the natural frequency. Finite Element (FE) modelling and experimental verification have been used to study the vibration characteristics of the proposed BRC tubes. An FE model is combined with an optimization loop to improve the natural frequency with respect to the fibre angles within the laminate of the bistable tubes. The results demonstrate that the introduction of step changes in laminate thickness at certain locations, and careful selection of fibre angles can significantly improve the natural frequency.

This paper will describe the development, testing, and preliminary flight results of a 16 m2 drag-deorbiting sail system launched as part of the SSO-A mission in late 2018. The sail consists of four triangular quadrants of 12 micron thick mylar metalized membrane, which are `Z' folded before being wrapped around a central free-spinning hub. A brushless DC motor is used to drive the extension of four co-coiled bistable carbon booms, which in turn pull the wrapped sail off the hub during deployment. The packaged module sits within a housing with four fold-out panels, and is attached to an outer panel of the host spacecraft. On-board electronics and a battery pack allow completely independent operation from the host spacecraft bus, with the system programmed to deploy at a predetermined time after a separation indication. Two commercial units of the sail system were launched on Monday December 3 2018, on 260 kg, and 1250 kg host spacecraft to a Sun Synchronous orbit at an altitude of 575 km, with deployment of the sails taking place shortly after launch. The time to demise for these systems is estimated to be between 5 years for the 260 kg, and 13 years for the 1250 kg spacecraft, with the spacecraft already having lost 750 m and 300 m of altitude respectively.

Tape springs are a type of thin-walled deployable boom that are used extensively in the space industry to deploy sensors, drag sails and antennas. When a tape spring is stowed it coils into a cylindrical shape and so deployment drums are manufactured as cylinders to match. The consequence of this is that when the tape spring is deployed a portion of the cross section remains flat against the cylindrical drum. This has the effect of reducing the stiffness of the tape spring. In this paper a Finite Element (FE) model is presented to capture this reduction. An experimental method for validating the FE model is also presented on which Beryullium-Copper (BeCu) and glass fibre polypropylene (PP) composite tape springs were tested. The FE model is able to predict the rotational stiffness of the BeCu tape springs more accurately than the composite tape springs. The disagreement in the case of the composite tape springs is attributed to inaccuracies in the available data for the mechanical properties, and the assumption that the tape spring does not compress through the thickness. Increasing the drum length has been shown to decrease the rotational stiffness due to increased flattening at the root. BeCu tape springs show an increase of 82\% in the rotational stiffness when the flattened drum region reduces from 90\% to 30\% of the tape springs' width. Glass fibre PP tape springs with a layup of [$\pm$34/0/$\pm$34] show an increase of 65\% when the drum length percentage reduces from 82\% to 27\%. A parametric study showed that the rotational stiffness can be significantly improved with the introduction of local root reinforcing plies.

The EC FP7 RemoveDebris mission aims to be one of the world's first Active Debris Removal (ADR) missions to demonstrate key technologies in-orbit in a cost-effective ambitious manner, including: net capture, harpoon capture, vision-based navigation, dragsail de-orbitation. The mission will utilise two CubeSats as artificial debris targets to demonstrate the technologies. In early 2018, the main 100 kg satellite will launch to the International Space Station (ISS) where it will be deployed via the NanoRacks Kaber system into an orbit of around 400 km. The mission comes to an end later in 2018 with all space entities having been de-orbited. The mission contains several mechanisms and inflatable systems that will be addressed in the paper: CubeSat deployable, dragsail mechanisms, harpoon target extender. One of the ejected CubeSats, DS-1, which acts as the artificial debris and is used in the net experiment, deploys an inflatable structure to maximise volume and enable faster de-orbitation. The dragsail system utilises a two stage deployment - firstly using a 1 metre inflatable rigidisable mast to move the sail system away from the main platform - then deploying a series of coiled carbon fibre booms to draw out a 10 metre squared aluminised Kapton sail. Finally, the harpoon is fired at a deployable target plate, which is extended 1.5 metres away from the platform by a deployable carbon fibre boom. The varying mechanisms technologies in the mission form a mix of core technologies (such as the dragsail) and in-space support technologies (such as the target extender). The paper will examine for the selected mechanisms the design of the systems and details of the testing methodology and results for ground (both functional and environmental). The paper will also detail some in-flight mission results with regards to the experiments. Examination of mechanism performance will also be addressed, along with usage and scalability potential for future ADR missions. Future mega-satellite constellations are now being proposed, where hundreds to thousands of satellites are being launched into orbit. A coherent strategy, along with technological and platform developments, is needed for de-orbiting, re-orbiting, or servicing of such constellations. The RemoveDebris mission is a vital prerequisite to achieving the ultimate goal of a cleaner Earth orbital environment, and is a core step in the development of active removal vehicles, or on-orbit servicing vehicles of the future.

While a solar or drag sail segment is packaged, packaging folds are introduced across the segment. Understanding elastic-plastic behaviour in polymers undergoing large displacements is important in forming processes due to the tendency of shape alteration or ploua springback once the forming force is removed. By tracking the stresses and strains during folding, residual stress distribution in the polyner segment is known. A numerical model has been derived for the compression of a polymer film in between two plates. The polymer film is modelled as an isotropic cantilever under the action of a bending moment undergoes elastic-plastic, relaxation and springback. The fundamental Euler-Bernoulli equation has been adapted for large displacements. The results show good agreement at small and large curvatures for the displacement of the foil.

Carbon fibre reinforced plastics (CFRP) can be found as structural components in various space applications, including the field of ‘gossamer’ structures used as deployable masts, antennas or hinges. Many of these applications are missions in low Earth orbit (LEO), which is a particularly hazardous environment for polymers and organic materials, such as epoxy resins used in CFRP manufacturing. The incorporation of silicon derivatives in epoxy resin based CFRPs in order to create hybrid organicinorganic networking has been suggested as a way to prolong the life span of ultra-thin composite structures. Two ways of modification were considered during this study; incorporation of polyhedral oligomeric silsesquioxane (POSS) nanoparticles to create so called nanocomposites, and a mixture of POSS with a flexible polydimethylsiloxane (PDMS) in order to achieve a smooth, silicon-rich protective surface. Both mono-functional and octa-functional POSS were selected and their compatibility with aliphatic amine/epoxy resin system was evaluated. The conducted experiment was inspired by the Design of Experiments (DoE) theory to validate the degradation of properties. The suggested method allows the magnitude of individual effects that contribute to the composite ageing and the effectiveness of various silicon derivatives to be evaluated. The results of this study contribute to the development of protection strategies which could help lower the rate of LEO induced degradation of ultra-thin CFRP masts.

In this work, thin carbon fibre reinforced plastic (CFRP) structures were coated with an organic-inorganic resin system for improved resistance to the low Earth orbit (LEO) environment. Thin structures of this type have been proposed for use in solar sails and other large deployable structures. The ultra-light, long extendible members were primarily composed of aromatic, high stiffness epoxy resin (TGDDM) cured with aromatic polyamines. This resin system was chosen because the high aromatic content provides excellent stiffness and creep resistance that are critical for this application. However, the resin’s aromaticity contributes to degradation by ultraviolet radiation and oxidation. The proposed solution involves shielding aromatic rings and organic chemical bonds that are prone to degradation by UV rays, with a cycloaliphatic resin system additionally reinforced with silicon nanostructures. By applying surface coating a significant decrease in roughness was observed and the surface degradation due to UV radiation prevented.

Deployable coilable booms have many advantages for use in space, but these kinds of structures sometimes experience a deployment failure mode called `blos- soming'. Blossoming of a coiled boom occurs when the boom stops deploying, and instead unwinds and expands within the deployer. This can occur even in the presence of sprung rollers used to constrain the coil. In the blossoming process, friction between the layers of the coil plays an important role that has only been brie y considered in previous work. In order to be able to model and predict the onset of this phenomenon more precisely, the pressure distribution between adjacent layers of the coil must be known. This paper establishes a numerical model to investigate the pressure distribution within a coiled open- section tape spring boom, then combines this result with theoretical analysis to produce an estimate of the maximum tip force that a deploying boom can with- stand before the onset of blossoming. The effect of the roller springs' stiffness and the boom friction coefficient are also taken into account in the simulation. The results of the theoretical analysis and numerical simulation are compared with previous experimental results to provide some practical verification.

The vibration characteristics of cantilevered straight and curved carbon/epoxy bistable reeled composites (BRCs) have been investigated. The tube length, cross-section radius, subtending angle, longitudinal curvature and number of plies - design parameters were investigated for their effects on the vibration modes. The boom length affects the frequency the most, which is found to be inversely proportional to the square of boom length, in addition to ABAQUS simulation results showing that frequency is proportional to curvature. Short, three-ply carbon/epoxy samples were manufactured and tested. A regime change from short (48.5cm) to slender (≈150cm) tubes was observed, signified by curved tubes exhibiting higher vibration modes in a particular plane than the straight ones in simulation - highlighting the scalability of curved BRC applications. Recommendations for the upcoming CleanSpace One, EPFL space mission which uses curved tubes for its capture mechanism, are discussed. Dynamic stability analysis was performed by simulating increasing rotary accelerations, causing the cantilevered BRCs attached to a spacecraft to rotate. A failure point derived from the Budiansky-Hutchinson criterion was developed to determine the maximum rotation acceleration - the critical value by which the tube loses stability.

In this paper, a homodyne single-beam interferometer for three degrees of freedom (DoF) measurement is used to assess the dimensional stability of a deployable telescope for small spacecraft. The interferometric system is based on a Michelson interferometer concept, and the number of components is kept to a minimum. The rig is composed of a HeNe laser at 632.8 nm, two lenses, a prism, a beamsplitter and a CMOS camera. This makes the setup very attractive for low-cost and low-complexity solutions, and its performance can be readily improved by upgrading the hardware according to need. The algorithm is based on the Discrete Fourier Transform (DFT) of the spatial interference pattern detected by a CMOS sensor. Spectral information on fringe density and orientation can be translated into both relative displacements and tilts. The system can easily measure displacements with nanometer resolution and angle variations with microrad resolution. The developed architecture was suitable to determine the thermal deformations of the optical payload. Maximum relative displacements of about 30 microns and angle variations of the order of 0.1 mrad were obtained experimentally, with good repeatability.