Hoskin A, Viquerat AD (2016) An Analysis of a Coiled Tape Spring During Compression and Extension,
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.
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.
Viquerat AD, schenk M, Sanders B, Lappas V (2014) Inflatable Rigidisable Mast for End-of-Life Deorbiting System,
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.
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.
Viquerat AD (2006) A Continuous Wave Doppler Radar System for Collision Avoidance Applications,
This document contains a final year Engineering Thesis in the design and construction of a set of sensors capable of detecting obstacles at a distance. The device is intended for use with small air and ground vehicles for obstacle detection and tracking. The device has varying applications depending on the sensor configuration, and can function as either a proximity detector, or as a speed and position tracker. The hardware is designed to be inexpensive, small and lightweight. This project involved the design and construction of a sensor simulator, a sensor array pattern, electronics to power and operate transmitters and receivers, and microcontroller software.
The key features of the system are its small size, low weight, low power requirements and its ability to employ a synthetic aperture technique known as Doppler beam-sharpening to refine its field of view. This is believed to be the first application of Doppler beam-sharpening in a collision avoidance sensor application.
Viquerat AD, Blackhall L, Reid A, Sukkarieh S, Brooker G (2008) Reactive Collision Avoidance for Unmanned Aerial Vehicles Using Doppler Radar, Field and Service Robotics 42 pp. 245-254 Springer
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.
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.
Viquerat AD, Schenk M (2016) Viscoelastic Effects in Metal-Polymer Laminate Inflatable Structures,
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.
Viquerat AD, Schenk M, Lappas V, Sanders B (2015) Functional and Qualification Testing of the InflateSail Technology Demonstrator,
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.
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.
Secheli G, Viquerat AD, Lappas V (2015) An Examination of Crease Removal in Rigidizable Inflatable Metal-Polymer Laminate Cylinders,
Secheli G, Viquerat AD, Aglietti G (2016) Mechanical Development of a Novel Inflatable and Rigidizable Structure, 3rd AIAA Spacecraft Structures Conference Proceedings American Institute of Aeronautics and Astronautics
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.
© 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.
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.
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.
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
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.
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.
Polynomial continuation, a branch of numerical continuation, has been applied to several primary problems in kinematic geometry. The objective of the research presented in this document was to explore the possible extensions of the application of polynomial continuation, especially in the field of deployable structure design. The power of polynomial continuation as a design tool lies in its ability to find all solutions of a system of polynomial equations (even positive dimensional solution sets). A linkage design problem posed in polynomial form can be made to yield every possible feasible outcome, many of which may never otherwise have been found. Methods of polynomial continuation based design are illustrated here by way of various examples. In particular, the types of deployable structures which form planar rings, or frames, in their deployed configurations are used as design cases. Polynomial continuation is shown to be a powerful component of an equation-based design process. A polyhedral homotopy method, particularly suited to solving problems in kinematics, was synthesised from several researchers? published continuation techniques, and augmented with modern, freely available mathematical computing algorithms. Special adaptations were made in the areas of level-k subface identification, lifting value balancing, and path-following. Techniques of forming closure/compatibility equations by direct use of symmetry, or by use of transfer matrices to enforce loop closure, were developed as appropriate for each example. The geometry of a plane symmetric (rectangular) 6R foldable frame was examined and classified in terms of Denavit-Hartenberg Parameters. Its design parameters were then grouped into feasible and non-feasible regions, before continuation was used as a design tool; generating the design parameters required to build a foldable frame which meets certain configurational specifications. iv Two further deployable ring/frame classes were then used as design cases: (a) rings which form (planar) regular polygons when deployed, and (b) rings which are doubly plane symmetric and planar when deployed. The governing equations used in the continuation design process are based on symmetry compatibility and transfer matrices respectively. Finally, the 6, 7 and 8-link versions of N-loops were subjected to a witness set analysis, illustrating the way in which continuation can reveal the nature of the mobility of an unknown linkage. Key features of the results are that polynomial continuation was able to provide complete sets of feasible options to a number of practical design problems, and also to reveal the nature of the mobility of a real overconstrained linkage.
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.
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
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.
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 (H150cm) 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.
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.
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.
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.
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.
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.
Within this study, a development of a protection strategy for ultra-thin CFRP structures from degrading effects of low Earth orbit (LEO) is presented. The proposed strategy involves an application of a modified epoxy resin system on outer layers of the structure, which is cycloaliphatic in its chemical character and reinforced with POSS nanoparticles. The core of the CFRP structure is manufactured using a highly aromatic epoxy resin system which provides excellent mechanical properties, however, its long-term ageing performance in space is not satisfactory, and hence a surface treatment is required to improve its longevity.
The developed resin system presented in this thesis is a hybrid material, designed in such a way that its individual constituents each contribute to combating the detrimental effects of radiation, atomic oxygen (AO), temperature extremes and vacuum induced outgassing of exposed material surfaces while operating in LEO. The cycloaliphatic nature of the outer epoxy increases UV resistance and the embedded silicon nanoparticles improve AO and thermal stability.
During the study, a material characterization of the developed cycloaliphatic epoxy resins was performed including the effects of nanoparticles on morphology, curing behaviour, thermal-mechanical properties and surface chemistry. Following on that, the efficacy of the modified resin system on space-like resistance was studied. It was found that when the ultra-thin CFRP structures are covered with the developed resin system, their AO resistance is approximately doubled, UV susceptibility decreased by 80\% and thermal stability improved by 20\%.
Following on the successful launch of the InflateSail mission earlier this year, which demonstrated a sail deployment and a controlled de-orbiting, the findings of this study are of importance for the future generation of similar, but significantly longer missions. Ensuring resistance of CFRP structures in a highly corrosive LEO environment is a critical requirement to make their use in space applications truly feasible.
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.
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.
Deployable structures play an important role in space applications as they minimise the volume required by large structures such as antennas, solar panels, reflectors or de-orbiters. A low cost and mass option, relies on the use of airtight inflatable structures. Over the years several rigidization methods have been developed, each with their strengths and weaknesses, however due to the simplicity of aluminium based metal-polymer laminates, this class of shells have been successfully flown on a series of legacy missions. Metal polymer laminates are typically three-ply constructions where two foils of ductile annealed aluminium sandwich a polymer core. Structures such as sphere and columns may be constructed from flat sheets of material. The envelopes are then packaged. Pressurised gas, typically nitrogen is realised into the envelope to achieve deployment. To rigidize the structure the pressure is further increased to a value slightly higher than the yield point of the metal foils. By conducting repeated rigidization experiments it was observed that residual fold creases remain present in metallic shell. As metal laminates rely on the structural integrity of the shell for strength, it is important that the extent of the initial imperfections is known. The collapse load of laminate columns is significantly reduced by this effect. If care is taken during packaging and construction of these structures, packaging residual creases remain the largest source of imperfections.
To observe closely the folding process, a 3D laser and SEM images have been taken at various steps during folding. To understand this mechanism these results were compared against the results from literature. It has been found that for a `Z' folded column the longitudinal creases flatten more than the circumferential creases.
A numerical model has been derived for the elastic-plastic bending and springback of a metal film and metal-polymer-metal laminate. In the presented work this approach replicates the introduction of a typical `V' fold and relaxation once the load has been removed. The system of differential equations was solved in MATLAB using ode45. To simplify the analysis a bilinear stress-strain profile with plane strain has been attributed to the metal film. The results have been validated with good agreement against experimental results and FEA analysis conducted in ABAQUS. Two three ply aluminium-polymer-aluminium flight ready laminates have been used as the experimental benchmark. The derived model may be adapted for different laminate configurations. It is known that it is difficult to quantify the mechanical properties of thin aluminium films, in particular the Young?s modulus. Several results from literature are discussed and the solutions proposed is outlined.
The lessons learn from this research project have been applied to the development of a novel rigidizable aluminium-BoPET based deployable structure. The structure consists of six laminate booms to connect to form a cuboid structure with a cross-sectional are of 0.5 m2. The structure and support systems were design to occupy the volume of single CubeSat Unit. The deployable will be flown on the RemoveDebris ADR technology demonstrator.
This thesis describes progression towards developing an enhanced design methodology for laminated composite bonded joints in aerospace applications. The premise of a universal failure criterion is impractical given the number of adhesive-adherend configurations available. However, for a finite number of joint configurations, design rules can be developed based on experimental test data and detailed finite element 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. While complex damage modelling techniques are available, the additional material data and analysis time required renders them not suitable for the vast majority of time-sensitive industrial applications.
Initially, the work presented in this thesis experimentally studied the effect of the substrate material, substrate layup, adhesive material and adhesive thickness on several laminated composite bonded joint configurations. The corresponding failure surfaces were extensively 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 each joint?s respective failure loads to identify critical parameters, which would later be used to develop a Critical Parameter Method for evaluating joint performance.
Once these parameters were consolidated, they were validated against a unique set of joints. The critical parameter approach was able to predict joint strength with an average error of 26% compared experimental strength. Traditional FE criterions presented an average error of 61% compared to experimental strength. After further consolidation, joint strength prediction reduced to within 3% of experimental strength using the Critical Parameter Method, representing a substantial improvement in predictive capabilities.
In recent years extremely small satellites have been developed in response to trends in the space industry to achieve more for less cost. Extremely lightweight and efficiently packaged deployable structures are essential for achieving large-scale applications including communication antennas, solar arrays, and in recent years, deorbiting drag-sails.
This thesis is motivated for developing novel deployable helical antennas for space-based maritime surveillance. The helical antenna technology provides packaging efficiency and radio frequency characteristics superior to the latest efforts of international research groups. To achieve this, the research presented focuses on developing the proven bistable composite slit tube (BCST) deployable technology. These are open-section tubular structures which can be deployed and rolled up into a compact coil, analogous to a tape measure, but do not require constraint to remain stowed. This behaviour is referred to as bistability and enables lightweight and relatively simple deployable structures for spacecraft applications.
New forms of BCST are modelled through the introduction of additional curvatures, manufactured and described in this work with two new subcategorisations established: toroidal and helical. Toroidal BCSTs are doubly curved with both principal curvatures initially non-zero in the deployed stress-free state. Helical BCSTs are doubly curved and twisted out-of-plane. Investigations into the effects of geometrical parameters and laminated composite material properties on the bistable coils of both types are presented. The results provide an understanding of bistable behaviour in new forms of BCST previously neglected in the literature, which is almost exclusively focused on straight forms. As a consequence of this research, new deployable structure technologies are envisaged in the areas of compact terrestrial shelters and small satellite communications.
Taylor Benjamin, Underwood Craig, Viquerat Andrew, Fellowes Simon, Duke Richard, Stewart Brian, Aglietti Guglielmo, Bridges Christopher, Schenk M, Massimiani Chiara, Masutti D, Denis A (2018) Flight Results of the InflateSail Spacecraft and Future Applications of Drag Sails, 32nd Annual AIAA/USU Conference on Small Satellites pp. 1-12
AIAA/Utah State University
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.
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.
Underwood Craig, Viquerat Andrew, Schenk Mark, Taylor Ben, Massimiani Chiara, Duke Richard, Stewart Brian, Fellowes Simon, Bridges Chris, Aglietti Guglielmo, Sanders Berry, Masutti Davide, Denis Amandine (2019) InflateSail De-Orbit Flight Demonstration Results and Follow-On Drag-Sail Applications, Acta Astronautica - Proceedings of the 69th International Astronautical Congress (IAC)
International Astronautical Federation (IAF)
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
sail. InflateSail's primary mission was to demonstrate the effectiveness of using a drag sail in Low Earth Orbit (LEO)
to dramatically increase the rate at which satellites lose altitude and re-enter the Earth's atmosphere and it was one of
31 satellites that were launched simultaneously on the PSLV (polar satellite launch vehicle) C-38 from Sriharikota,
India on 23rd June 2017 into a 505km, 97.44o
Shortly after safe deployment in orbit, InflateSail automatically activated its payload. Firstly, it inflated its metrelong
metal-polymer laminate tubular mast, and then activated a stepper motor to extend four lightweight bi-stable
rigid composite (BRC) booms from the end of the mast, so as to draw out the 3.1m x 3.1m square, 12mm thick
polyethylene naphthalate (PEN) drag-sail. As intended, the satellite immediately began to lose altitude, causing it to
re-enter the atmosphere just 72 days later ? thus successfully demonstrating for the first time the de-orbiting of a
spacecraft using European inflatable and drag-sail technologies.
The InflateSail project was funded by two European Commission Framework Program Seven (FP7) projects:
DEPLOYTECH and QB50. DEPLOYTECH had eight European partners including DLR, Airbus France, RolaTube,
Cambridge University, and was assisted by NASA Marshall Space Flight Center. DEPLOYTECH?s objectives were
to advance the technological capabilities of three different space deployable technologies by qualifying their
concepts for space use. QB50 was a programme, led by VKI, for launching a network of 50 CubeSats built mainly by
university teams all over the world to perform first-class science in the largely unexplored lower thermosphere.
The boom/drag-sail technology developed by SSC will next be used on a third FP7 Project: RemoveDebris,
launched in 2018, which will demonstrate the capturing and de-orbiting of artificial space debris targets using a net
and harpoon system. This paper describes the results of the InflateSail mission, including the observed effects of
atmospheric density and solar activity on its trajectory and body dynamics. It also describes the application of the
technology to RemoveDebris and its potential as a commercial de-orbiting add-on package for future space missions.
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.
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.
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.
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.
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.
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.