Maintaining missions in proximity of small bodies involves extensive orbit determination and ground station time due to the current ground-in-the-loop approach. The prospect of having multiple concurrent missions around different targets requires the development of concepts and capabilities for autonomous proximity operations. Developments in on-board navigation by landmark maps paved the way for autonomous guidance at asteroids. The missing elements for achieving this goal are gravity models, simple enough to be easily used by the spacecraft to steer itself around the asteroid, and guidance laws that rely on such inherently simple models. In this research, we identify a class of models that can represent some characteristics of the dynamical environment around small bodies with sufficient accuracy to enable autonomous guidance. We found that sets of three point masses are suitable to represent the rotational equilibrium points generated by the balance of gravity and centrifugal acceleration in the body-fixed frame. The equilibrium point at the lowest Jacobi energy can be viewed as the energy-gateway to the surface. Information of the location and energy of this point can then be used by a control law to comply with a condition of stability against impact for orbital trajectories. In this thesis, we show an optimisation process for the derivation of three-point mass models from higher order ones and compare the profile of the Zero-Velocity curves between the two models. We define an autonomous guidance law for achieving body fixed hovering in proximity of the asteroid while ensuring that no impact will occur with the small body during the manoeuvre. Finally, we discuss the performance of this approach by comparing it with another autonomous guidance law present in literature and we suggest possible future developments.
Taylor Benjamin, Duke Richard, Stewart Brian, Massimiani Chiara, Djamane F, Bridges Christopher, Aglietti Guglielmo, Lassakeur Abdelmadjid, Amine Ouisb M, Cherif Ladouze M, Meftah K, Underwood Craig, Chikouche A, Hamed DEB (2017)AlSat-Nano: Knowledge Transfer to Operational Partnership, In: 68th International Astronautical Congress Proceedings
International Astronautical Federation
The AlSat-Nano mission is a joint endeavour by the UK and Algeria to build and operate a 3U CubeSat. The project was designed to provide training to Algerian students, making use of UK engineering and experience. The CubeSat was designed and built by the Surrey Space Centre (SSC) of the University of Surrey and hosts three UK payloads with operations run by the Algerian Space Agency (ASAL). The educational and CubeSat development were funded by the UK Space Agency (UKSA), whilst the UK payloads were self-funded. Launch and operations are funded by ASAL. This paper illustrates the development of the programme, the engineering of the satellite and the development of collaborative operations between the SSC and ASAL.
The term microvibrations generally refers to accelerations in the order of micro-gs and which manifest in a bandwidth from a few Hz up to say 500-1000 Hz. The need to accurately characterise this small disturbances acting on-board modern satellites, thus allowing the design of dedicated minimisation and control systems, is nowadays a major concern for the success of some space missions. The main issues related to microvibrations are the feasibility to analytically describe the microvibration sources using a series of analysis tools and test experiments and the prediction of how the dynamics of the microvibration sources couple with those of the satellite structure. In this thesis, a methodology to facilitate the modelling of these phenomena is described. Two aspects are investigated: the characterisation of the microvibration sources with a semi-empirical procedure which allows derivation of the dynamic mass properties of the source, also including the gyroscopic effect, with a significantly simpler test configuration and lower computational effort compared to traditional approaches; and the modelling of the coupled dynamics when the source is mounted on a representative supporting structure of a spacecraft, including the passive and active effects of the source, which allows prediction of the structure response at any location. The methodology has been defined conducting an extensive study, both experimental and numerical, on a reaction wheel assembly, as this is usually identified as the main contributory factor among all microvibration sources. The contributions to the state-of-the-art made during this work include: i) the development of a cantilever configured reaction wheel analytical model able to reproduce all the configurations in which the mechanism may operate and inclusive of the gyroscopic effect; ii) the reformulation of the coupling theory which allows retrieving the dynamic mass of a microvibration source over a wide range of frequencies and speeds, by means of the experimental data obtained from measurements of the forces generated when the source is rigidly secured on a dynamometric platform and measurements of the accelerations at the source mounting interface in a freefree suspended boundary condition; iii) a practical example of coupling between a reaction wheel and a honeycomb structural panel, where the coupled loads and the panel response have been estimated using the mathematical model and compared with test results, obtained during the physical microvibration testing of the structural panel, showing a good level of agreement when the gyroscopic effect is also taken into account.
Taylor Benjamin, Underwood Craig, Viquerat Andrew, Fellowes Simon, Denis A, Bridges Christopher, Duke Richard, Stewart Brian, Aglietti Guglielmo, Schenk M, Massimiani Chiara, Masutti D (2018)Flight Results of the InflateSail Spacecraft and Future Applications of Drag Sails, In: 32nd Annual AIAA/USU Conference on Small Satellitespp. 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.
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 (2018)InflateSail De-Orbit Flight Demonstration Results and Follow-On Drag-Sail Applications, In: 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 deployable drag sail. InflateSail's primary mission was to demonstrate the effectiveness of using a drag sail in Low Earth Orbit (LEO) to dramatically increase the rate at which satellites lose altitude and re-enter the Earth's atmosphere and it was one of 31 satellites that were launched simultaneously on the PSLV (polar satellite launch vehicle) C-38 from Sriharikota, India on 23rd June 2017 into a 505km, 97.44o Sun-synchronous orbit. Shortly after safe deployment in orbit, InflateSail automatically activated its payload. Firstly, it inflated its metrelong metal-polymer laminate tubular mast, and then activated a stepper motor to extend four lightweight bi-stable rigid composite (BRC) booms from the end of the mast, so as to draw out the 3.1m x 3.1m square, 12m thick polyethylene naphthalate (PEN) drag-sail. As intended, the satellite immediately began to lose altitude, causing it to re-enter the atmosphere just 72 days later – thus successfully demonstrating for the first time the de-orbiting of a spacecraft using European inflatable and drag-sail technologies. The InflateSail project was funded by two European Commission Framework Program Seven (FP7) projects: DEPLOYTECH and QB50. DEPLOYTECH had eight European partners including DLR, Airbus France, RolaTube, Cambridge University, and was assisted by NASA Marshall Space Flight Center. DEPLOYTECH’s objectives were to advance the technological capabilities of three different space deployable technologies by qualifying their concepts for space use. QB50 was a programme, led by VKI, for launching a network of 50 CubeSats built mainly by university teams all over the world to perform first-class science in the largely unexplored lower thermosphere. The boom/drag-sail technology developed by SSC will next be used on a third FP7 Project: RemoveDebris, launched in 2018, which will demonstrate the capturing and de-orbiting of artificial space debris targets using a net and harpoon system. This paper describes the results of the InflateSail mission, including the observed effects of atmospheric density and solar activity on its trajectory and body dynamics. It also describes the application of the technology to RemoveDebris and its potential as a commercial de-orbiting add-on package for future space missions.
Eckersley S., Saunders C., Gooding D., Sweeting M., Whiting C., Ferris M., Friend J., Forward L., Aglietti G., Nanjangud A., Blacker P., Underwood C., Bridges C., Bianco P. (2018)In-Orbit Assembly of Large Spacecraft Using Small Spacecraft and Innovative Technologies, In: Proceedings of the 69th International Astronautical Congress (IAC)
International Astronautical Federation (IAF)
The size of any single spacecraft is ultimately limited by the volume and mass constraints of currently available launchers, even if elaborate deployment techniques are employed. Costs of a single large spacecraft may also be unfeasible for some applications such as space telescopes, due to the increasing cost and complexity of very large monolithic components such as polished mirrors.
The capability to assemble in-orbit will be required to address missions with large infrastructures or large instruments/apertures for the purposes of increased resolution or sensitivity. This can be achieved by launching multiple smaller spacecraft elements with innovative technologies to assemble (or self-assemble) once in space and build a larger much fractionated spacecraft than the individual modules launched.
Up until now, in-orbit assembly has been restricted to the domain of very large and expensive missions such as space stations. However, we are now entering into a new and exciting era of space exploitation, where new mission applications/markets are on the horizon which will require the ability to assemble large spacecraft in orbit. These missions will need to be commercially viable and use both innovative technologies and small/micro satellite approaches, in order to be commercially successful, whilst still being safety compliant. This will enable organisations such as SSTL, to compete in an area previously exclusive to large commercial players. However, inorbit assembly brings its own challenges in terms of guidance, navigation and control, robotics, sensors, docking mechanisms, system control, data handling, optical alignment and stability, lighting, as well as many other elements including non-technical issues such as regulatory and safety constraints. Nevertheless, small satellites can also be used to demonstrate and de-risk these technologies.
In line with these future mission trends and challenges, and to prepare for future commercial mission demands, SSTL has recently been making strides towards developing its overall capability in “in-orbit assembly in space” using small satellites and low-cost commercial approaches. This includes studies and collaborations with Surrey Space Centre (SSC) to investigate the three main potential approaches for in-orbit assembly, i.e. deployable structures, robotic assembly and modular rendezvous and docking. Furthermore, SSTL is currently developing an innovative small ~20kg nanosatellite (the “Target”) as part of the ELSA-d mission which will include various rendezvous and docking demonstrations. This paper provides an overview and latest results/status of all these exciting recent in-orbit assembly related activities.
The launch phase is the most demanding mechanical environment typical satellites experience. In order to verify that a payload or piece of equipment will survive the expected loads experienced during launch, it is subject to prescribed vibration environments. However, current vibration testing methods tend to overtest. This means the harshest environment a satellite and its equipment must survive is the testing, not the launch. Consequentially, design process compromises are made, moving the focus from surviving the launch to surviving the testing. Vibration testing involves shaking the test article in each of the three standard directions (X, Y and Z) according to the provided testing specifications. These specifications are based of the single launch environment which is split into three for practical reasons, but which leads to overtesting. One of the causes for equipment overtesting is that items are normally tested along its three orthogonal axes (i.e. X, Y and Z). However, the body axes of the equipment are not always in line with the structure it is attached to. Even if the body axes do align, the dynamics of the coupled system mean any vibration at the base of the larger structure is unlikely to be acting all on the same axis (or axes) at the interface between the satellite and equipment. Another key difference between the testing environment and launch environment is the direction of the vibrations. The launch vibration environment is a single 3D environment, while testing is usually comprised of three single axis vibrations tests. This thesis presents two alternative testing methods that separately, or together, can create a test campaign which better matches the environment the piece of equipment would see during launch. The first method, the Angle Optimisation Method, looks at testing the piece of equipment is mounted at an offset angle on to the shaker rather than the traditional three orthogonal mounting directions. The method optimises the testing angle for the piece of equipment such that testing responses are closer to those seen when the equipment is attached to the higher level assembly. This method focuses on covering the maximum Root Mean Square (RMS) values for each quantity (e.g. sum of interface forces, and acceleration at centre of mass) obtained from the coupled system tests - resulting in a test campaign of one to three separate tests, each with altered input directions. This results in RMS values much closer to the desired higher level testing values than the traditional testing. The second method, the Dual Input Method, looks at adding a secondary smaller vibration source at a specific location on the test item. The method finds the best location to attach the second vibration source that produces a more representative testing of the piece of equipment when compared to the higher level testing. It also determines what the input should be at this specific point. This method looks at improving the correlation of the Operational Deflecting Shapes (ODS) of the equipment when tested in isolation and when attached to the higher level assembly. Response Vector Assurance Criterion (RVAC) is used for the correlation of the ODS. Two case studies were undertaken to demonstrate the benefits of these methods. The first was a computational case study that both methods were applied to. In this case study the Angle Optimisation method was able to reduce the amount of over testing by up to 70% compared to the traditional testing method. While the Dual Input method was able to improve the correlation between the equipment and coupled system responses by nearly 50%. The second case study was an experimental application of the Angle Optimisation Method. This case study successfully showed that it was possible to implement this method as a physical test. A custom angled interface plate was manufactured to the specifications determined by the Angle Optimisation method. In addition to showing the successful implementation of this method, the over testing was reduced by roughly 50% when compared to the traditional method.
Dynamic variability in nominally identical structures is an issue widely studied in structural dynamics community. Small uncertainties and manufacturing tolerances can significantly affect the dynamic behaviour of spacecraft payload. The aim of this paper is the investigation of such dynamic variability generated from both mechanical actuators and spacecraft structure itself. Both aspects will be tackled in this paper, suggesting two distinct approaches able to take into account these variations. First, vibration sources will be analysed by using real space mechanisms data (i.e. reaction wheels) and applying the proposed methodology in different cases. Finally, spacecraft structure variability will be addressed by looking at the dynamic behaviour of the satellite from a subsystem point of view. Such dynamic variability will then be assessed by the definition of specific margin of uncertainty.
Disturbances generated by reaction wheels on board the spacecraft are among the most substantial . Hence they play a crucial role when microvibration budget has to be assessed. This paper aims at characterising the effects of RW on the structure by focusing on the format of the disturbance input matrix of these components. In particular the case of single and multiple wheel accounted for. In the first one the responses are evaluated at some specific locations of the reaction wheel where their disturbance is amplified, i.e. harmonics. In the second case a more realistic scenario is considered with several wheels to be characterised and the effects of neglecting some terms of the disturbance input matrix are discussed. Finally a sensitivity analysis is carried out to quantify in which extent changes in the input matrix can alter the response. A preliminary methodology is then suggested to characterise a large num ber of wheels.
Driven by the increasingly stringent stability requirement of some modern payloads (e.g. the new generations of optical instruments) the issue of accurate spacecraft micro-vibration modelling has grown of importance. This article focuses on the dynamic coupling between a source of micro-vibration (e.g. reaction wheel) and a structure, taking into account the uncertainties related to both parts. In this context, an alternative to the Monte Carlo Simulation for complex structures has been developed, consisting in sub-structural approach to perturb the natural frequencies of specific subsystem reduced with the Craig-Bampton method. In order to prove the validity of the method and its application to the theory of the coupling, benchmark examples and practical applications will be described. © 2013 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
In this paper, the mathematical framework for a computationally efficient stochastic finite element method (FEM) is outlined. It is devised for a range of applications in structural dynamics, where uncertainties need to be reliably dealt with in the context of reduced model formulations. It allows random mass and stiffness matrices to be robustly generated at the subsystem level in component mode synthesis (CMS) applications. The technique is validated for the particularly challenging case of mid-frequency FEM-FEM vibroacoustic analysis of a spacecraft structure. Results are compared against both test data and full parametric Monte-Carlo simulation. Finally, the method’s applicability to coupled vibroacoustic problems utilising hierarchical matrix boundary element method (BEM) acoustic formulations is evaluated.
Driven by the increasingly stringent stability requirement of some modern payloads (e.g. the new generations of optical instruments) the issue of accurate spacecraft micro-vibration modeling has grown increasingly important. In this context micro-vibrations are low level mechanical disturbances occurring at frequencies from a few Hertz up to 1000 Hz. As the frequency content of these phenomena extends beyond the first few modal frequencies, FEA predictions become less accurate and alternative methods have to be considered. Other modeling and analysis techniques have been investigated and applied to vibration problems (Stochastic Finite Element Method (e.g. Monte Carlo Simulation), Statistical Energy Analysis (well-established method for high frequency ranges) and the Hybrid FE-SEA), with the aim of investigating medium and high frequency behavior. This work is part of a project whose aim is to establish appropriate procedures for the modeling and analysis of micro-vibration and validate these procedures against experimental data. All the methods cited above are implemented in this study and compared with experimental results, in order to assess the performance of the various methodologies for micro-vibration problems, covering the whole frequency range up to 1000 Hz. Some comparisons between experimental and computational results are performed using the MAC. Some other analyses, like linearity, reciprocity or effect of the harness are also described. The bench work model that has provided the experimental data is the satellite platform SSTL 300 and this paper outlines these related test campaigns.
Contracted by the European Commission in the frame of the EU’s Seventh Framework Programme for Research (FP7), a wide European consortium has been working since 2013 towards the design of a low cost in-orbit demonstration called RemoveDEBRIS. With a targeted launch date in the second quarter of 2016, the RemoveDEBRIS mission aims at demonstrating key Active Debris Removal (ADR) technologies, including capture means (net and harpoon firing on a distant target), relative navigation techniques (vision-based navigation sensors and associated algorithms), and deorbiting technologies (drag sail deployment after the mission followed by an uncontrolled reentry). In order to achieve these objectives, a micro satellite testbed will be launched into a Low Earth Orbit, where it will deploy its own dedicated targets and CubeSats to complete each demonstration. As part of its System Engineering role, Airbus Defence and Space has been conducting the Mission Analysis studies for this unprecedented mission. This paper will present a description of the RemoveDEBRIS demonstration objectives and scenario and will present in detail some specific mission related analyses and trade-offs that have driven the mission design.
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.
Reliable and efficient vibroacoustic loads prediction is often critical in structural design, yet it remains a challenging task for many applications. Spacecraft structures are characterised by extensive use of composite materials, complex connections between components and various non-trivial geometrical features. Accurate treatment necessitates the construction of highly detailed numerical models, traditionally employing deterministic representations. Simultaneously, the broadband acoustic excitation due to the diffuse sound field experienced during launch requires modelling the fluid domain and solving the resulting elasto-acoustic interaction at = multiple frequencies. To alleviate the computational demand implications for large problem sizes, substructuring and reduction techniques for the structural domain are commonplace, component mode synthesis (CMS) being a framework widely adopted in the aerospace industry. Nevertheless, despite ongoing research, the topic still presents a range of difficulties when a universal, robust method of accounting for model uncertainties is sought. In this study, two CMS based approaches are proposed and evaluated. Firstly, the Craig-Bampton stochastic method (CBSM) is improved via a set of modifications enhancing its efficiency, and subsequently adapted for use in a vibroacoustic setting. Optimal perturbation levels and scope of validity of the technique are established against a probabilistic structural analysis (PSA) simulation for a spacecraft structure. Secondly, a novel stochastic finite element method (FEM) is presented. The underlying mathematical foundation is derived so that uncertainty can naturally be controlled at the subsystem level, in partitions of the corresponding condensed mass and stiffness matrices. This decomposition based approach ensures that realisations of the random matrices have key properties such as positive (semi)definiteness strictly preserved, guaranteeing complete robustness. The method is validated with a spacecraft test case, comparing its predictions against PSA, the improved CBSM and experimental data. A coupling scheme with a hierarchical matrix accelerated boundary element method is formulated, resulting in the construction of a complete stochastic vibroacoustic solver.
Nominal operability of satellites can be significantly affected by low level vibrations in the range of micro-g generated by on board mechanisms. These are referred to as micro-vibrations and can considerably jeopardise the regular functioning of very high precision sensors. Hence, it is essential to study their features in terms of characterisation and analysis, which are greatly affected by structural uncertainties due, among other causes, to the manufacturing and assembly tolerances. In this thesis the variability of micro-vibrations arising from structural uncertainties is targeted, through the implementation of a macroscopic approach that is able to take into account different sources of structural uncertainties. In particular the work done in this doctorate is split into two main areas. First the characterisation of micro-vibration sources is studied; these include all the devices on board the spacecraft that can generate micro-vibrations. A methodology to investigate the effects of manufacturing defects on the dynamics of reaction wheels is described, by focusing on the frequency domain representation of their disturbance. Through this approach, it is possible to define a group of nominally-identical devices by means of a single disturbance input matrix. Such achievement can be beneficial in space applications as it allows an easier assessment of micro-vibrations on the spacecraft. Second, the issue of structural uncertainty is addressed in terms of transmission path from the source to the receiver. The focus here is shifted towards the development of an analysis methodology that can target the issue of structural uncertainties by satisfying the stringent computational requirements for aerospace applications. In particular, the main achievement obtained in this thesis stays in the development of an uncertainty quantification methodology that can be used to provide an estimation of perturbation parameters used in the Craig-Bampton Stochastic Method.
Accurate resolution imaging from Satellites involves large amounts of data that has to be stored on board the spacecraft computer. The data can be stored on Hard Disc Drives. However survival to the mechanical environment existing during the spacecraft launch and to the space environment during satellite operations are two major challenges in the use of HDDs for spacecraft applications. This paper describes the process that generated the design of an enclosure that has allowed conventional Personal Computer's HDDs to be used on board the SSTL spacecraft BEIJING-1. The design philosophy is discussed and the extensive test campaigns that supported the selection of a suitable HDD are described. The focus of the work was the design and implementation of a suspension system to reduce to acceptable levels the random vibration environment experienced by the HDD. The tests carried out on the suspension system showed that this was able to reduce by approximately 50% the rms acceleration experienced by the units. Thus allowing their use on the spacecraft. The spacecraft was launched in October 2005, and to date the HDD units are operating correctly.
The paper addresses the problem of actively attenuating a particular class of vibrations, known as microvibrations, which arise, for example, in panels used on satellites. A control scheme that incorporates feedback action is developed which operates at a set of dominant frequencies in a disturbance spectrum, where the control path model is estimated online. Relative to earlier published techniques, a new feature of the presented controller is the use of the inverse Hessian to improve adaptation speed. The control scheme also incorporates a frequency estimation technique to determine the relevant disturbance frequencies with higher precision than the standard fast Fourier transform (FFT). The control scheme is implemented on an experimental test-bed and the total achieved attenuation, as measured from the experiments, is 26 dB. The low computational demand of the control scheme allows for single chip controller implementation, a feature which is particularly attractive for potential applications areas, such as small satellites, where there are critical overall weight restrictions to be satisfied while delivering high quality overall performance. © IMechE 2005.
The issue of model reduction is one that must often be overcome in order to perform the necessary checks as part of the spacecraft Finite Element Model (FEM) validation process. This work compares different reduction methods; specifically the popular and long-standing Guyan method, and the potentially more accurate System Equivalent Reduction Expansion Process (SEREP). The influence of sensor set location on the quality of the reduced model has also been considered, and the commonly applied methods to maximize kinetic energy and effective independence have been applied. These investigations have taken the form of studies involving two large, unique, scientific spacecraft. The computational results are compared with experimental results that are also detailed in the paper. The findings highlight the potential issues with the accuracy of a Guyan reduced model in replicating the full system dynamics, even with a reasonably large sensor set. It is shown that this can be improved slightly in some circumstances through implementation of sensor set placement optimization techniques. The SEREP method is shown to have the benefit of being more accurate at replicating the full system behavior than the more traditional Guyan method, while also producing higher diagonal values in cross-orthogonality comparisons between FEM and test.
Aglietti G, Schaffner J, Ward C, Curiel ADS, Sweeting M (2005)Surrey small satellite transfer vehicle, In: International Astronautical Federation - 56th International Astronautical Congress 20054pp. 2169-2178
The concept for a Small Satellite Transfer Vehicle is presented. This vehicle is targeted at the deployment of complex groups, constellations or formations of nano- or picosats into an orbit as a secondary payload. The concept provides the capability to deploy this secondary payload into a significantly different orbit from the primary payload on the launch, and has the capability to provide additional services such as spacecraft dispersion along an orbit. It aims to minimise the overheads associated with multiple small spacecraft that are manifested on a single launcher, and offers a number of benefits over the more typical ad-hoc launcher accommodation. The Cubesat PPOD accommodation concept is employed as it provides a suitable open standard that is well used, tested and documented. A mission scenario is introduced and analysed and the system design is described.
Dynamic models are widely used in many branches of science and engineering, and it has been argued that many of the shortfalls with these models are due to the fact that the physics of joint dynamics are not fully understood. This makes the phenomenon very hard to model theoretically from first principles. Experimental analyses are therefore widely used to underpin any work in this area. This study aims to build on the previous experimental work based on simple beam joints and analyzes the damping trends for metal panels. This incorporates torsional effects into the system creating more complex displacements of the joint. Five panel configurations are investigated using an experimental approach that minimizes all external influences on the dynamics of the panels. Each mode loss factor is determined from these experimental tests and compared with the most established theoretical model in the field. The corresponding joint displacements and decay trends are also analyzed, producing indications as to the likely dominant source of damping, suggesting that the mode shapes can be categorized based on their displacement and dominant damping source.
This paper examines several available analytic and experimental methods to determine the orthotropic material properties of honeycomb. Fifteen published sets of simple equations for the material properties were reviewed and their values calculated for a specific honeycomb aluminum core. The same core was tested with ASTM standard methods and the agreement between the theoretical material properties and the experimental results was considered. To reduce the time and cost for the experimental determination, a simple technique for measuring the main dynamic material properties of honeycomb is introduced. A good agreement was found between the major theoretical out-of-plane material properties of honeycomb, the experimental ASTM methods, and the presented dynamic approach. © ASCE.
A main aim of spacecraft design is cost reduction that can either be achieved by reducing the cost of the spacecraft or by lowering the cost to launch. One proposed technology to reduce the mass and therefore lower the launch costs is the multifunctional power structure concept that incorporates the secondary spacecraft power supply into the load carrying structure. Here a short introduction of the dynamic analysis and optimization of such a structure is presented. The manufacture and testing of a multifunctional power panel is discussed in detail and its dynamic response is compared to a conventional honeycomb panel. The multifunctional design successfully combined the structural and power storage functions. It provided a similar dynamic response to conventional spacecraft structures and improved the energy density.
© 1999 EUCA.The suppression of microvibrations (low amplitude vibrations with frequencies in the range 1 to 1000 Hz) is becoming increasingly important in spacecraft and other applications and can only be achieved (in most cases) by active feedback control schemes. This paper describes a Lagrange-Rayleigh-Ritz method which has been used to develop a state space description of the generic case of a vibrating panel with piezo-electric patches as actuators and sensors, disturbances, and a payload. The resulting models are used here to design H∞ based active feedback control schemes for disturbance attenuation.
For next generation microsatellites and nanosatellites, new design approaches will be required to significantly increase their payload to mass fraction. One proposed technology is the multifunctional design concept that incorporates spacecraft subsystems into the load carrying structure. The focus of the research is the multifunctional power structure which replaces conventional battery systems in a spacecraft. An analytical and finite element analysis of ten multifunctional sandwich structures is presented. The out-of-plane material properties are discussed and a parameter optimization of the ten sandwich panels is carried out to optimize their frequency to density ratio. The best configuration for an optimized multifunctional power structure is then identified from the analytical and finite element investigation. The optimized design provides a similar predicted dynamic response as a conventional honeycomb sandwich panel, and can be considered a serious alternative for future spacecraft. Copyright © 2006 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
New design approaches will be required to increase the payload to mass fraction for future satellite generations. The multifunctional design concept, where spacecraft subsystems are integrated into the load bearing structure of the satellite, is one considered technology. This paper describes the design, analysis and manufacture of a particular multifunctional power structure with a special focus on its dynamic response. An analytical and a finite element analysis of ten proposed multifunctional power structures, based on a sandwich panel configuration, are presented. The theoretical out-of-plane material properties for the investigated designs are derived with the help of the virtual displacement method. These theoretical properties are compared to finite element models and subsequently used in a parameter optimisation of the dynamic response of the ten introduced sandwich panels. The optimisation allows the identification of the most favourable multifunctional power structure. The experimental dynamic response of a manufactured multifunctional power panel is presented and compared to a conventional honeycomb panel for a successful evaluation of the introduced multifunctional approach. The results of this work show the ability of the presented multifunctional design to successfully combine the structural and power storage functions which makes the multifunctional power structure an excellent design approach for future space missions.
This paper investigates the feasibility and economical advantages that could be offered by a new facility for the production of solar energy. The basic concept is to exploit a high altitude aerostatic platform to support Photovoltaic (PV) modules to substantially increase their output by virtue of the significantly enhanced solar radiation at the operating altitude of the aerostat. The electric energy is then transmitted to the ground using the aerostat mooring cable. The technical feasibility of the concept is demonstrated by using standard equations and realistic values for the relevant engineering parameters that describe the technical properties of the materials and subsystems. There are, nevertheless, issues to be addressed to improve the performance; however, none of these issues is deemed to negate the technical feasibility of this concept. A preliminary cost model is illustrated and using unit cost data for the various materials and subsystems it is shown that it is possible to identify a specific size that minimizes the cost of energy produced. This cost could be considerably lower than what can be achieved by solar panels based on the ground in the UK. Copyright © 2008 John Wiley & Sons, Ltd.
Conventional spacecraft subsystems are designed and manufactured separately, and are integrated only during the final stages of satellite development. This requires containers for the subsystems' hardware, mechanical interfaces, panels, frames, bulky wire harnesses, etc., which add considerable mass and volume. As all subsystems are generally secured to the structure, the multifunctional structure approach aims at merging these elements into the structure, so that the structure also carries out some of the typical functions of the subsystems (e.g. electrical energy storage). The main advantages are as follows: (i) removal of the bolted mechanical interfaces and most of the subsystems' containers; (ii) reduction of the satellite structure mass, as the strength of the parts of the subsystem imbedded into the structure are exploited, and substitute purely structural parts; (iii) reduction of the overall satellite volume, as elements such as battery packs or electronic harnesses can be built into the structure's volume. There are still issues that need to be addressed to allow a wider utilization of multifunctional structures. However, the development of concurrent engineering approaches, to carry out an integrated design of the spacecraft, together with advances in the subsystems' disciplines, will help to promote the further diffusion of multifunctional structures. © 2007 SAGE Pulications.
Coupled Loads Analyses (CLAs), using finite element models (FEMs) of the spacecraft and launch vehicle to simulate critical flight events, are performed in order to determine the dynamic loadings that will be experienced by spacecraft during launch. A validation process is carried out on the spacecraft FEM beforehand to ensure that the dynamics of the analytical model sufficiently represent the behavior of the physical hardware. One aspect of concern is the containment of the FEM correlation and update effort to focus on the vibration modes which are most likely to be excited under test and CLA conditions. This study therefore provides new insight into the prioritization of spacecraft FEM modes for correlation to base-shake vibration test data. The work involved example application to large, unique, scientific spacecraft, with modern FEMs comprising over a million degrees of freedom. This comprehensive investigation explores: the modes inherently important to the spacecraft structures, irrespective of excitation; the particular ‘critical modes’ which produce peak responses to CLA level excitation; an assessment of several traditional target mode selection methods in terms of ability to predict these ‘critical modes’; and an indication of the level of correlation these FEM modes achieve compared to corresponding test data. Findings indicate that, although the traditional methods of target mode selection have merit and are able to identify many of the modes of significance to the spacecraft, there are ‘critical modes’ which may be missed by conventional application of these methods. The use of different thresholds to select potential target modes from these parameters would enable identification of many of these missed modes. Ultimately, some consideration of the expected excitations is required to predict all modes likely to contribute to the response of the spacecraft in operation.
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.
Tape springs, defined as thin metallic strips with an initially curved cross section, are an attractive structural solution and hinge mechanism for small satellite deployable structures due to their low mass, low cost, and general simplicity. When mounted at skewed angles to the hinge line, the tapes can be subjected to complex folds involving both bending and twisting of the tape. These folds have been experimentally investigated and theories have been developed to model the resulting opening moment. However, the opening moments of these theories are not equivalent to the opening moment about the hinge line, which is the parameter required in satellite deployment applications. This paper derives a method to determine the hinge moment from the previous theories and compares the theoretical predictions with experimental and finite element results. It uses this model to investigate the predicted hinge moment trends for full deployments of 180°. The model is then applied to a practical spacecraft hinge application. © 2007 ASCE.
Multifunctional spacecraft power structures are an incorporation of energy storage and generation into structures on a spacecraft. For the mass and volume saving benefits to be realised, the technology must be shown to be viable throughout the spacecraft's lifetime. Firstly, commercially available batteries where built into a structural panel and tested to determine the battery's capability to withstand the manufacturing cycle and the effect upon the mechanical characteristics of the panel. Secondly, a mathematical model was created to determine the temperatures a battery would experience in various earth orbits. It was found that spacecraft in most low earth orbits will require thermal control and that the addition of a phase change material is a feasible control solution. Copyright © 2008 by ASME.
In the current world of engineering, structural vibration problems continue impact the design and construction of a wide range of products. Amid the parameters that determine the dynamic behaviour of a structure the one that takes into account the dissipation of energy resulting in the decay of the vibration is the least understood and the most difficult to quantify . The estimation of damping factors is of interest in most branches of engineering sciences. In the field of aircraft structures the damping directly affects the fatigue life, a parameter which is applied conservatively due to the inherent complexity in modelling the damping of built up structures and the potentially catastrophic consequences of a fatigue failure. One of the most important problems is the limited knowledge of how joints affect the damping of the complete structure. This work therefore addresses this issue and focuses on the damping of joints in metal plates as part of a larger project to investigate the damping of built up structures. Various plate configurations are experimentally investigated using two different approaches. The results from the configurations are compared and discussed along with the advantages and disadvantages of each experimental approach. This enables a link to be identified between the damping magnitudes and the mode shapes and joint stiffnesses.
A multi-functional structure saves mass from a spacecraft by incorporating other functional subsystems into the structure. By using the structural properties of a non-structural element, inert structure may be eliminated, and the requirement to allot internal volume to the subsystem in question is removed. The current paper describes a multi-functional structure based on the secondary power system. By using commercially available plastic lithium-ion cells to form the core of a sandwich panel, inert mass is eliminated from both the structure and from the battery enclosure. The feasibility of the proposed multi-functional structure is demonstrated though vibration testing on a single cell, and the successful manufacture of a test panel. The work goes on to quantify the potential mass savings that may be achieved by using a multi-functional structure of this type. By varying a set of spacecraft attributes, the study identifies that small spacecraft with high power requirements have the potential to gain the most benefit from using a multi-functional structure of this type. © IMechE 2008.
Mathematical models of structural dynamics are widely used and applied in many branches of science and engineering, and it has been argued that many of the shortfalls with these models are due to the fact that the physics of joint dynamics are not properly represented. Experimental analyses are, therefore, widely used to underpin any work in this area. The most renowned model for predicting the damping resulting from air pumping is based on a significant quantity of experimental data and was generally developed and applied to high frequency vibrations of jointed or stiffened panels. This publication applies this model to low frequency panel vibrations by assessing the accuracy of the model for these systems. It is concluded that the theoretical model for high stiffness joints, although generally over approximating the damping magnitude, gives a good conservative estimate of the increase in damping due to air pumping for low frequency vibrations. © IMechE 2008.
The technology of multifunctional structures, when applied to spacecraft batteries, allows mass and volume to be saved through the removal of parasitic components of the battery. Electrochemical cells, in this case commercially available cells of the plastic lithium-ion type, may be incorporated directly into structural sandwich panels, removing the need for a secondary structure (i.e. the battery enclosure and its interface) to mount the cells. Placing the cells within the structure also removes them from the bus, reducing its volume and thus further reducing overall mass. Notwithstanding the relatively low mechanical properties exhibited by the batteries that have been investigated during our work, this paper shall demonstrate how, by optimizing the layout of the cells within a panel, adequate structural performance may be maintained. Copyright © 2008 by the American Institute of Aeronautics and Astronautics, Inc.
Multifunctional structures offer savings in spacecraft mass and volume by combining the functional elements of subsystems with structural components. However, the benefit of these mass and volume savings can be outweighed by the cost of the materials and processes required for their manufacture. This paper presents a compromise, taking advantage of some of the mass and volume savings associated with multifunctional structures, whilst remaining relatively low in price. By taking commercial electrical cells, and integrating them into the structure, the parasitic mass of the power subsystem is eliminated and the structural properties of the cells harnessed to produce a multifunctional powerstructure. In this work, prismatic plastic lithium-ion cells are used as part of the core of a carbon fibre sandwich panel, along with aluminium honeycomb. This makes use of the batteries' structural properties and also removes their volume from the spacecraft bus to an area that would otherwise be filled with inert material. The paper assesses the feasibility of this concept. Firstly, a vibration test has been successfully conducted to prove the ability of such cells to survive the launching environment. Secondly, a panel has been successfully manufactured using these cells as described, without serious damage to the cells' performance.
Multistable composite elements are a convenient approach to realise morphing or shape-adaptable structural systems. This property is particularly important, for example, in the aeronautical industry where morphing structures can be exploited for better aircraft performance and operational versatility. In this paper it is proposed to study the dynamics of a bistable square plate with pinned boundary conditions using a simplified single-degree-of-freedom (SDOF) model. Initially the numerical load-deflection characteristic of the centre of a plate pinned at the four corners is obtained from a finite element model (FEM) using ABAQUS. This curve was then adjusted to account for the change in material properties due to exposure to moisture and others ambient variables. The dynamic response of the plate was simulated by solving the equation of motion numerically. A test rig was designed and built. The bistable plate was hinged to two beams which are rigid in the vertical direction but allow for horizontal displacement. As predicted, the measurements show that the response of the plate to an harmonic excitation of the base is periodic for a low amplitude of excitation. For large excitation, a snap-through (passage from one stable state to the other) takes place. Ideally, if the boundary condition were symmetric, the chaotic passage from one stable state to another could be observed (as for a system with a double-well potential). However, in practice ideal boundary conditions cannot be achieved and the load-deflection characteristic is not symmetric. As a result, the frequency and the amplitude at which the snap occurs depends on the initial stable configuration. © 2008 by the Katholieke Universiteit Leuven Department of Mechanical Engineering All rights reserved.
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.
This paper illustrates a procedure to calculate the response of a tethered spherical aerostat to gusts, including the effect of structural nonlinearity and accounting for some of the fluid - structure interaction between the aerostat and tether line. The procedure developed and presented here is based on a full three-dimensional dynamic finite element model, with aerodynamic loads calculated from the relative velocity between a time-varying input airflow and resulting structural velocities. Exact solutions for the static response and a simplified dynamic model, both developed to validate the results of the procedure illustrated in this paper, are also derived and described in detail. The dynamic responses to gusts are compared with the equivalent steady-state solution to assess the approximations of the static solutions. Particular emphasis is placed on the output rotation of the aerostats to quantify disturbances on the pointing stability produced by gusts.
Forshaw JL, Aglietti GS, Navarathinam N, Kadhem H, Salmon T, Pisseloup A, Joffre E, Chabot T, Retat I, Axthelm R, Barraclough S, Ratcliffe A, Bernal C, Chaumette F, Pollini A, Steyn WH (2016)RemoveDEBRIS: An in-orbit active debris removal demonstration mission, In: Acta Astronautica127pp. 448-463
Since the beginning of the space era, a significant amount of debris has progressively been generated. Most of the objects launched into space are still orbiting the Earth and today these objects represent a threat as the presence of space debris incurs risk of collision and damage to operational satellites. A credible solution has emerged over the recent years: actively removing debris objects by capturing them and disposing of them. This paper provides an update to the mission baseline and concept of operations of the EC FP7 RemoveDEBRIS mission drawing on the expertise of some of Europe's most prominent space institutions in order to demonstrate key active debris remove (ADR) technologies in a low-cost ambitious manner. The mission will consist of a microsatellite platform (chaser) that ejects 2 CubeSats (targets). These targets will assist with a range of strategically important ADR technology demonstrations including net capture, harpoon capture and vision-based navigation using a standard camera and LiDAR. The chaser will also host a drag sail for orbital lifetime reduction. The mission baseline has been revised to take into account feedback from international and national space policy providers in terms of risk and compliance and a suitable launch option is selected. A launch in 2017 is targeted. The RemoveDEBRIS mission aims to be one of the world's first in-orbit demonstrations of key technologies for active debris removal and is a vital prerequisite to achieving the ultimate goal of a cleaner Earth orbital environment.
The modal assurance criterion (MAC) and normalized cross-orthogonality (NCO) check are widely used to assess the correlation between the experimentally determined modes and the finite element model (FEM) predictions of mechanical systems. Here, their effectiveness in the correlation of FEM of two types of multi-physics systems, namely, viscoelastic damped systems and a shunted piezoelectric system are investigated using the dynamic characteristics obtained from a nominal FEM, that are considered as the ‘true’ or experimental characteristics and those obtained from the inaccurate FEMs. The usefulness of the MAC and NCO check in the prediction of the overall loss factor of the viscoelastic damped system, which is an important design tool for such systems, is assessed and it is observed that these correlation methods fail to properly predict the damping characteristics, along with the responses under base excitation. Hence, base force assurance criterion (BFAC) is applied by comparing the ‘true’ dynamic force at the base and inaccurate FEM predicted force such that the criterion can indicate the possible error in the acceleration and loss factor. The effect of temperature as an uncertainty on the MAC and NCO check is also studied using two viscoelastic systems. The usefulness of MAC for the correlation of a second multi-physics FEM that consists of a shunted piezoelectric damped system is also analyzed under harmonic excitation. It has been observed that MAC has limited use in the correlation and hence, a new correlation method – current assurance criterion – based on the electric current is introduced and it is demonstrated that this criterion correlates the dynamic characteristics of the piezoelectric system better than the MAC.
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.
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.
Inflatable technology for space applications is under continual development and advances in high strength fibres and rigidizable materials have pushed the limitations of these structures. This has lead to their application in deploying large-aperture antennas, reflectors and solar sails. However, many significant advantages can be achieved by combining inflatable structures with structural stiffeners such as tape springs. These advantages include control of the deployment path of the structure while it is inflating (a past weakness of inflatable structure designs), an increased stiffness of the structure once deployed and a reduction in the required inflation volume. Such structures have been previously constructed at the Jet Propulsion Laboratory focusing on large scale booms. However, due to the high efficiency of these designs they are also appealing to small satellite systems. This article outlines ongoing research work performed at the University of Southampton into the field of small satellite hybrid inflatable structures. Inflatable booms have been constructed and combined with tape spring reinforcements to create simple hybrid structures. These structures have been subjected to bending tests and compared directly to an equivalent inflatable tube without tape spring reinforcement. This enables the stiffness benefits to be determined with respect to the added mass of the tape springs. The paper presents these results, which leads to an initial performance assessment of these structures.
Due to their high specific strength and high specific stiffness properties the use of honeycomb panels is particularly attractive in spacecraft structures. However, the dynamic loads produced during the launch of a satellite can subject the honeycomb cores of these sandwich structures to numerous stress cycles, potentially leading to early fatigue failures. Knowledge of the fatigue behavior of these honeycomb cores is thus a key requirement when considering their use in spacecraft structural applications. This paper presents the findings of an experimental test campaign carried out to investigate the static and fatigue behaviors of aluminum hexagonal honeycomb cores subject to in-plane shear loads. The investigation involved carrying out both static and fatigue tests using the single block shear test method. These tests were conducted for both of the two principal cell orientations (L and W) and from the results S-N fatigue curves for both cell orientations are presented, confronted and discussed. These results are also discussed in relation to the observed damage and failure modes which have been reported for the statically tested specimens and for the fatigue tested specimens at various stages of fatigue life. Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc.
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 severe mechanical environment that electronic components experience during spacecraft launch necessitates that their failure probability be assessed. One possible approach is to create an accurate model of the Printed Circuit Boards (PCBs) dynamic response; subsequently the failure probability can be determined by comparing the response model with corresponding failure criteria for the electronic components. In principle the response model can be achieved by a very detailed Finite Element (FE) model of the PCB which would include the mass and stiffness of all components present on the PCB. Unfortunately this approach requires an excessive effort; therefore it is rarely pursued by the designer. Past research has shown that assumptions can be made about the mass and stiffness that allow simpler models to be created that still achieve appropriate levels of accuracy. However, the accuracy of these simplified models has not yet been quantified over a range of possible design cases. This paper will quantify how increasing levels of modelling simplifications decrease the accuracy of PCB FE models.
Forshaw JL, Aglietti Guglielmo, Salmon T, Retat I, Roe M, Burgess C, Chabot T, Pisseloup A, Phipps A, Bernal C, Chaumette F, Pollini A, Steyn WH (2016)Review of Final Payload Test Results for the RemoveDebris Active Debris Removal Mission
High-performance space-based optical systems typically require structures that exhibit high levels of dimensional stability over their lifetimes. To better understand the mechanisms for dimensional instability, a novel series of tests were carried out at Rutherford Appleton Laboratory on a breadboard high-stability optical bench structure. Goals of the testing were to assess the relative stability of a pair of reference surfaces and to determine the contributions of various structural elements and joints to dimensional instability. The breadboard was subject to an intensive environmental test campaign that included thermal cycling under vacuum and random vibration testing. Metrology was performed throughout the campaign to assess the dimensional stability response to the various environmental loads. The metrology requirement was challenging, with measurements of micron-level displacement and arcsecond-level tilt over 1-meter distances being necessary in situ during environmental testing. This issue was resolved using a combination of techniques: a contacting coordinate measurement machine, laser interferometry, and optical autocollimation. The greatest levels of instability were produced by random vibration testing, though evidence of a bedding-in process implies that vibratory conditioning could be used to improve stability. Copyright © 2008 by the American Institute of Aeronautics and Astronautics. Inc. All rights reserved.
A structure becomes a multifunctional power structure when in addition to meeting structural requirements it also performs functions associated with the electrical power system. With the structure performing these functions, some separate discreet components may no longer be required. Thus the parasitic structures that support them and the bus volume for these components are no longer required, reducing both mass and volume of the spacecraft. This paper focuses on the inclusion of commercial lithium polymer batteries into a sandwich panel which comprises the structure of a wing mounted solar array. It is shown that the thermal environment in earth orbit is hostile to the batteries. As such, a local thermal control system is required; with its authority targeted at preventing overcooling during eclipse. Phase change materials are proposed as a method to increase the thermal inertia of the structure by exploiting the latent heat. Through numerical simulation, it is shown that phase change materials are a relatively heavy solution. It is demonstrated that as the transition temperature rises, the amount of phase change material increases and that the optical properties of the structure can be altered to reduce the mass of phase change material required to more feasible levels.
Micro-vibration on board a spacecraft is an important issue that affects payloads requiring high pointing accuracy. Although isolators have been extensively studied and implemented to tackle this issue, their application is far from being ideal due to the several drawbacks that they present, such as limited low-frequency attenuation for passive systems or high power consumption and reliability issues for active systems. In the present study, a novel 2-collinear-DoF strut with embedded electromagnetic shunt dampers (EMSD) is modelled, analysed and the concept is physically tested. The combination of high-inductance components and negative-resistance circuits is used in the two shunt circuits to improve the EMSD micro-vibration mitigation and to achieve an overall strut damping performance that is characterised by the elimination of the resonance peaks and a remarkable FRF final decay rate of −80 dB dec–1. The EMSD operates without requiring any control algorithm and can be comfortably integrated on a satellite due to the low power required, the simplified electronics and the small mass. This work demonstrates, both analytically and experimentally, that the proposed strut is capable of producing better isolation performance than other well-established damping solutions over the whole temperature range of interest.
As an intermediate solution between Glaser's satellite solar power (SSP) and ground-based photovoltaic (PV) panels, this paper examines the collection of solar energy using a high-altitude aerostatic platform. A procedure to calculate the irradiance in the medium/high troposphere, based on experimental data, is described. The results show that here a PV system could collect about four to six times the energy collected by a typical U.K.-based ground installation, and between one-third and half of the total energy the same system would collect if supported by a geostationary satellite (SSP). The concept of the aerostat for solar power generation is then briefly described together with the equations that link its main engineering parameters/variables. A preliminary sizing of a facility stationed at 6 km altitude and its costing, based on realistic values of the input engineering parameters, is then presented. © 2009 IEEE.
Test-analysis models are used in the validation of the nite element models of spacecraft structures. Here, a probabilistic approach is used to assess the robustness of a system equivalent reduction expansion process based testanalysis model when experimental and analytical modes contain different levels of inaccuracy. The approach is applied to three spacecraft models, and Monte Carlo simulations were used to determine the sensitivity of the normalized cross-orthogonality check to the system equivalent reduction expansion process reduced matrix. The effect of parameters used in this reduction and the amount of inaccuracies that can be tolerated in the modes before failing the normalized cross-orthogonality check were also determined. The results show that the probability to pass the normalized cross-orthogonality check is highly determined by the number of modes used in the reduction. The relation between capability of the nite element models to predict the frequency-response function and the quality of the model validation determined using normalized cross-orthogonality check is also investigated, and it is observed that the quantities are not always correlated. This study also shows that the sensor locations can be optimally chosen using the system equivalent reduction expansion process reduced mass matrix, and this can increase the probability to pass the normalized cross-orthogonality check.
Microvibrations at frequencies between 1 and 1000 Hz generated by on-board equipment can propagate through a satellite's structure and hence significantly reduce the performance of sensitive payloads. This paper describes a Lagrange-Rayleigh-Ritz method for developing models suitable for the design of active control schemes. Here Loop Transfer Recovery based controller design methods are employed with this modeling strategy.
Aglietti Guglielmo, Wicks A, Barrington-Brown AJ (1999)SIL SGS-2.4 S-band ground station, In: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering213(4)pp. 265-272
The Space Innovations Limited SGS-2.4 ground station is a flexible low-cost system designed specifically for small satellite missions in a low Earth orbit using S- and X-band communications links. The ground station is compatible with ESA and NASA communications standards. These features, as well as its high tracking speeds, make the system widely compatible and capable of supporting many mission types. The paper describes the overall system configuration and covers the mechanical design of the antenna and its drive system. © IMcchE 1999.
Mathematical models (Finite Element Model (FEM)) of spacecraft structures are validated via a correlation process with results from dynamic test campaigns. Typically the Modal Assurance Criterion (MAC) and Normalised Cross Orthogonality (NCO) are used to quantify the quality of the match between models predictions and test results, and appropriate thresholds are set to ensure that a model that passes the test gives sufficiently good predictions of the behaviour of the real structure. However these criteria focus on comparing mode shapes, and the results of this correlation can be misleading. If an excellent correlation is achieved (e.g. MAC very close to unit value), the modal model is an excellent representation of the real structural behaviour; but most often there are models where MAC or NCO is just below the threshold and these are deemed not fit for purpose and requiring some FEM updating to push them above the validation threshold. The main contribution of this work is to show with practical examples of real spacecraft FEMs that a model that passes MAC and NCO checks is not necessarily better at predicting some of the relevant structural responses (e.g. reactions at the base of the craft) than a model that does not pass the checks. The examples show that some of the responses are highly uncorrected with MAC and NCO results and therefore the temptation to extrapolate the results of these validation checks to the "quality" of the model - quality intended in terms of capability of model to predict relevant responses - should be rejected. Using Monte-Carlo simulations, it is also shown that, in the System Equivalent Reduction Expansion Process (SEREP), it is better to include only the target modes in the reduction process. Copyright © (2012) by the International Astronautical Federation.
In this work, three techniques for the mathematical modelling of a piezoelectric actuated thin panel, namely the finite element method, a Lagrange Rayleigh-Ritz method, and a mechanical impedance-based method, are briefly presented. An accurate experimental implementation of a piezoelectric actuated simply-supported panel, whose dynamics have been simulated using the mathematical models, is described in detail. Since the differences between the results produced by the various mathematical models are very small, the accuracy of the experimental set-up is crucial. The results obtained via the numerical simulations are then compared with test results in order to assess the accuracy of the various modelling techniques.
It is well documented that reaction wheels are among the most significant microvibration sources in space applications. These components, despite being nominally identical, can show differences in the generated signals due to manufacturing imperfections in their internal elements, such as ball bearing, internal and external race. In this article a methodology to account for those variations in microvibration predictions is proposed, aiming at generating a disturbance input matrix that encompasses the effects of a family of reaction wheels. With such a tool, it is possible to provide a more accurate microvibration budget at an early stage of the mission, reducing the uncertainty margin usually applied to quantify reaction wheel effects on the structure. As a consequence better designs are produced faster and cheaper. This allows for more flexibility in the mission design and reduces the degree of uncertainties in the predictions. Furthermore, it is shown that the proposed approach is able to characterise the effects of the entire family of wheels by considering only a limited number. The methodology is validated by assessing the microvibration excitation on different structures, including a real space structure with various reaction wheel mounting configurations.
Driven by the increasingly stringent stability requirement of some modern payloads (e.g. the new generations of optical instruments) the issue of accurate spacecraft micro-vibration modelling has grown of importance. This article focuses on the dynamic coupling between a source of micro-vibration (e.g. reaction wheel) and a structure, taking into account the uncertainties related to both parts. In this context, an alternative to the Monte Carlo Simulation for complex structures has been developed, consisting in sub-structural approach to perturb the natural frequencies of specific subsystem reduced with the Craig-Bampton method. In order to prove the validity of the method and its application to the theory of the coupling, benchmark examples and practical applications will be described. © 2012 AIAA.
Reaction Wheel Assemblies (RWAs) are actuators commonly used for satellite attitude control. During their functioning RWAs produce undesired disturbances transmitted through satellite structure and can significantly affect the performance of vibration-sensitive payload. In this paper, microvibrations generated by a RWA in hard-mounted boundary condition are modelled and validated by RWA steady and transient speed test. The two test methods are compared and their differences on the results are discussed. The end-to-end disturbance modelling of a hard-mounted RWA is also accomplished, and a new method of determining higher harmonic responses with the consideration of structural modes and damping is developed. © The author(s) and/or their employer(s), 2012.
Microvibrations, at frequencies between 1 and 1000 Hz, generated by on board equipment, can propagate throughout a spacecraft structure and affect the performance of sensitive payloads. To investigate strategies to reduce these dynamic disturbances by means of active control systems, realistic yet simple structural models are necessary to represent the dynamics of the electromechanical system. In this paper a modeling technique which meets this requirement is presented, and the resulting mathematical model is used to develop some initial results on active control strategies. Attention is focused on a mass loaded panel subjected to point excitation sources, the objective being to minimize the displacement at an arbitrary output location. Piezoelectric patches acting as sensors and actuators are employed. The equations of motion are derived by using Lagrange's equation with vibration mode shapes as the Ritz functions. The number of sensors/actuators and their location is variable. The set of equations obtained is then transformed into state variables and some initial controller design studies are undertaken. These are based on standard linear systems optimal control theory where the resulting controller is implemented by a state observer. It is demonstrated that the proposed modeling technique is a feasible realistic basis for in-depth controller design/evaluation studies.
In this study, the effectiveness of model correlation methods such as modal assurance criterion (MAC) and normalized cross orthogonality check (NCO) in the prediction of forced response characteristics of spacecraft structure are assessed using synthetic modal parameters as well as physically altered finite element models of two real spacecraft structures. It is observed that, neither MAC nor NCO can assure the finite element model capability to represent the acceleration response on the structure and the force transmitted to the base within the acceptable limits during the base excitation. It is also observed that, sometimes model with lower values of MAC or NCO represent the response characteristics better than that of a model with high value of MAC or NCO. A synthetic model is used for the robustness study of spacecraft models under base excitation and this model can approximately represent the response characteristics.
Tape springs, defined as thin metallic strips with an initially curved cross-section, are an attractive structural solution and hinge mechanism for small satellite deployable structures because of their low mass, low cost, and general simplicity. They have previously been used to deploy booms and array panels in various configurations that incorporate a two-dimensional deployment of the tape. However, applications currently exist that incorporate three-dimensional tape springs folds. To accurately model the deployment of an appendage mounted with tape spring hinges, it is necessary to accurately model the opening moments produced from the material strains in the tape spring fold. These moments are primarily a function of curvature. This publication uses a photographic method to analyse the curvature assumptions of two-dimensional tape spring folds and to define the curvature trends for three-dimensional tape spring folds as a basis for calculating the opening moment. It is found that although a variation in the curvature can be seen for three-dimensional tape spring folds, its effect is secondary to the tape thickness tolerance. Therefore, constant curvature models are concluded to be accurate enough for general tape fold applications. © IMechE 2007.
The accumulation of space debris in low Earth orbits poses an increasing threat of collisions and damage to spacecraft. As a low-cost solution to the space debris problem the Gossamer Deorbiter proposed herein is designed as a scalable stand-alone system that can be attached to a low-to-medium mass host satellite for end-of-life disposal from low Earth orbit. It consists of a 5 m by 5 m square solar/drag sail that uses four bistable carbon fiber booms for deployment and support. Prior to deployment of the gossamer structure, a telescopic enclosure system is used to displace the sail from the host craft in order to extend the sail without hindrance from the host peripherals, and also provide passive stabilization. The principal advantage of an entirely passive operational mode allows the drag augmentation system to act as a "fail-safe" device that would activate if the spacecraft suffers a catastrophic failure. Several scenarios are analyzed to study the potential application and performance of the system to current and future missions. A detailed breakdown of the mechanical subsystems of the Gossamer Deorbiter is presented, as well as the characterization process of the deployable booms and sail membrane and the full qualification testing campaign at component and system levels. Finally, the performance scalability of the concept is analyzed. © 2014 IAA. Published by Elsevier Ltd. All rights reserved.
In this paper, the conventional design of an enclosure for electronic equipment for space application is reviewed and an alternative type of construction is proposed. The alternative design is based on the use of carbon fibre reinforced plastic (CFRP) sandwich panels for the construction of the enclosure, and the substitution of the printed circuit board (PCB) antivibration frames (AVFs) with antivibration rods (AVRs). To put this work into context, the requirements applicable to the structural design of this type of unit are briefly reviewed. Standard structural analyses have been performed on the conventional enclosure and then repeated for the proposed configuration, in order to demonstrate its compliance with the fundamental mechanical requirements. The issues concerning the radiation protection offered by the enclosure are discussed, and some solutions to this potential problem are briefly presented. The work demonstrates the possibility of achieving a saving of about 20 per cent on the overall mass of the unit. Finally, the cost of the proposed enclosure is assessed and compared with the conventional design for various missions.
Micro-vibrations on board spacecraft are an issue of growing importance, as some modern payloads, and in particular the new generations of optical instruments require extreme platform stability. These low level mechanical disturbances are usually created by the functioning of mechanical equipment (sources) such as reaction wheels, antenna pointing mechanisms cryo-coolers etc., and cover a wide frequency range. Because of the low level of the vibrations and their wide frequency range, the modeling and analysis of micro-vibrations poses a challenge as the typical structural modeling techniques used in this sector (Finite Element Method (FEM) and Statistical Energy Analysis (SEA)) are reliable only in some areas of the frequency spectrum. The FEM is well suited for low level frequencies; whereas energy methods (e.g. SEA or Energy Finite Element Method EFEA) are suited for high-frequency problems; in the mid-frequency range, finally, other methods (e.g. Hybrid FEA-SEA) tend to be used, even if they’re still not well-established such as the ones named before. However the issue is that there is no single method that can address micro-vibrations in the whole frequency range. In this paper, the methods cited above will be very briefly reviewed and their use in specific micro-vibration prediction problems will be investigated in detail and compared with experimental results. In practice the work presented here uses the Finite Element Method as base-line method to investigate the whole frequency range (say up to 1000 Hz). The FEM predictions are then compared with the experimental results, showing that at medium and high frequencies the response start to deviate significantly from the FEA predictions. The high frequency behavior of the structure will be analyzed using SEA. The mid-frequency range, finally, will be tackled from both directions: from the high frequency side using the Hybrid FE-SEA, whereas from the low frequency side the capability of the standard FEM will be extended using stochastic FEM. The tests are carried out using the structural qualification model of an SSTL satellite bus that has been used to support a high resolution camera. The computational transfer functions and those from the experimental activity will be finally compared using the Modal Assurance Criteria (MAC).
One of the most significant drivers in satellite design is the minimization of mass, in the attempt to reduce the large costs involved in the launch. With technological advances across many fields it is now widely known that very low mass satellites can perform a wide variety of missions. However, the satellite power requirement does not reduce linearly with mass, creating the need for efficient and reliable small satellite deployable structures. One structural solution for this application is tape springs. Tape springs have been previously studied by many countries for space applications focusing on two dimensional systems. This work studies the possible impact of using tape springs folded in three dimensions. By initially analytically determining the static moments created, simple deployment models can be constructed for tape springs in free space. By determining the impact of these moments about an array fold line, a dynamic model of an array can be created which is directly comparable to the two dimensional system. The impact of the three dimensional fold can then be determined.
Sandwich panels have a very high stiffness to weight ratio, which makes them particularly useful in the aerospace industry where carbon fibre reinforced plastics and lightweight honeycomb cores are being used in the construction of floor panels, fairings and intake barrel panels. In the latter case, the geometry of the panels can be considered doubly curved. This paper presents an introduction to an ongoing study investigating the dynamic response prediction of acoustically excited composite sandwich panels which have double curvature. The final objective is to assess and hopefully produce an up to date set of acoustic fatigue design guidelines for this type of structure. The free vibration of doubly curved composite honeycomb sandwich panels is investigated here, both experimentally and theoretically, the latter using a commercially available finite element package. The design and manufacture of three test panels is covered before presenting experimental results for the natural frequencies of vibration with freely supported boundary conditions. Once validated against the experimental results, the theoretical investigation is extended to study the effects of changing radii of curvature, orthotropic properties of the core, and ply orientation on the natural frequencies of vibration of rectangular panels with various boundary conditions. The results from the parameter studies show curve veering, particularly when studying the effect of changing radii and ply orientation, however, it is not clear whether this phenomenon is due to the approximation method used or occurs in the physical system.
An investigation on the structural performance of inserts within honeycomb sandwich panels is presented. The investigation only considers metallic inserts in all aluminum sandwich panels and emphasis is placed on the structural performance difference between hot bonded and cold bonded inserts. The former are introduced during panel manufacture while the latter are potted into existing panels. The investigation only deals with the static performance of the two insert systems subject to loads in the normal direction to the facing plane, which corresponds to their main mode of operation. The experimental part of the work presented involved carrying out pullout tests on hot bonded and cold bonded reference samples by loading them at a centrally located insert. As expected the hot bonded reference samples outperformed the cold bonded reference samples in terms of load carrying capabilities. An analytical model which allows the prediction of shear stress distribution in a circular sandwich panel normally loaded at a centrally located insert is used in an analytical approach for calculating the load carrying capability of inserts. The results from this analytical approach were found to correlate well with the experimental ones for the hot bonded inserts but not for the cold bonded inserts which actually failed at a significantly lower load than was predicted. Copyright © 2008 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
For satellite applications the determination of the correct dynamic behaviour and in particular the structural damping is important to assess the vibration environment for the spacecraft subsystems and ultimately their capability to withstand the launch vibration environment. Therefore, the object of this investigation is to experimentally analyse a range of aluminium panel configurations to study the effect of joints on the damping of the complete structure. The paper begins with a full description of the experimental method used to accurately determine the modal loss factors for each of the panel configurations analysed. Nine different panels were used in the experimental tests, six of which incorporate lap joints variations. The joint parameters investigated include fastener type, bolt torque, fastener spacing, overlap distance and the effect of stiffeners. The damping results of ten different joint variants are presented for each of the first twelve modes of vibration. This data is directly compared to the damping factors of an equivalent monolithic panel. Various specific conclusions are made with respect to each of the joint parameters investigated. However, the primary conclusion is that the mode shape combined with the joint stiffness and joint location can be suggestive as to the likely magnitude increase of the modal loss factor. © 2009 Elsevier Ltd.
Microvibrations, generally defined as low amplitude vibrations at frequencies up to 1 kHz, are now of critical importance in a number of areas. One such area is on-board spacecraft carrying sensitive payloads, such as accurately targeted optical instruments or micro-gravity experiments, where the microvibrations are caused by the operation of other equipment, such as reaction wheels, necessary for its correct functioning. It is now well known that the suppression of such microvibrations to acceptable levels requires the use of active control techniques which, in turn, require sufficiently accurate and tractable models of the underlying dynamics on which to base controller design and initial performance evaluation. Previous work has developed a systematic procedure for obtaining a finite-dimensional state space model approximation of the underlying dynamics from the defining equations of motion which has then been shown to be a suitable basis for robust controller design. This modeling approach is based on the use of Lagrange's equations of motion and is one of a number of possible models possible in this area. In this paper, we describe the experimental validation of this model prior to experimental studies to determine the effectiveness of the designed controllers with the objective of establishing the effectiveness of this procedure both stand alone and against alternatives.
The emerging field of multifunctional structure (MFS) technologies enables the design of systems with reduced mass and volume, thereby improving their overall efficiency. It requires developments in different engineering disciplines and their integration into a single system without degrading their individual performances. MFS is particularly suitable for aerospace applications where mass and volume are critical to the cost of the mission. This article reviews the current state of the art of multifunctional structure technologies relevant to aerospace applications.
Test planning and post-test correlation activity have been issues of growing importance in the last few decades and many methodologies have been developed to either quantify or improve the correlation between computational and experimental results. In this article the methodologies established so far are enhanced with the implementation of a recently developed procedure called Virtual Testing. In the context of fixed-base sinusoidal tests (commonly used in the space sector for correlation), there are several factors in the test campaign that affect the behaviour of the satellite and are not normally taken into account when performing analyses: different boundary conditions created by the shaker’s own dynamics, non-perfect control system, signal delays etc. All these factors are the core of the Virtual Testing implementation, which will be thoroughly explained in this article and applied to the specific case of Bepi-Colombo spacecraft tested on the ESA QUAD Shaker. Correlation activity will be performed in the various stages of the process, showing important improvements observed after applying the final complete methodology.
Microvibration management onboard spacecraft with high stability requirements has drawn increasing interest from engineers and scientists, and this paper discusses a reaction wheel design that allows a significant reduction of mid- to high-frequency microvibrations and that has been practically implemented in industry. Disturbances typically induced by mechanical systems onboard a spacecraft (especially rotating devices such as reaction wheel assemblies and momentum wheel assemblies) can severely degrade the performance of sensitive instruments. Traditionally, wheel-induced high-frequency (over 100-200 Hz) vibrations, generated by a combination of phenomena from bearing noise to dynamic amplifications due to internal resonances, are especially difficult to control. In this paper, the dynamic behavior of a newly designed wheel assembly, with a cantilevered flywheel configuration supported by a soft-suspension system, is investigated. The wheel assembly's mathematical model is developed and later verified with vibration tests. Wheel-assembly-induced lateral and axial microvibrations are accurately measured using a seismic-mass microvibration measurement system, which represents an alternative to typical microvibration measurement setups. Finally, the performance of this wheel assembly in terms of microvibration emissions is compared with a traditional design (with a rigid suspension) through comparison of frequency spectra, and it is shown that this design produces significantly lower vibrations at high frequency. Copyright © 2010 by Zhe Zhang.
For satellite applications the determination of the correct dynamic behaviour and in particular the structural damping is important to assess the vibration environment for the spacecraft subsystems and ultimately their capability to withstand the launch vibration environment. Therefore, the object of this investigation is to experimentally analyse a range of aluminium panel configurations to study the effect of joints on the damping of the complete structure. The paper begins with a full description of the experimental method used to accurately determine the modal loss factors for each of the panel configurations analysed. Nine different panels were used in the experimental tests, six of which incorporate lap joints variations. The joint parameters investigated include fastener type, bolt torque, fastener spacing, overlap distance and the effect of stiffeners. The damping results of ten different joint variants are presented for each of the first twelve modes of vibration. This data is directly compared to the damping factors of an equivalent monolithic panel. Various specific conclusions are made with respect to each of the joint parameters investigated. However, the primary conclusion is that the mode shape combined with the joint stiffness and joint location can be suggestive as to the likely magnitude increase of the modal loss factor.
Due to their high strength to weight ratio and stiffness to weight ratio the use of honeycomb panels is particularly attractive in spacecraft structures. Honeycomb panels are often used in secondary satellite structures such as equipment platforms and solar arrays, but they can also be used as part of the primary structure of a satellite. Indeed honeycomb panel assemblies can be, and are, used to produce efficient and cost-effective primary structures. These types of structures have been used for some time for numerous satellites; however, their development still poses some challenges ranging from the structural performance of the panels themselves to the problem of connecting them to other panels or structural elements. These challenges are faced each time a new satellite is being developed adding cost to the design process. Furthermore, often due to strict timescales in the development process, some of the uncertainties which naturally arise from these challenges cannot always be completely addressed. To compensate for this, conservative design approaches often need to be taken with the ultimate effect of lowering the efficiency of the structure's final design. To meet these challenges and provide a better knowledge base for future satellite development projects a number of research activities have been, and are still, under way at the University of Southampton. The aim of this paper is to describe these research activities and present the key results. ©2010 IEEE.
The influence of the nine orthotropic material properties of honeycomb on the dynamic response of a finite element model of a simple supported sandwich plate are examined. Fifteen available theories from the literature for the material properties of honeycomb are reviewed and their values calculated for a Hex Web 5.2-1/4-25(3003) Aluminium core. The agreement between the theoretical material properties and the major ASTM (American Society for Testing and Materials) standard test methods is investigated. A new and simple technique is described for measuring the dynamic shear moduli of honeycomb materials and its values are compared with those presented in the literature.
There is a wide range of launch opportunities currently available for small- and medium-sized satellites, although the launcher constraints and costs are quite variable. In order to reduce the cost of small satellites it is desirable to define a spacecraft bus which is compatible with as wide a range of vehicles as possible. One of the most promising methods of providing this compatibility for mini-satellites in the range 100-500 kg is by using a stack configuration where the mini-satellite is positioned between a launcher primary payload (LPP) and the launch vehicle. This paper describes the design process for the structure of MiniSIL™, a small satellite bus which is being designed and produced by Space Innovations Limited. © IMechE 1999.
Microvibrations, at frequencies between 1 and 1000 Hz, generated by on board equipment, propagate throughout a spacecraft structure affecting the performance of sensitive payloads. The purpose of this work is to investigate strategies to model and reduce these dynamic disturbances by active control. Initial studies were performed by considering a mass loaded panel where the disturbance excitation source consisted of point forces, the objective being to minimise the displacement at an arbitrary output location. Piezoelectric patches acting as sensors and actuators were used. The equations of motion are derived by using Lagrange's equation with modal shapes as Ritz functions. The number of sensors/actuators and their location is variable. The set of equations obtained is then transformed into state variables and some initial controller design studies have been undertaken. These are based on feedback control implemented using a full state feedback and an observer which reconstructs the state vector from the available sensor signal. Here, the basics behind the structural modelling and controller design will be described. This preliminary analysis will also be used to identify short to medium term further work.
This article examines the new practice of Virtual Shaker Testing (VST), starting from its motivation to its practical implementations and future possible implications. The issues currently experienced during large satellites’ vibration testing are discussed, examining practical examples that highlight the coupling existing between the item under test and facility, and that are the basis for the motivation behind the new methodology (i.e. VST). VST is proposed as a way to bypass some of these issues, and here its use as a pre and post shaker test tool is discussed. In the article VST is applied to real test cases (Airbus’ large spacecraft Bepi Colombo, built for the European Space Agency's first mission to Mercury), showing computations and real physical test data to illustrate the advantages of the methodology. These are mostly in terms of de-risking of the physical test campaigns (due to the capability to simulate realistically the future physical test thus reducing the probability of aborts and stops during the runs), and an improvement of the quality of the correlation process and related FEM update (resulting from the capability to separate the dynamics of the satellite from the effects of the test equipment); ultimately providing a tool to address questions arising from test response observations, which are many. This tool also offers the possibility to improve vibration testing using 6 DOF facilities. The article is concluded articulating a possible way forward to take maximum advantage of the new methodology, drawing a parallel with the current Satellite/Launch Vehicle Coupled Load Analysis cycles, and proposing a different design and validation philosophy.
Precision structures for space-based optical systems are typically subjected to brief periods of random vibration during the launch and ground testing phases. Such events pose a potential threat to the dimensional stability of such structures, which may be required to maintain positional tolerances on large optics in the low 10s of microns to meet optical performance requirements. Whilst there is an abundance of information in the literature on structural instability caused by hygrothermal cycling, there appears to have been little work done on the effects of random vibration. This issue has recently been addressed at RAL with a series of tests aimed at characterizing the behavior of dimensional instability in structures for high-resolution Earth-imaging cameras subject to random vibration. Firstly, a breadboard model of a typical "conventional" CFRP-based optical payload structure was produced and subjected to a range of environmental tests. The effects of random vibration were compared to those of other environmental stressors (such as thermal vacuum testing) and found to be significant. Next, controlled tests were performed on specific structural areas in order to assess the specific contributions of each area to overall instability. These tests made use of novel test setups and metrology techniques to assess the dimensional stability response of material samples and bolted joints to random vibration exposure. The tests were able to measure dimensional instability, characterize it over a series of tests of increasing vibration levels, and assess variability in results between identical samples. Finally, a predictive technique using a Finite Element Model with nonlinear kinematic hardening was produced. A time domain solution was obtained, using an analogy to Miner's Rule to determine load cycle amplitudes. This model correlated reasonably well with test results. This paper presents this program of work, and the results. It also proposes ways to minimize and mitigate dimensional instability due to random vibration by design, analysis and procedural means.
Currently, vibroacoustic problems can be solved using a wide range of numerical techniques. In the low-frequency range, element-based deterministic methods, such as the Finite Element Method (FEM) and Boundary Element Method (BEM) are regularly employed to define the structural and acoustic domains, respectively. The fully coupled FEM-BEM is a classic, vastly popular method. In the high-frequency range probabilistic methods, such as Statistical Energy Analysis, tend to be more efficient and produce more reliable results. Although new techniques are becoming available (e.g. Hybrid FE-SEA Method), the characterisation of the mid-frequency behaviour still poses some challenges, as the computational cost of element-based techniques is often prohibitive, and the modal density is not sufficiently high for statistical approaches to be applicable. This paper discusses an approach aimed at improving the efficiency of the classic FEM-BEM method and potentially extending its usability to the mid-frequency band, specifically in the context of space-craft structural design. The iterative coupling between Craig-Bampton reduced finite element models and BEM is considered as an alternative to directly solving the FEM-BEM coupled equation, allowing the use of efficient procedures for either domain separately. A pre-process enabling the method’s computational implementation is presented, which is based on a manipulation of the reduced mass and stiffness matrices. It is used to allow the application of a distributed load to a Craig-Bampton condensed structure, while mitigating the need to retain a large number of physical degrees of free-dom. The efficiency of the aforementioned matrix modification procedure is compared to that of performing a full Craig-Bampton reduction, and its cost is expressed in terms of floating point operations. An iterative coupling scheme is used on a test-case structure for both a full physical model and a reduced one to verify the concept, and check whether convergence is susceptible to initial conditions, such as the shape of the acoustic field. Finally, the perturbation of the condensed matrices is shown to produce results consistent with those for the full physical model, while substantially reducing the computational effort required for the simulation.
Micro-vibrations on board spacecraft are an issue of growing importance, as some modern payloads, and in particular the new generations of optical instruments require extreme platform stability. These low level mechanical disturbances are usually created by the functioning of mechanical equipment (sources) such as reaction wheels, antenna pointing mechanisms cryo-coolers etc., and cover a wide frequency range. Because of the low level of the vibrations and their wide frequency range, the modeling and analysis of micro-vibrations poses a challenge as the typical structural modeling techniques used in this sector (Finite Element Method (FEM) and Statistical Energy Analysis (SEA)) are reliable only in some areas of the frequency spectrum. The FEM is well suited for low level frequencies; whereas energy methods (e.g. SEA or Energy Finite Element Method EFEA) are suited for high-frequency problems; in the mid-frequency range, finally, other methods (e.g. Hybrid FEA-SEA) tend to be used, even if they're still not well-established such as the ones named before. However the issue is that there is no single method that can address micro-vibrations in the whole frequency range. In this paper, the methods cited above will be very briefly reviewed and their use in specific micro-vibration prediction problems will be investigated in detail and compared with experimental results. In practice the work presented here uses the Finite Element Method as base-line method to investigate the whole frequency range (say up to 1000 Hz). The FEM predictions are then compared with the experimental results, showing that at medium and high frequencies the response start to deviate significantly from the FEA predictions. The high frequency behavior of the structure will be analyzed using SEA. The mid-frequency range, finally, will be tackled from both directions: from the high frequency side using the Hybrid FE-SEA, whereas from the low frequency side the capability of the standard FEM will be extended using stochastic FEM. The tests are carried out using the structural qualification model of an SSTL satellite bus that has been used to support a high resolution camera. The computational transfer functions and those from the experimental activity will be finally compared using the Modal Assurance Criteria (MAC).
Micro-vibration is a low level disturbance, which cannot be controlled or reduced by the Attitude and Orbit Control System of a spacecraft. It can emanate from various sources on a typical spacecraft, notably subsystems with moving parts such as reaction wheels or cooler mechanisms. Micro-vibration can also result from thermo-elastic effects due to stick-slip from differential expansion of parts. It causes problems for sensitive payloads especially high resolution cameras where the demand for higher resolution (which drives stability requirements), has made analysis and control of microvibrations relevant for a larger number of satellites. The availability of mathematical models to represent the disturbance sources in a format that can themselves be coupled to the mathematical model of the structure to perform end-to-end analysis to obtain predictions of the stability level at the receiver is crucial. In the simplest form the sources can be represented with forces and moments of appropriate characteristics, adding also some inertia at the source location. But even for this simplistic implementation it is necessary to have available the details of the forces and moment produced by a source, and these can be either calculated from schematic models of the functioning of the device, or experimentally determined, or a mixture of the two. Typical test techniques applied to micro-vibration measurement /characterization will be described, highlighting advantages and drawbacks of the various methods. A simple experimental apparatus for the measurement of the micro-vibrations emitted by a reaction wheel is presented. A mathematical model of the reaction wheel disturbances is also presented together with its coupling to a typical spacecraft structural model.
In this paper attention is focused on a simply supported panel, with twin patches of piezoelectric material bonded on opposite faces of the panel acting as actuators. A test rig comprising of an aluminium alloy panel has been designed and built. Particular attention has been placed in designing the rig to reproduce as accurately as possible a simple support along all four edges. The deign and analysis of the rig were carried out using the Finite Element (FE) method, and the results of the FE analysis are then compared and validated against the experimental results. A Mechanical Impedance based method, and the Lagrange Rayleigh-Ritz Method were then used to produce mathematical models of the actively controlled panel. These two techniques are chosen as representative of commonly used techniques in the production of mathematical models for active control design studies. The results obtained from the numerical simulations were compared with experimental results in order to assess the accuracy and sensitivity of the modelling techniques.
Inflatable technology for space applications is under continual development and advances in high strength fibers and rigidizable materials have pushed the limitations of these structures. This has lead to their application in deploying large-aperture antennas, reflectors and solar sails. However, many significant advantages can be achieved by combining inflatable structures with structural stiffeners such as tape springs. These advantages include control of the deployment path of the structure while it is inflating (a past weakness of inflatable structure designs), an increased stiffness of the structure once deployed and a reduction in the required inflation volume. Such structures have been previously constructed at the Jet Propulsion Laboratory focusing on large scale booms. However, due to the high efficiency of these designs they are also appealing to small satellite systems. This article outlines ongoing research work performed at the University of Southampton into the field of small satellite hybrid inflatable structures. Inflatable booms have been constructed and combined with tape spring reinforcements to create simple hybrid structures. These structures have been subjected to bending tests and compared directly to an equivalent inflatable tube without tape spring reinforcement. This enables the stiffness benefits to be determined with respect to the added mass of the tape springs. The paper presents these results, which leads to an initial performance assessment of these structures. © 2010 Elsevier Ltd.
The design concept of multifunctional structures allows spacecraft to be more efficient. By creating structural components that can fulfil extra subsystem functions in addition to meeting structural requirements, mass and volume can be saved. The presented example of this is a wing solar array that contains the spacecraft's batteries. The batteries are commercial off the shelf components, used for their lower costs. The paper focuses on understanding the thermal environment and the response of the multifunctional structure to that environment. With this information, it is shown that the concept can be implemented in many orbits without any particular thermal control and that many more orbits could be used with thermal control of a suitable authority. Possible thermal control solutions are reviewed and recommendations for a passive system are made. © 2010 Elsevier Masson SAS. All rights reserved.
In recent years, frequency selective controllers were developed to offer an alternative to the well studied and widely used filtered reference least mean square algorithm. The motivation came from the fact, that many disturbance spectra are dominated by a finite set of frequencies (tones). Hence it was concluded that noticeable reduction could be achieved by just controlling the dominant tones. The main advantages of frequency selective controllers are that stability can be shown easily and the controller is of lower order compared to standard FIR filters. Initially, the here presented controllers were tested in sound experiments and excellent behaviour was demonstrated. Then a vibrating plate test rig was acquired and it was possible to move on to examine the control behaviour for microvibrations. A lump mass was fixed to the plate to create an unbalanced behaviour. Furthermore, 4 piezoelectric components were glued to the plate. The arrangement is, that 2 piezoelectric elements are opposite each other on either side of the plate. Different experimental trials were carried out. It was started with driving one piezoelectric element as a disturbance generator and the use of two more as actuator and accelerometer respectively. Model free and model based controllers were implemented. All controller types achieved an overall attenuation for the whole frequency band of more than 10 dB, not only for the controlled frequencies justifying the approach taken.
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.
This paper develops a modelling technique for equipment load panels which directly produces (adequate) models of the underlying dynamics on which to base robust controller design/evaluations. This technique is based on the use of the Lagrange's equations of motion and the resulting models are verified against those produced by a finite Element Method Model.
This article discusses the microvibration analysis of a cantilever configured reaction wheel assembly. Disturbances induced by the reaction wheel assembly were measured using a previously designed platform. Modelling strategies for the effect of damping are presented. Sine-sweep tests are performed and a method is developed to model harmonic excitations based on the corresponding test results. The often ignored broadband noise is modelled by removing spikes identified in the raw signal including a method of identifying spikes from energy variation and band-stop filter design. The validation of the reaction wheel disturbance model with full excitations (harmonics and broadband noise) is presented and flaws due to missing broadband noise in conventional reaction wheel assembly microvibration analysis are discussed.
This paper describes the validation of a Lagrange-Rayleigh-Ritz technique for the mathematical modeling of mass loaded panels for active control studies. The validation has been carried out by comparing the results produced by a Lagrange-Rayleigh-Ritz model with the results produced by a finite-element model and experimental data. Attention was focussed on a simply supported panel with a lumped mass constrained to its surface to simulate the presence of equipment mounted on the panel and twin piezoelectric patches bonded to the panel, working as sensors and actuators. The design of the experimental rig is described in detail, and a test campaign was carried out to obtain a set of transfer functions characteristic of this plant. The experimental data are then used to validate the predictions of the mathematical model. In particular, it was demonstrated that the Lagrange-Rayleigh-Ritz model was able to reproduce accurately the dynamics of the plant requiring a relatively small number of degrees of freedom.
The modal assurance criterion and normalized cross-orthogonality check are widely used to assess the correlation between the experimentally determined dynamic characteristics and the finite element model predictions. In this paper, the effectiveness of these criteria on the base excitation responses of three spacecraft models is carried out. The dynamic characteristics obtained from a nominal finite element model are considered as experimental or true characteristics, and those obtained from a model produced by introducing errors in the nominal model are considered as analytically predicted characteristics. It is observed that these criteria are not suitable, particularly when the model is used to predict forced response characteristics such as the force transmitted to the base, peak acceleration response, and dynamic displacement in the spacecraft. Thus, a qualitative indicator named as base-force assurance criterion is defined by comparing the experimentally determined dynamic force at the base and the finite element predicted force such that the criterion can state the possible error in the peak acceleration and the dynamic displacement under the base excitation. The method is applied to assess the performance of three spacecraft structures, and the results show that new criterion can better correlate with the acceleration and the dynamic displacement error than the conventional criteria. Copyright © 2013 by K. K. Sairajan and G. S. Aglietti.
The study presented in this paper concerns the development of an algorithm, based on finite element analysis, for the dynamic simulation of a tethered lighter-than-air balloon when subjected to operational conditions. The main features of the algorithm are described, highlighting the advantages of this approach when performing dynamic analysis. The input parameters considered in the method are derived from experimental and simulated data, which are elaborated to obtain the static and dynamic atmospheric properties in terms of mean windspeed profile, discrete gusts, and continuous turbulence. In particular, the algorithm is employed to perform a thorough analysis of a specific high-altitude tethered platform operating in a realistic design scenario. The dynamic behavior of the system is evaluated in terms of displacements and tether forces. The results, even though they are obtained for a specific example application, demonstrate the general viability of the algorithm proposed for the evaluation of the dynamic response of high-altitude tethered platforms and for the preliminary assessment of the technical feasibility of these systems. © Copyright 2010.
One of the most significant drivers in satellite design is the minimization of mass to reduce the large costs involved in the launch. With technological advances across many fields, it is now widely known that very low-mass satellites can perform a wide variety of missions. However, the satellite power requirement does not reduce linearly with mass, creating the need for efficient and reliable small satellite deployable structures. One possible structural solution for this application is tape springs. Tape springs have been previously studied in many countries for space applications focusing on two-dimensional systems. This work studies the possible impact of using tape springs folded in three dimensions. By first analytically determining the static moments created, simple deployment models can be constructed for tape springs in free space. Then, determining the impact of these moments about an array fold line allows the creation of a dynamic model of an array that is directly comparable to the two-dimensional system. The impact of the three-dimensional fold can then be determined.
Microvibrations, generally defined as low-amplitude vibrations at frequencies up to 1 kHz, are of critical importance in a number of areas. It is now well known that, in general, the suppression of such microvibrations to acceptable levels requires the use of active control techniques which, in turn, require sufficiently accurate and tractable models of the underlying dynamics on which to base controller design and initial performance evaluation. Previous work has developed a systematic procedure for obtaining a finite-dimensional state-space model approximation of the underlying dynamics from the defining equations of motion, which has then been shown to be a suitable basis for robust controller design. In this paper, the experimental validation of this model prior to experimental studies is described in order to determine the effectiveness of the designed controllers. This includes details of the experimental rig and also the use of methods for assessing the safety of the resulting structure against uncertain parameters. © IMechE 2004.
During launch, a spacecraft undergoes loads ranging from quasi-static to highly transient or harmonic low frequency events, from higher frequency shock loads to acoustic excitations. In order to reproduce such a dynamic diversity, fixed base sinusoidal tests, wide band acoustic loading and different regimes of shock testing are implemented in the test facilities. In this article, the main focus is on fixed base sinusoidal tests, fundamental for a number of reasons, including demonstrating that the satellite can withstand the low frequency dynamic environment and validating the mathematical model which will then be also used for coupled load analysis purposes. For the latter, a post-test correlation process is carried out and the basic assumption is trusting the experimental results obtained from shaker testing. In reality, some of these assumptions (e.g. “infinitely” stiff boundary and inertial properties of the shaker) are not correct, as for the kind of applications treated in this article experimental results are significantly affected by boundary flexibilities, modes of the shaker/head expander and non-perfect implementation of the control algorithm in the electronic hardware. In the last decade, there has been a growing interest in virtual testing, with the long-term view to use simulation as substitute for the majority of testing, but currently under investigation for pre-test response predictions and post-test correlation. Here, the satellite is mathematically modelled along with the shaker and the control system. In this article, in particular, a simulation capability of longitudinal closed loop control simulation of the ESA electrodynamic shaker (QUAD) flexible body coupled with a test specimen (Bepi Colombo) flexible model is developed. This shows how significant the differences are when looking at the analytical results from two different perspectives (standard Finite Element Analysis and Virtual Testing implementation). The focus of this article is specifically on post-test correlation: correlation methods are used for both procedures and results show significant improvements when the satellite Finite Element Model undergoes the virtual testing approach.
Deployable structures are required for many satellite operations, to deploy booms for communications or area deployment for power generation, and many sophisticated mechanisms have been developed for these types of structures. However, tape springs, defined as thin metallic strips with an initially curved cross-section, are an attractive structural solution and hinge mechanism for satellite deployable structures because of their low mass, low cost and general simplicity. They have previously been used to deploy booms and array panels in various configurations that incorporate small two-dimensional tape hinges, but they also have the potential to be used in greater numbers to create larger, more geometrically complicated deployable structures. The aim of this work is to investigate a computationally efficient method of simulating these tape spring based deployable structures and to determine the limitations of the analysis approach. The study focuses on a specific deployable structure layout that incorporates 148 separate tape spring elements in three fold lines. The static and dynamic properties of the system are initially investigated experimentally allowing the basic parameters of the theoretical model to be determined accurately. It was found that the simulated tape pair stiffness was a key parameter affecting the dynamic properties of the model and the peak shock accelerations. It was concluded that the model was capable of closely simulating the dynamic 'snap through' behaviour of the wall. However, the torsional stiffness around the axis normal to the plane of the structure was found to be too large, resulting in over approximations of the peak shock accelerations.
Microvibrations of a RWA are usually studied in either hard-mounted or coupled conditions, although coupled wheel-structure disturbances are more representative than the hard-mounted disturbances. The coupled analysis method of the wheel-structure is not as well developed as the hard-mounted one. A coupled disturbance analysis method is proposed in this paper. One of the most important factors in coupled disturbance analysis - the accelerance or dynamic mass of the wheel is measured and results are validated with an equivalent FE model. The wheel hard-mounted disturbances are also measured from a vibration measurement platform particularly designed for this study. Wheel structural modes are solved from its analytical disturbance model and validated with the test results. The wheel-speed dependent accelerance analysis method is proposed.
High precision stable structures are potentially vulnerable to dimensional instability induced by exposure to random vibration. There appears to have been little work in the literature to understand or mitigate structural dimensional instability induced by random vibration. To gain more insight into this issue, a novel test was recently developed to assess the plastic strain response in the 10 to 10 range for structural materials subjected to specific random vibration loads. The test was based on a four-point bending configuration with an applied random base excitation. Two types of material were tested-an Al alloy and a CFRP. This paper presents the test setup and results in detail. The Al alloy samples were found to grow slightly in length during testing, due to a small non-symmetry in the applied load. An FEA model of the test setup was solved in the time domain for a sequence of cyclic loads whose amplitude was based on their probability of exceedance in the random environment. This model, using nonlinear kinematic hardening, was able to predict the residual strain response observed during testing with good accuracy. The main implication of this finding is that ultra stable structures subject to random vibration should be assembled in the most strain-free state possible to avoid loss of dimensional stability due to cyclic hardening. © 2012 Elsevier Inc. All rights reserved.
Reaction wheel assemblies are one of the most important microvibration sources on typical modern satellites. In this paper microvibrations induced by a cantilevered reaction wheel assembly are modelled and validated against microvibration test results. The disturbance model is developed using energy method. A microvibration measurement platform is designed to measure its disturbances. Disturbance test results are analyzed in detail. The peculiar dynamic characteristics such as nonlinearity and high damping of harmonic responses in the test results are discussed. Estimations of damping values used in the disturbance model are introduced. A new method developed to model harmonic excitations is discussed. Furthermore, novel methods to identify harmonics and extract model parameters from test results are presented. The empirical modeling method developed for broadband noise excitations are also introduced and validated. © (2012) Trans Tech Publications, Switzerland.
A multifunctional structure reduces the mass and volume of a spacecraft through the removal of parasitic components of a functional structure, such as the purely structural "packaging" of the battery. Commercially available cells, of the plastic lithium-ion type, may be incorporated into structural sandwich panels, eliminating the need for a secondary structure in the battery subsystem. Locating the cells in the structure also removes them from the bus, which reduces the volume of the craft and, consequently, reduces the mass of the primary structure. Although the batteries studied in this work exhibited low mechanical properties, this paper will show that, by placing the cells carefully within the sandwich panel, structural performance is not compromised. Finite element models show a reduction in peak stress and deformation in multifunctional panels, compared to a conventional design, when a favourable layout is selected. Less conclusive results for peak acceleration, however, suggest that this type of multifunctional structure may not be appropriate for all applications. Comparison of the finite element modelling technique with a real panel's behaviour shows that the deformation and stress predicted by the model is consistent with reality, whilst the acceleration is reliable for low frequencies. © 2010 Elsevier Ltd. All rights reserved.
Purpose - The purpose of this paper is to assess the suitability of various methods for the reduction of a large finite element model (FEM) of satellites to produce models to be used for correlation of the FEM with test results. The robustness of the cross-orthogonality checks (COC) for the correlation process carried out utilizing the reduced model is investigated, showing its dependence on the number of mode shapes used in the reduction process. Finally the paper investigates the improvement in the robustness of the COC that can be achieved utilizing optimality criteria for the selection of the degrees of freedom (DOF) used for the correlation process. Design/methodology/approach - A Monte Carlo approach has been used to simulate inaccuracies in the mode shapes (analysis and experimental) of a satellite FEM that are compared during the COC. The sensitivity of the COC to the parameters utilized during the reduction process, i.e. mode shapes and DOFs, is then assessed for different levels of inaccuracy in the mode shapes. Findings - The System Equivalent Expansion Reduction Process (SEREP) has been identified as a particularly suitable method, with the advantage that a SEREP reduced model has the same eigenvalues and eigenvector of the whole system therefore automatically meeting the criteria on the quality of the reduced model. The inclusion of a high number of mode shapes in the reduction process makes the check very sensitive to minor experimental or modelling inaccuracies. Finally it was shown that utilizing optimality criteria in the selection of the DOFs to carry out the correlation can significantly improve the probability of meeting the COC criteria. Research limitations/implications - This work is based on the FEM of the satellite IT>Aeolus/IT>, and therefore the numerical values obtained in this study are specific for this application. However, this model represents a typical satellite FEM and therefore the trends identified in this work are expected to be generally valid for this type of structure. Practical implications - The correlation of satellite FEM with test results involves a substantial effort, and it is crucial to avoid failures of the COC due to numerical issues rather than real model inaccuracies. This work shows also how an inappropriate choice of reduction parameters can lead to failure of the COC in cases when there are only very minor differences (e.g. due to minor amount of noise in the results) between analytical and test results. Vice versa, the work also shows how the robustness of the reduced model can be improved. Originality/value - The paper shows how the robustness of the correlation process for a satellite FEM carried out utilising a SEREP reduced model needed to be investigated, to demonstrate the suitability of this method to reduce large FEM of satellites. Copyright © 2012 Emerald Group Publishing Limited. All rights reserved.
Active control techniques are often required to mitigate the micro-vibration environment existing on board spacecraft. However, reliability issues and high power consumption are major drawbacks of active isolation systems that have limited their use for space applications. In the present study, an electromagnetic shunt damper (EMSD) connected to a negative-resistance circuit is designed, modelled and analysed. The negative resistance produces an overall reduction of the circuit resistance that results in an increase of the induced current in the closed circuit and thus the damping performance. This damper can be classified as a semi-active damper since the shunt does not require any control algorithm to operate. Additionally, the proposed EMSD is characterised by low required power, simplified electronics and small device mass, allowing it to be comfortably integrated on a satellite. This work demonstrates, both analytically and experimentally, that this technology is capable of effectively isolating typical satellite micro-vibration sources over the whole temperature range of interest
Due to constantly increasing requirements for more precise and high-resolution instrumentations, microvibration prediction represents an issue of growing importance. Hence the need of reliable analysis tools which can evaluate microvibrations effects efficiently. This paper describes how to tackle the issue of structural uncertainties in microvibration predictions. In particular, uncertainties related to the microvibration sources are analysed as well as those linked to the modelling of the structure. A methodology to define the worst case of vibration produced by on board sources is presented and compared to experimental data. Additionally, an approach to quantify the uncertainties in the Finite Element model is also described.
In this paper, a novel stochastic finite element method is introduced. The concept is based on a set of strict necessary and sufficient requirements for nonnegative definiteness of Hermitian matrices with a 2x2 block partitioning, expressed in terms of their constituent submatrices. Further mathematical constructions are suggested, permitting the robust and efficient construction and sampling of random mass and stiffness matrices. The method allows uncertainty to be controlled at the subsystem level in dynamic substructuring problems. Different levels of randomness can be applied to off-diagonal partitions of the component matrices without interfering with the remaining blocks, or the key mathematical properties of the global matrix. Sparsity pattern of the ‘nominal’ deterministic matrix is preserved. The method is validated with a spacecraft test case in a vibroacoustic load scenario. Very good results are demonstrated against direct parametric Monte Carlo simulation, while computational time is reduced by nearly 3 orders of magnitude.
In recent years, driven by the increasingly stringent stability requirements imposed by some satellites’ payloads (e.g., the new generation of optical instruments), the issue of accurate onboard spacecraft microvibration modeling has attracted significant interest from engineers and scientists. This paper investigates the microvibration-induced phenomenon on a cantilever-configured reaction wheel assembly including sub- and higher harmonic amplifications due to modal resonances and broadband noise. A mathematical model of the reaction wheel assembly is developed and validated against experimental test results. The model is capable of representing each configuration in which the reaction wheel assembly will operate, whether it is hard mounted on a dynamometric platform or suspended free–free. The outcomes of this analysis are used to establish a novel methodology to retrieve the dynamic mass of the reaction wheel assembly in its operative range of speeds. An alternative measurement procedure has been developed for this purpose, showing to produce good estimates over a wide range of frequencies using a less complex test campaign compared with typical dynamic mass setups. Furthermore, the gyroscopic effect influence in the reaction wheel assembly response is thoroughly examined both analytically and experimentally. Finally, to what extent the noise affects the convergence of the novel approach is investigated.
© 2015 Elsevier Ltd. All rights reserved. It is well documented that at frequencies beyond the first few modes of a system, the Finite Element Method is unsuitable to obtain efficient predictions. In this article, it is proposed to merge the efficiency of the Craig-Bampton reduction technique with the simplicity and reliability of Monte Carlo Simulations to produce an overall analysis methodology to evaluate the dynamic response of large structural assemblies in the mid-frequency range. The method (Craig-Bampton Stochastic Method) will be described in this article with a benchmark example shown and implemented in the theory of the dynamic coupling extended to the case when multiple sources of microvibrations act simultaneously on the same structure. The methodology will then be applied to a real practical application involving the modern satellite SSTL 300 S1.
In this paper, a full methodology to deal with microvibration predictions onboard satellites is described. Two important aspects are tackled: 1) the characterization of the sources with a pragmatic procedure that allows integrating into the algorithm the full effect of the sources, including their dynamic coupling with the satellite structure; 2) the modeling of the transfer function source receivers with a technique named in this paper as the Craig-Bampton stochastic method, which allows prediction of a nominal response and variations due to structural uncertainties as accurate as full Monte Carlo simulations but at a fraction of the computational effort. The paper then includes a statistical study of the data from the structural dynamic testing of the five identical craft of the Rapid-Eye constellation to set the magnitude of the uncertainties that should be applied in the analysis. Finally, the computational procedure is applied to the new high-resolution satellite SSTL-300-S1 and the predictions compared with the experimental results retrieved during the physical microvibration testing of the satellite, which was carried out at the Surrey Satellite Technology Limited facilities in the United Kingdom.
Efficient vibroacoustic response prediction on complex structures, such as spacecraft, represents a challenging task, even for the computers and numerical techniques of today. This is particularly evident in the mid-frequency range, where structures begin exhibiting chaotic behaviour, rendering element-based techniques inefficient or unreliable. In this article, an efficient random formulation for reduced finite element method (FEM) models is proposed, such that Monte Carlo simulations can be carried out robustly within practically acceptable timeframes. The introduced novel non-parametric stochastic FEM is inherently compatible with various existing component mode synthesis techniques. It is particularly well adapted to use with popular modal reduction approaches, such as the Craig-Bampton method. The mathematical framework for the method is outlined, enabling the deterministic reduced matrices to be robustly perturbed at the subsystem level. Properties, such as matrix positive-(semi)definiteness, mean system eigenvalues, and representation accuracy are preserved. This new stochastic FEM is validated against a full parametric Monte-Carlo simulation and test data of a real spacecraft structure, establishing its reliability and computational efficiency. In the proposed coupled FEM-BEM approach, the acoustic domain is modelled with hierarchical matrix accelerated collocation BEM. This alleviates the memory requirements for the large, dense BEM matrices, and the need for spatial discretisation of acoustic FEM. The full implementation is outlined for a simple geometry discretised with high a density mesh, showing consistent convergence of the employed iterative solver.
This paper addresses the susceptibility of finite element models to uncertainty in frequency ranges with relatively high modal density, particularly in the con- text of vibroacoustic analysis. The principal idea is based on a stochastic fi- nite element method (FEM) technique called Craig-Bampton stochastic method (CBSM). It is a parametric Monte Carlo simulation (MCS) approach that can be performed at a fraction of the otherwise potentially impractical computational cost, due to the use of reduced rather than full system matrices. An enhanced formulation of the CBSM, significantly improving its efficiency by exploiting the block structure of the condensed model’s stiffness and mass matrices is derived. The improved method is adapted for use with distributed loads, such as diffuse sound field excitation. Its practical implementation is illustrated through a simple theoretical example followed by a high-complexity spacecraft structure case. In both cases solutions are compared to those of a classic MC simulation of the non-condensed models. Through an extensive parametric survey, recommendations are given on the ideal perturbation levels and underlying statistical distributions for the improved CBSM’s random variables. The proposed technique shows a very strong agreement with the benchmark MC results. Computational time reductions of over 1 and 3 orders of magnitude against the original CBSM and the MC simulation, respectively, are demonstrated.
The term "microvibrations" generally refers to accelerations in the region of micro-g, occurring over a wide frequency range, up to say 500-1000 Hz. The main issues related to microvibrations are their control and minimisation, which requires their modelling and analysis. A particular challenge is posed in the mid-frequency range, where many of the micro-vibration sources on board a spacecraft tend to act. In this case, in addition to the typical issues related to predicting responses in the mid-frequency, the low amplitude of the inputs can produce further non-linear behaviour which can manifest as uncertainties. A typical example is the behaviour of cables secured onto panels; when very low forces are applied, the presence of harness can influence the characteristics of the panel in terms of stiffness and damping values. In these circumstances, the cables themselves couple with the panel, hence become paths for vibration transmission. The common practise is to model such cables as Non-Structural Mass; however, this paper illustrates that this method does not yield accurate results. In order to demonstrate this, an experimental campaign was conducted investigating a honeycomb panel, which was tested bare and with different configurations of harness secured to it. The results of this experimental campaign showed significantly different behaviour of the structure depending on the amplitude of the loads and the frequency. In particular, it was found that the effects the addition of the cables had on the panel were different depending on the frequency range considered. Based on this observation, a general methodology to deal with the whole frequency range is presented here and the basis to extend it to the case of more complex structures is also proposed.
The ROV-E project is a three year Framework 7 project dedicated to the exploration of multifunctional technologies for use on Mars rovers. As part of this the University of Southampton has been looking at the development of a Multifunctional Power Structure. This is a structure that combines aspects of the electrical power system into a single component, removing the parasitic structures needed to support distributed discrete components in the Mars rover bus. Looking to exploit the cost benefits of using off the shelf components, commercially available lithium polymer cells have been exposed to structural, temperature and pressure environments and have proved to be robust throughout the lifecycle of the panel. A concept optimisation has shown that the design of the panel is a trade-off between not only capacity and strength, but also between mechanical loading of the cells and panel stiffness. A representative panel was manufactured, showing that it is possible to respect the limits of the selected cells and still create a valid component of the rover. This panel was then used experimentally to assess the failure methods of the cells, revealing that the cells are more likely to suffer performance loss due to bending than accelerations. The work has then moved onto an assessment of the thermal control support required by multifunctional power structures in the mars environment. A study of current thermal control technology found that there exists no low mass solutions for maintaining cell temperature if the panel is part of a deployed structure. Even when the panel is part of the rover primary bus, only heaters were shown to be a mass efficient solution. The study did reveal that phase change materials could be an effective thermal control for intermittently used electronics. The final section of work completed was a study into the embedding of phase change materials into sandwich panel cores, which found the concept workable, provided the material is concentrated under the component. Copyright © 2013 by the International Astronautical Federation.
The ROV-E project is a three year European Union Framework 7 project, which began in January 2011, dedicated to the research and development of lightweight technologies for exploration rovers. As part of this the University of Southampton, along with other consortium members, have been looking into the development of a Multifunctional Power Structure (MFPS). This is a structure that combines aspects of the electrical power system into a single panel component, removing the unnecessary mass of additional structures and containers required to support distributed discrete components inside a rover. The specific components imbedded into the multifunctional panel include: power generation (photovoltaic cells), control electronics and power storage. The main focus of the research at the University of Southampton was the power storage function of the panel, which aimed at exploiting the cost benefits of using off the shelf components by using commercially available lithium polymer battery cells. Initial validation testing exposed these cells to structural, temperature and pressure environments which proved the robustness of the cells throughout the predicted lifecycle of the multifunctional panel. An initial representative honeycomb panel incorporating battery cells was constructed to validate the manufacturing process. This panel was then used experimentally to assess the failure methods of the cells, revealing that the cells are more likely to suffer performance loss due to bending than accelerations. Following on from the initial validation testing a full MFPS was designed and optimised before being subjected to mechanical and thermal environments. This paper focuses on the final design and testing of this complete MFPS. Although the testing encountered various unforeseen problems, the batteries were both mechanically and thermally validated as part of the complete MFPS.
This paper presents a set of preliminary calculations to assess the technical feasibility of a lighter than air tethered platform that would be used as a support for a photovoltaic array, in order to harness the solar power at high altitude and transmit it to the ground via the mooring cable. This solution would allow to collect significantly more solar radiation if compared to a traditional ground based photovoltaic array, especially in countries where the solar resource is scarce due to their location and weather conditions. The technical feasibility of the system is preliminarily evaluated with the use of a simplified model, which simulates the response of a tethered spherical balloon to the environmental conditions that can be found during the ascent and at operational altitude. The contribution of the main components (PV array, transmission tether...) to the overall weight of the system and their influence on the total lift are evaluated. On the basis of the results obtained the possibility of introducing additional components to optimize the performance of the system is investigated. ©2009 IEEE.
Mercer JF, Kiley AM, Aglietti GS (2014)BepiColombo: sine test FEM correlation experiences, In: Sas P, Moens D, Denayer H, (eds.), PROCEEDINGS OF INTERNATIONAL CONFERENCE ON NOISE AND VIBRATION ENGINEERING (ISMA2014) AND INTERNATIONAL CONFERENCE ON UNCERTAINTY IN STRUCTURAL DYNAMICS (USD2014)pp. 835-849
This paper reflects on some of the experiences of correlating the Finite Element Model (FEM) of ESA's BepiColombo spacecraft to sine sweep test measured data, in order to validate the model for subsequent Coupled Loads Analysis (CLA). Post-model update procedures can take a significant number of man hours to complete, without necessarily resulting in a final FEM which is notably more representative of the real structure than the initial FEM. The long term research intention is to use the lessons learnt from BepiColombo and other spacecraft correlations to work towards the containment of the FEM correlation process, this paper addresses part of this on-going research effort. This is to be achieved through: investigating the current techniques/algorithms for FEM test analysis correlation and validation; identifying their limitations; thus ultimately developing a methodology which identifies where model updates are or are not necessary to achieve a model which is validated for key participating modes in flight predictions or during any sine test notching.
This article discusses the coupled microvibration analysis of a cantilever configured Reaction Wheel Assembly with soft-suspension system. A RWA-seismic mass coupled microvibration measurement system is presented and its model validated against test results. The importance of the RWA driving point accelerances in coupled microvibration analysis is thoroughly discussed. A RWA accelerance measurement system has been designed to measure the driving point accelerances in both static (flywheel not spinning) and dynamic (flywheel spinning) conditions. Analytically, RWA static accelerance is obtained by frequency response analysis of a finite element model. The traditionally ignored gyroscopic effects in the accelerances are included in the model and their effects with respect to traditional models are shown both theoretically and experimentally. Although at high angular speed, when nonlinearities in the microvibrations prevent an accurate simulation, it is shown that the predicted microvibrations match more closely with the test results when considering gyroscopic effects in RWA accelerances than those predicted using the traditional method. The presented coupled microvibration analysis method is also very efficient in practice and is applicable in an industrial environment. © 2013 Elsevier Ltd. All rights reserved.
Driven by the increasingly stringent stability requirement of some modern payloads (e.g. the new generations of optical instruments) the issue of accurate spacecraft micro-vibration modelling has grown of importance. This article focuses on the dynamic coupling between a source of micro-vibration (e.g. reaction wheel) and a structure, taking into account the uncertainties related to both parts. In this context, an alternative to the Monte Carlo Simulation for complex structures has been developed, consisting in sub-structural approach to perturb the natural frequencies of specific subsystem reduced with the Craig-Bampton method. In order to prove the validity of the method and its application to the theory of the coupling, benchmark examples and practical applications will be described.
Microvibrations of a satellite reaction wheel assembly are commonly analysed in either hard-mounted or coupled boundary conditions, though coupled wheel-to-structure disturbance models are more representative of the real environment in which the wheel operates. This article investigates the coupled microvibration dynamics of a cantilever configured reaction wheel assembly mounted on either a stiff or flexible platform. Here a method is presented to cope with modern project necessities: (i) need of a model which gives accurate estimates covering a wide frequency range; (ii) reduce the personnel and time costs derived from the test campaign, (iii) reduce the computational effort without affecting the quality of the results. The method involves measurements of the disturbances induced by the reaction wheel assembly in a hard-mounted configuration and of the frequency and speed dependent dynamic mass of the reaction wheel. In addition, it corrects the approximation due to missing speed dependent dynamic mass in conventional reaction wheel assembly microvibration analysis. The former was evaluated experimentally using a previously designed and validated platform. The latter, on the other hand, was estimated analytically using a finite element model of the wheel assembly. Finally, the validation of the coupled wheel-structure disturbance model is presented, giving indication of the level of accuracy that can be achieved with this type of analyses.
An increasing demand for high-quality, low cost Earth imagery has led to the requirement for improved structural stability of the satellite instruments providing the imagery. This translates into camera structures capable of maintaining very high levels of dimensional stability over their lifetimes. There are several adequate technical solutions to this problem-for instance onboard mechanisms that can re-align optical components operationally, or unconventional materials such as silicon carbide. This paper investigates the current state of the art in solutions which are based on conventional materials and joining techniques to maintain high levels of stability without the need for mechanisms. The purpose of this article is to give an overview of the most important factors that should be taken into account when designing stable structures for space applications. Whilst the use in Earth observation has been highlighted in this paper the overview is applicable to all applications that use space optics. For completeness, the methods to verify the design have also been briefly described. © 2009 Elsevier Ltd. All rights reserved.
The current paper examines the feasibility of using a high altitude tethered aerostat as a platform for producing a substantial quantity of electric energy and transmitting it to Earth using the mooring cable. Based on realistic values for the relevant engineering parameters that describe the technical properties of the materials and subsystems, a static analysis of the aerostat in its deployed configuration has been carried out. The results of the computations, although of a preliminary nature, demonstrate that the concept is technically feasible. There are, nevertheless, issues to be addressed to improve the performance. However none of these issues is deemed to negate the technical feasibility of this concept. A test case is investigated in terms of preliminary sizing of the aerostat, including mooring cable and solar cell coverage, and it shows the capability to deliver power to the ground in excess of 95 kW. A brief assessment of the cost has also been carried out to investigate the potential gains offered by this system to produce solar electric energy. © IMechE 2008.
Forshaw Jason, Aglietti Guglielmo, Salmon T, Retat I, Roe M, Burgess C, Chabot T, Pisseloup A, Phipps A, Bernal C, Chaumette F, Pollini A, Steyn WH (2017)Final Payload Test Results for the RemoveDebris Active Debris Removal Mission, In: Acta Astronautica138pp. 326-342
Since the beginning of the space era, a significant amount of debris has progressively been generated in space. Active Debris Removal (ADR) missions have been suggested as a way of limiting and controlling future growth in orbital space debris by actively deploying vehicles to remove debris. The European Commission FP7-sponsored RemoveDebris mission, which started in 2013, draws on the expertise of some of Europe’s most prominent space institutions in order to demonstrate key ADR technologies in a cost effective ambitious manner: net capture, harpoon capture, vision-based navigation, dragsail de-orbiting. This paper provides an overview of some of the final payload test results before launch. A comprehensive test campaign is underway on both payloads and platform. The tests aim to demonstrate both functional success of the experiments and that the experiments can survive the space environment. Space environmental tests (EVT) include vibration, thermal, vacuum or thermalvacuum (TVAC) and in some cases EMC and shock. The test flow differs for each payload and depends on the heritage of the constituent payload parts. The paper will also provide an update to the launch, expected in 2017 from the International Space Station (ISS), and test philosophy that has been influenced from the launch and prerequisite NASA safety review for the mission. The RemoveDebris mission aims to be one of the world’s first in-orbit demonstrations of key technologies for active debris removal and is a vital prerequisite to achieving the ultimate goal of a cleaner Earth orbital environment.
Spacecraft overtesting is a long running problem, and the main focus of most attempts to reduce it has been to adjust the base vibration input (i.e. notching). Instead this paper examines testing alternatives for secondary structures (equipment) coupled to the main structure (satellite) when they are tested separately. Even if the vibration source is applied along one of the orthogonal axes at the base of the coupled system (satellite plus equipment), the dynamics of the system and potentially the interface configuration mean the vibration at the interface may not occur all along one axis much less the corresponding orthogonal axis of the base excitation. This paper proposes an alternative testing methodology in which the testing of a piece of equipment occurs at an offset angle. This Angle Optimisation method may have multiple tests but each with an altered input direction allowing for the best match between all specified equipment system responses with coupled system tests. An optimisation process that compares the calculated equipment RMS values for a range of inputs with the maximum coupled system RMS values, and is used to find the optimal testing configuration for the given parameters. A case study was performed to find the best testing angles to match the acceleration responses of the centre of mass and sum of interface forces for all three axes, as well as the von Mises stress for an element by a fastening point. The angle optimisation method resulted in RMS values and PSD responses that were much closer to the coupled system when compared with traditional testing. The optimum testing configuration resulted in an overall average error significantly smaller than the traditional method. Crucially, this case study shows that the optimum test campaign could be a single equipment level test opposed to the traditional three orthogonal direction tests.
This paper addresses the characterisation and analysis of a 2-collinear-DoF strut with embedded electromagnetic shunt dampers. The use of a negative resistance in the shunt circuit has been proved to considerably enhance the damping performance of this kind of electromagnetic dampers. The analytical model is reported and the theoretical results are compared with other damping methods. This work demonstrates the feasibility of achieving a remarkable decay rate of -80 dB/decade with a device that is smaller than previously-presented active struts and does not require complex electronics to operate.
This paper presents a preliminary assessment of the potential advantage that a high altitude solar collector could bring compared with the traditional ground based photovoltaic systems. This advantage mainly derives from the reduced attenuation of the solar radiation as it travels through the atmosphere especially if clouds are present above the location considered. A sun beam traveling through clear atmosphere is considered first using an existent model to calculate the daily irradiation at different altitudes in clear sky conditions. The results obtained are then integrated with experimental data describing cloud distributions versus altitude, and finally the contribution of the diffused radiation is also included to give a realistic evaluation of the total actual irradiation at a specific altitude. The results are obtained for a specific location in the UK, where the experimental data have been acquired. The general conclusions, however, involving the potential of high altitude solar collectors, can be extended to other countries in Europe with similar climates. Finally, the main issues involved in the design and development of a flying platform for the exploitation of the solar energy are presented and the technical feasibility of the system is discussed. © 2010 by ASME.
In the recent years, micro-vibrations have been an issue of growing importance, due to the high-stability requirements imposed by some modern payloads. These low level mechanical disturbances, occurring at frequencies from sub hertz up to 1000 Hz, are created by different sources in the spacecraft (e.g. reaction wheels) and how to model the micro-vibration environment is currently under investigation. In this paper, a methodology is presented, involving analyses techniques such as FEA (reliable at low frequencies), Monte Carlo Simulation (precise but still computationally demanding and time consuming) and Modal Hybridization (a Stochastic Finite Element Method which involves perturbation of modes and natural frequencies and will be used to refine the general methodology). The various modelling techniques also require a particular attention when dealing with micro-vibrations. For instance, mechanical equipment typically on board a spacecraft (such as harness, thermal straps etc.) affects the response caused by low level disturbances and a FEM which models them with simple non-structural mass appears to be not accurate enough. Another important aspect of the modelling is the coupling between the sources and the tested structure. All the methods described above will be applied to a bench-work model represented by the satellite platform SSTL 300 (the relative testing campaign will be also described in this paper) and comparisons between the experimental and the computational results will be performed.
An effective investigation of alternative control strategies for the reduction of vibration levels in satellite structures requires realistic, yet efficient, structural models to simulate the dynamics of the system. These models should include the effects of the sources, receivers, supporting structure, sensors, and actuators. In this paper, a modeling technique which meets these requirements is developed and some active control strategies are briefly investigated. The particular subject of investigation is an equipment-loaded panel and the equations of motion are derived using the Lagrange-Rayleigh-Ritz (LRR) approach. The various pieces of equipment on the panel are mounted on active or passive suspensions, and resonators are used to represent the internal dynamics of the mounted equipment. Control of the panel, which transmits vibrations from sources to receivers, is by means of piezoelectric patches and the excitation consists of dynamic loads acting on the equipment enclosures and/or directly on the panel. The control objective is to minimize the displacement at an arbitrary output location. The LRR model developed is verified against one produced by using the finite-element method. Finally, some initial controller design studies are undertaken to investigate and compare the effectiveness of different control strategies (e.g., minimization at the source, along the vibration path, or at the receiver).
The vibration study of a general three-layer conical sandwich panel based on the h-p version of the finite element method is presented in this paper. No restriction is placed on the degree of curvature of the shell, thereby relaxing the strictures associated with shallow shell theory. The methodology incorporates a new set of trigonometric functions to provide the element p-enrichment, and elements may be joined together to model either open conical panels, or complete conical frusta (circumferentially connected, but open at each end). The full range of classical boundary conditions, which includes free, clamped, simply supported and shear diaphragm edges, may be applied in any combination to open and closed panels, thereby facilitating the study of a wide range of conical sandwich shells. The convergence properties of this element have been established for different combinations of the h- and p-parameters, thereby assuring its integrity for more general use. Since very little work has been reported on the vibration characteristic of either circumferentially closed or open conical sandwich panels, the main thrust of this work has been to present and validate an efficient modelling technique, rather than to perform numerous parameter and/or sensitivity studies. To this end, some new results are presented and subsequently validated using a commercially available finite element package. It is shown that for results of comparable accuracy, models constructed using the h-p formulation require significantly fewer degrees of freedom than those assembled using the commercial package. Some preliminary experimental results are also included for completeness. © 1999 Academic Press.
Copyright ©2014 by the International Astronautical Federation. All rights reserved.Microvibrations of a reaction wheel assembly are commonly investigated in either hard-mounted or coupled boundary conditions, although coupled wheel-structure disturbances are more representative than the hard-mounted disturbances. With the aim to reproduce the dynamics between a reaction wheel and its supporting structure, the dynamic mass (or its inverse, the accelerance) of the wheel and the driving point accelerance of the supporting structure have to be evaluated. This usually involves a series of experiments to characterise the hardware and produce exemplary models. Here a methodology is presented which has been shown to produce good estimates over a wide frequency range using a less complex test campaign. In addition, a practical example of coupling between a reaction wheel assembly and a structural panel, where the coupled loads have been estimated using the mathematical model and compared with experimental results, will be presented. Moreover, indications of the level of accuracy that can be expected from this type of analyses will be given herein.
Flight and ground segment software in university missions is often developed only after hardware has matured sufficiently towards flight configuration and also as bespoke codebases to address key subsystems in power, communications, attitude, and payload control with little commonality. This bespoke software process is often hardware specific, highly sequential, and costly in staff/monitory resources and, ultimately, development time. Within Surrey Space Centre (SSC), there are a number of satellite missions under development with similar delivery timelines that have overlapping requirements for the common tasks and additional payload handling. To address the needs of multiple missions with limited staff resources in a given delivery schedule, computing commonality for both flight and ground segment software is exploited by implementing a common set of flight tasks (or modules) which can be automatically generated into ground segment databases to deliver advanced debugging support during system end-to-end test (SEET) and operations. This paper focuses on the development, implementation, and testing of SSC’s common software framework on the Stellenbosch ADCS stack and OBC emulators for numerous missions including Alsat-1N, RemoveDebris, SME-SAT, and InflateSail. The framework uses a combination of open-source embedded and enterprise tools such as the FreeRTOS operating system coupled with rapid development templates used to auto-generate C and Python scripts offline from ‘message databases’. In the flight software, a ‘core’ packet router thread forwards messages between threads for inter process communication (IPC). On the ground, this is complemented with an auto-generated PostgreSQL database and web interface to test, log, and display results in the SSC satellite operations centre. Profiling is performed using FreeRTOS primitives to manage module behaviour, context, time and memory – especially important during integration. This new framework has allowed for flight and ground software to be developed in parallel across SSC’s current and future missions more efficiently, with fewer propagated errors, and increased consistency between the flight software, ground station and project documentation.
The Finite Element Method (FEM) has become the most utilized tool to carry out structural analysis. It is implemented in various software packages which are commonly used in Industry. FEA gives accurate predictions up until the first few structural modes of vibration, where the behaviour of real structures is quite deterministic. In the high frequency range statistical approaches are more suitable, and here Statistical Energy Analysis (SEA) has been applied quite successfully. In the mid-frequency range FEM predictions start to become unreliable, and SEA is not applicable as some of its basic assumptions are not verified. This paper has been developed in the context of a project concerning analyses of transmission of micro-vibrations in satellite structures. In addition to the ones related to the mid-frequency range, micro-vibrations introduce other issues: being very small entities, the related uncertainties are more substantial. Because of the large bandwidth of the frequency range related to micro-vibrations, their modelling and analysis pose a challenge, in particular in the mid-frequency range, where many of the micro-vibration sources on board a spacecraft tend to act. In this context, this paper will deal with two different aspects: on one side the development of a method aimed at reducing the computational effort nowadays involved to overtake the mid-frequency issue (we propose to merge the efficiency of the Craig-Bampton reduction with the simplicity and reliability of the Monte Carlo Simulation for the various subsystems to produce an overall analysis algorithm); on the other side the validation of finite element models of satellite structures (the effectiveness of Modal Assurance Criteria and Normalised Cross Orthogonality on the response prediction of spacecraft models is here carried and a new criterion, Base Force Assurance Criterion, is defined using the experimentally determined dynamic force at the base and the finite element predicted force). The method (Craig-Bampton Stochastic Method, CBSM) will be described in this article; with a benchmark example shown. A proof of the validity of the method in a real industrial application will also be presented, which will be performed by comparing the results obtained in applying the CBSM and the MCS to results obtained during an experimental campaign. This campaign has been carried out on the spacecraft SSTL 300 (made available by the company Surrey Satellite Technologies Limited in Guildford, UK).
Electromagnetic dampers (EMD) have been widely studied and designed in the control of vibrating structures. Yet, their use for space applications has been almost negligible, due mainly to their high ratio of system mass over damping force produced. The development of shunted circuits, and in particular negative impedances, has allowed higher currents to flow in the device, thus obtaining an increased damping performance. However, the need for a thermal analysis has become crucial in order to evaluate the power and temperature limits of EMDs, and hence allow a more efficient optimization of the whole device. This paper presents a multiphysics Finite Element Analysis (FEA) of an EMD in which the thermal domain is integrated with the electromagnetic and mechanical domains. The influence of the temperature on the device parameters and overall performance in the operative temperature and frequency range of a space mission is shown. It follows a design optimization of an electromagnetic shunted damper for 5-kg SDOF to obtain a second-order filter. In particular, the analytical results are compared with the typical transfer function of a viscoelastic material. This paper demonstrates the feasibility to achieve the same slope of -40 dB/dec while considerably decreasing the magnitude of the characteristic resonance peak of viscoelastic materials. © (2015) COPYRIGHT Society of Photo-Optical Instrumentation Engineers (SPIE). Downloading of the abstract is permitted for personal use only.
In order to examine the dynamic response of spacecraft during launch, Coupled Loads Analyses (CLAs), which couple a Finite Element Model (FEM) of the spacecraft with a model of the launch vehicle, are performed to simulate critical flight events. For the CLA results to be trusted, it is necessary to first develop a high level of confidence in the spacecraft FEM. This confidence is achieved by conducting appropriate test-FEM correlation and update activities making use of data gathered during vibration testing of the physical hardware. One major point of concern is the containment of the correlation and update effort in terms of mode count/modal domain. As such, this work is concerned with the assessment of the effectiveness of various target mode selection criteria. Findings are presented for initial investigations conducted using FEM data for a large, unique, scientific spacecraft developed by the European Space Agency (ESA). The work presented herein is the initial stage, and a larger study would be required to draw conclusions on the most effective means of containing the modal domain for correlation and update activities to those natural frequencies/modes which are most likely to contribute significantly in response to flight event level loading conditions.
Taylor Benjamin, Massimiani C, Duke Richard, Stewart Brian, Fellowes Simon, Bridges Christopher, Aglietti Guglielmo (2017)The AlSat-1N CubeSat Mission, In: AAS Advances in the Astronautical Sciences163
The stringent stability requirements imposed by advanced, high-resolution payloads have produced an increased interest in the development of better-performing micro-vibration isolators. Several devices aimed at mitigating micro-vibrations have been studied and implemented, but their application is still far from being ideal due to the several drawbacks that they present, such as limited low-frequency attenuation for passive systems or high power consumption and reliability issues for active systems. This research focuses on the modelling and testing of Electromagnetic Shunt Dampers (EMSD) characterised by the use of negative impedance converter circuits. An electromagnetic damper is a self-excited device that exploits the interaction between a moving magnetic field and a conductive material to provide a reaction force to the applied motion. An EMSD presents several advantages, but the high ratio of system mass over damping force produced has limited its application in space missions. The use of a negative resistance can considerably lower this ratio since it produces an overall reduction of the circuit resistance that results in an increase of the induced current in the closed circuit and thus the damping performance. In this thesis, the development of a multiphysics, multi-parametric model of an EMSD is presented and accurately corroborated by an extensive test campaign. This damper can be classified as a semi-active damper since the negative resistance circuit does not require any control algorithm to operate. In terms of damping performance, this research demonstrates that an EMSD applied to a 1-DoF system is capable of behaving, throughout the whole temperature range of interest, like a 2nd-order mechanical filter in which the resonance peak is eliminated and the roll-off slope is -40 dB/dec. Additionally, the proposed EMSD is characterised by low required power, simplified electronics and small device mass that could allow it to be comfortably integrated on a satellite. This study presents also a possible novel 2-collinear-DoF system design with embedded EMSDs. This isolator is capable of achieving a remarkable final decay rate of -80 dB/dec while completely eliminating the two resonance peaks due to the high attenuation performance of the dampers. Moreover, other aspects of the proposed 2-collinear-DoF system are investigated in order to assess not only the damping performance but also its features at system level. This work demonstrates that the fundamental advantages of this system can make it a viable, competitive alternative to other actively controlled struts.
Mathematical finite element models (FEMs) of spacecraft are relied upon for the prediction of loads experienced during launch and flight events. It is essential that the spacecraft is able to survive the launch environment without sustaining damage which could inhibit its ability to carry out its mission. Therefore, ensuring that these FEMs give a realistic representation of the physical spacecraft structural dynamics is an important task. To achieve a high level of confidence in the FEM in question, a correlation activity is conducted. This is the process of applying various metrics to compare computational results, from analysis of the FEM, with corresponding data derived from measurements taken of the physical hardware during vibration testing. Subsequently, updates are applied to the FEM where necessary to achieve an acceptable level of correlation. It is possible for spacecraft FEM correlation exercises to take a considerable amount of time and effort without necessarily achieving an appreciable improvement in the final FEM. As such, this project has been conducted to address the need to ensure that the procedures being applied are as effective and efficient as possible. Various aspects of the spacecraft FEM correlation process have been investigated separately, and interactions between the different stages in the process have also been considered. Two large, unique, scientific spacecraft have been used as example applications in order to carry out these studies. As well as making use of computational results from the spacecraft FEMs, this project has also included comparisons to the results from the corresponding base-shake sine-sweep test campaigns conducted on these structures. A number of noteworthy, and industrially beneficial, findings relating to the effectiveness of the spacecraft FEM correlation process have resulted from these studies: the most appropriate techniques of modal parameter estimation for the considered spacecraft applications have been established; the potential benefits and relative merits of different pre-test sensor placement procedures have been explored; inaccuracies introduced through the use of a commonly applied FEM reduction method have been demonstrated and a superior alternative identified. In addition, the efficiency of the correlation and update process has also been addressed. This has mainly been achieved through investigations concerning the applicability of commonly used target mode selection criteria to spacecraft applications, and the potential benefits of a less widely applied method which takes into consideration the expected loading scenarios to be experienced by the considered structures.
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.
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.
Nowadays, a technology demonstrator platform popular amongst the research community given their relatively low cost and short development time are cubesats. Nevertheless, cubesats are by definition nano-satellites of small volume and mass, and therefore, they traditionally only allowed very limited sizes of any expandable structure onboard with final deployed areas in the order of a few square meters. This conflicts with the large areas required for efficient solar sails, making the demonstration of this exotic concept bound to more expensive missions with a dedicated launch. The applications that will be discussed throughout the thesis will be: three-axis stabilised solar sailing with a "rigid" support structure; and drag assisted deorbiting of a large host craft using a solar sail. Both of these applications still need validation in space, especially for Earth-bound missions. The main goal of this research effort is thus to satisfy the need of available deployable booms for their use on systems of unprecedented mass per unit area with cubesat-like mission constraints that will ultimately place more trust in gossamer concepts. For this, two novel rollable booms and their deployment mechanisms have been developed, one based on metallic tape-springs and the other on bistable composite slit tubes. Analyses and tests confirmed that the former boom has scalability problems related to stowage-induced boom axial curvature, and coil blossoming management. Reliable sail deployments of a 4 x 4 m^2 sail were achieved with them. The latter boom design solves previous scalability problems of bistable composite booms. The ground demonstrator tested deploys reliably a 5 x 5 m^2 sail, with the current compact boom design shown to be efficiently scalable for 100 m^2 class sails. To enable even larger sails with the bistable booms developed, a novel architecture named the completely stripped solar sail has been proposed. A simple experiment demonstrated the beneficial effect that dividing the sail into sets of parallel strips and using a continuous sail-boom attachment suspension configuration has towards scalability of the concept. A new structural characterisation programme developed means by which to characterise the slender booms properties. In addition, the test results validated and/or updated the imperfection seeded finite element models produced. These models are ultimately utilised in high-fidelity predictions of the performance of the solar sail booms under the established operational loads, as well as in the scalability analyses of the sail concepts proposed with them. Lastly, the first gossamer sail-based deorbiting system in it class, developed for medium mass (< 1000 kg) objects in Low Earth Orbit under an ESA contract, is introduced. Mission requirements, designs, and the purposely developed qualification programme are shown for the final system that reached TRL 5-6. The challenges and lessons learned from the ground testing of such lightweight structures are also documented with the aim of assisting future design and development efforts of similar concepts.
Coiled deployable booms have been used extensively in space and are a large part of the deployable space structures family. They have a wide variety of uses such as the deployment of instruments, gravity-gradient stabilisation masses and more recently solar sails. Most deployable booms are similar to a carpenter's tape measure in the way they are coiled in a retracted condition and then deploy to form the boom structure. There have been many developments in the optimisation of boom properties in the deployed state, by using different shape cross sections and by using different materials. The first metal tape spring booms have developed into the more modern booms with a variety of cross sections. One aspect that is common to all booms is the coiling and uncoiling process and the difficulties associated with this. Blossoming, where the boom starts to uncoil within the boom deployer, can lead to the jamming of the mechanism. The reasons behind blossoming have not been thoroughly investigated, leaving designers of booms, and boom housing mechanisms to try and mitigate this problem themselves, often by trial and error. This work investigates boom blossoming with the aim of better understanding the underlying mechanics so that more effective deployment systems can be designed in the future. A method is developed that uses the strain energy stored in coiled booms to find the maximum tip force that can be achieved before blossoming occurs. This method is also used to investigate the central spindle torque during blossoming. The effects that the coil geometry and the friction between the layers of the coiled booms have on blossoming are also investigated. The theory developed should enable the designers of tape spring deployers to estimate the tip force and central spindle torque of a tape spring boom in the design phase of projects and reduce the reliance on trial and improvement type testing once deployers have already been built.
In the past 50 years the scientists have been developing and analysing methods and new algorithms that optimise an interplanetary trajectory according to one or more objectives. Within this field, in 1963 Lawden derived, from Pontryagin's minimum principle, the so-called `primer vector theory'. The main goal of this thesis is to develop a theoretical understanding of Lawden's theory, getting an insight into the optimality of a trajectory when mid-course corrections need to be applied. The novelty of the research is represented by a different approach to the primer vector theory, which simplifies the structure of the problem.
Accessing the subsurface of planetary bodies with drilling systems is vital for furthering our understanding of the solar system and in the search for life and volatiles. The extremely stringent mass and sizing mission constraints have led to the examination of novel low-mass drilling techniques. One such system is the Dual-Reciprocating Drill (DRD), inspired by the ovipositor of the sirex noctilio, which uses the reciprocation of two halves lined with backwards-facing teeth to engage with and grip the surrounding substrate. For the DRD to become a viable alternative technique, further work is required to expand its testing, improve its efficiency and evolve it from the current proof-of-concept to a system prototype. To do this, three areas of research were identified. This involved examining how the drill head design affects the drilling depth, exploring the effects of ice content in regolith on its properties and drilling performance, and determining the benefits of additional controlled lateral motions in an integrated actuation mechanism. The tests performed in this research revealed that the cross-sectional area of the drill head was by far the most significant geometrical parameter with regards to drilling performance, while the teeth shape had a negligible effect. An ice content of 5 ± 1% in the regolith corresponded to an increase in drilling time and a clear change in the regolith's physical properties. Finally, it was demonstrated that the addition of lateral motions allowed the drill to achieve greater depths. This work has advanced both the understanding and design of the DRD considerably. It has continued the exploration of the geometrical and substrate parameters that affect drilling performance and provided the first characterisation of the properties of an icy lunar polar simulant. The construction and testing of the complex motion internal actuation mechanism has both evolved the DRD design and opened a new avenue through which the system can be further optimised.
Spin-stabilised control was a pioneer spacecraft control method before a three-axis stabilised control was widely used. Spin-stabilised spacecraft, or spinner, uses this simple but robust method in controlling or manoeuvring the body in the microgravity environment. Surrey Space Centre (SSC) is currently spearheading the research in this area for a prolate shape spinner as a result of its involvement in the Moon Lightweight Interior and Telecoms Experiment (LITE) surface penetrator mission development funded by Astrium. The penetrator is a missile-shaped spacecraft used for delivering required mission equipment to the subsurface of the intended planet or celestial body. It is spin-stabilised after release from the orbiter before it is slewed to achieve a desired angle while free falling to the surface. The work described in the thesis is the ongoing development of the slew control algorithms by SSC as mentioned above. State-of-the-art algorithms have been developed, namely the Half-Cone (HC) derived family and pulse-train family. These algorithms have been proven theoretically, but implementation in a real-time mission is yet to be done, except for the Rhumb Line slew control. In particular, this thesis addresses the issue of the asymmetric shape of the developed prolate spacecraft where, in theory, it has been assumed as perfectly symmetric. Three new algorithms based on the HC-derived family that consider this asymmetric factor are discussed. The significant improvement made by these novel algorithms is the mass reduction of the final residual nutation in an average of one tenth of the current algorithm. Further analysis is done to these new algorithms in terms of accuracy, energy efficiency, slew time, the effect of thruster response time and gravity. The performance of these new algorithms in controlling an extreme asymmetric case is described first before it is applied to a common prolate-shaped spacecraft. An attempt to develop a testbed is also discussed within the work.