Currently I carry out research in the area of microvibrations on board spacecraft, multifunctional spacecraft structures, correlation between FEM and testing, response of electronic packaging subjected to harsh vibration environment. My top priority will be to strengthen the collaboration between the Surrey Space Centre and SSTL to create new synergies and ensure a seamless relation between the research carried out in the Centre and the practical applications pursued by SSTL.
Further information on my activities is available on the Spacecraft Structures, Materials and Mechanisms page.
I am a Chartered Engineer and a Fellow of the Royal Aeronautical Society. I have active partnerships with industry and I have collaborated with all the most important UK companies operating in the space sector.
Find me on campus Room: BA U
© 2015 IAA. Published by Elsevier Ltd. All rights reserved.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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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
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.
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.
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.
© 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.
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 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 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.
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.
Due to their high specific strength and high specific stiffness properties the use of honeycomb panels is particularly attractive in spacecraft structures. However, the harsh environment produced during the launch of a satellite can subject the honeycomb cores of these sandwich structures to severe quasi-static and dynamic loads, potentially leading to static or early fatigue failures. Knowledge of the static and 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 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. As well as conducting tests for the more conventional principal cell orientations (L and W), results are also presented for tests carried out at intermediate orientations to investigate the variation of core shear strength with loading orientation. The results are further investigated using explicit non-linear finite element analysis to model the buckling failure mechanisms of the tested cores. © 2011 Springer Science+Business Media B.V.
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 resul
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.
An investigation on the structural performance of inserts within honeycomb sandwich panels is presented. The investigation 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 focuses on the static performance of the two insert systems subject to loads in the normal direction to the facing plane. 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. The experimental results were compared with results from an analytical model and results from a finite element model. Contrary to what was expected it was found from the experiments that the cold bonded inserts outperformed the hot bonded inserts in terms of load carrying capability. From the finite element study it was found that this was mainly due to the difference in stiffness of the different filler materials used in the two insert systems. © The Author(s), 2011.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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 spacecr
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.
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