Angelos Tsinas 15th Annual USU Conference on Small Satellites1

SSC01-IV-5

Efficient Satellite Structural Design Optimised for Volume Production

Angelos TsinasResearcher

Kingston UniversityFriars Avenue, Roehampton Vale

London UK, SW15 3DW0044-208-5472000 ext 4845

k949769@kingston.ac.uk

Dr Christopher WelchPrincipal Lecturer in Astronautics

School of Mechanical, Aeronautical and Production EngineeringKingston University

Tel: +44 181 547 8725 Fax: +44 181 547 7992C.S.Welch@kingston.ac.uk

Abstract:

In the past, the limited number and production volume of satellites has meant that their structural design hasbeen essentially a one-off procedure. As a result, the most common choice of primary structural medium hasbeen metal. Although eminently reliable and highly proven, this option has led to a comparatively highstructural mass fraction of 20-24%.

The emergence of communications satellite constellations creates the need for a complete reappraisal of currentdesign practices. Emphasis needs to be given in the implementation of volume production methods, alreadymatured through experience in the aviation industry, to manufacture of satellites. The prospect of new materialsand technology can offer reductions in the overall structural mass of satellites in the region of 15-20% which inturn can lead to significant overall mass savings and reduced launched costs. However the aim of massreduction can only be appreciated in terms of total cost savings, i.e. the net balance of the mass savings versusthe technological application cost should be positive.

This paper describes current approaches to the design of volume production satellites such as TechnologicalIslands, Virtual Factory, Multifunctional Surfaces, Short Accelerated Production of Satellites (SNAPSAT) inrespect of efficiency and economy. The implementation of volume production methods such as JIT and Taguchiin the area of satellite technology is also examined.

Alternative designs for mass production satellite structures are considered. Ideas described include aconventional truss of both composite and aluminium manufacture and a corrugated plate modelled upon themultifunctional surface.

Satellite StructuresManufacturing Issues

Past PracticesIn the past, satellite projects have been characterisedby a limited production runs. The bus structure usedto be mainly metallic, typically of a honeycomb ormonocoque thrust tube and surrounding shear panels.This in essence translated in an increased part count(especially in forms of connectors and fasteners),

resulting in a high cost and an increased percentageof the structural mass as part of the total mass. Thisweight allocation, although unavoidable, wasresponsible for increased launch costs, which weremore profound for GEO launches. Furthermore, evenfor quite successful designs such as the HS-601 GEOfamilies, the need for volume productionmethodology was offset by the large number ofdifferent projects. This resulted in a smallprocurement number per order so that the cost per busremained high.(1)

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Modern Constellations and the Need toDevise New Methods of Manufacture

Most modern satellite constellation designs feature adense network of LEO satellites in order to achievethe desired Earth coverage. The choice of LEO ismade on the grounds of a better signal quality withminimal signal latency, reduced launch costs and theneed for a more compact and economical satellite interms of size, mass and power allocation. Also theuse of new materials, culminating in the introductionof the composites and exotic alloys results inextremely high specific strength values which canoffer flexibility and, potentially, offer significant costsavings.

It is apparent that conventional manufacturingmethods can’t be implemented in such cases becauseof the increased manufacturing costs, pressing projectcompletion timescales, and unacceptably high launchcosts. The solution should come by mimickingsuccessful volume production methods similar tothese implemented in other high technologyindustries such as aeronautical, automotive andelectronics engineering. The difference, of course, isto be found in the absolute magnitude of the numbersproduced. However in aspects such as projectmanagement the situation is comparable. Theimplementation of modern production methods in thesatellite industry was a first step to a new directionand has not yet fully matured. The next sectiondescribes most of these methods and gives relevantexamples of current practices when applicable.

Volume Production Methods

Just In TimeJust In Time (JIT) is a systems approach todeveloping and operating a manufacturing system,based on the concept of waste elimination. (2) Itrequires that equipment, resources, and labour shouldbe readily available in the correct amounts required atany current moment. Its fundamental principle is thatonly the necessary units and the necessary quantitiesshould be manufactured. This is achieved by constantmonitoring of the production rate and adjustment forproduction needs. Besides waste elimination, thebenefits of JIT are increased productivity, workperformance and product quality while enabling lowcost.

Aspects of JIT include:• Integration and Optimisation• Quality Control• Reducing Manufacturing Cost• Producing Product on Demand• Developing Manufacturing Flexibility• Establishing Links with Customers and Suppliers

The main requirements to be fulfilled in order tosuccessfully implement JIT are:• Partnerships• Commitments• Contracts Supporting Partnerships• Developing JIT suppliers• Customer-Supplier proximity

Taguchi MethodsDr. Taguchi played an important role in shaping theattitude and philosophy of the Japanese industrial andmanufacturing strategy after the Second World War,which led to the economic growth and developmentof Japan. While Western attitude was focused onincreasing output rate, Taguchi introduced theconcept of Total Quality Management (TQM) inorder to maximise profits. The definition of quality isbased on whether the product is able to fulfil itsspecified mission at the least cost. (3)The space industry was late to adopt Taguchiprinciples, because of the low production numbers ofspacecraft families. However, the emergence ofcommunication satellite constellations projects hascreated the need for a different approach to the issueof manufacture of satellites. Indeed, both NASA andthe Department of Defence in the US, and theircounterparts in Europe and Japan have started toestablish TQM practices in their projects, focusing onthe relationship of performance with respect todevelopment cost. This attitude could be summarisedas “Cheaper, Smaller, Faster”, allowing the rapiddevelopment and deployment of new projects.The core of TQM is that quality should be achievedat the design level of the project. Once the productgoes into production, any modifications result inadditional cost and time delays. Achieving qualitycontrol is accomplished by any of the following threemeans:

1. System Design: System design ensures that theproduct is the built according to strictly laiddown specifications. Consequently, the outcomeis a project of increased reliability (but at ahigher cost, which may deter potential buyers).

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2. Parameter Design: In parameter design, thefactors that influence the performance aremonitored and weighted. As a result, the strongand weak points of the design are found andareas of specific consideration and interest arelocated. The ability of the product to resist theinfluence of the external parameters is defined as“robustness”, and the result as achieving a“robust design”. The comparison of theinfluencing parameters is accomplished using“orthogonal arrays”; a form of matrices, whichallow a quality score for every modification, thusallowing the selection of the most suitable

candidate. The external influence parameters aredefined as “noise”.

The process for a Taguchi optimised design has thefollowing steps:

• Determination of the Quality Characteristic to beOptimised

• Identify of the Noise Factors and Test Conditions

• Identify of the Control Factors and theirAlternative Levels

• Design the Matrix Experiment and establish theData Analysis Procedure

• Conduct Matrix Experiment

• Analyse Data and Determine Control Factors’Optimum Levels

• Predict Level’s Performance

3. Tolerance Design: Accepting wider tolerances indesign allows the manufacture of a cheaperproduct at the price of reduced reliability andvice versa. Concluding, it could be said thatTaguchi method dictates that quality controlefforts should begin at the early stage of designprocess, in order to conceive a configuration thatis less prone or more robust to externalinfluences, allowing for a reasonable cost.

Benefits of the Taguchi approach, therefore, includetime and resource savings, parameter influenceidentification in terms of operation performance andcost. This allows effective allocation of resources andtime, which translates as low cost, high qualitysolutions.

Virtual Factory

“Virtual factory” is a term coined by the Iridiumcompany in order to describe the operationalprocedures and methods that were applied in theIridium project.

The definition of virtual factory is: “ A team ofmultiple companies, each having a distinct worldclass core competence, that are leveraged in apartnership through collaboration and teamwork,resulting in a distinct competitive advantage.” (4)

The main issues that a virtual factory approach has toaddress are the compatibility and commonality of

metrics and procedures between vendors andsubcontractors; the establishment of good relationshipbased on trust, a well-established transportation anddelivery system etc. Proximity of the installations isan important parameter, but can be compensated bythe establishment of effective transportation and theassurance of total quality before the ordered productleaves the “sub-factory” facilities.

Indeed, although Iridium’s components came from avariety of sources in terms of origin and location, allthe main assembly and testing was carried out atMotorola’s facilities in Sunnyvale, CA. The level oftesting was limited to the subsystem/ completesatellite level, as all component testing is theresponsibility of the subcontractor or vendor.

Systems established during the implementation of the“virtual factory” approach included Six SigmaQuality, Design for Manufacturability, and DiscreteEvent Simulation. During the Process Developmentand Verification phase, the Iridium partnershipestablished a Quality System Review in order toassess and control the overall quality andperformance of the system. Quality was achieved andmonitored through a set of processes includingDesign of Experiments, Statistical Process Controletc.

Technological Islands

The introduction of the LEO communication satelliteconstellation concept has significantly altered aspectsof satellite manufacture. In particular the increase inproduction demand in a relatively short time scale hasproved that the traditional linear manufacturingapproach is no longer in position to comply with therequirements.

As a result, a new approach to Assembly, Integrationand Test has had to be sought.For Globalstar, Alenia’s response to this is based onthe concept of technological islands, which is definedas “Areas in the factory where a series of

Angelos Tsinas 15th Annual USU Conference on Small Satellites4

homogeneous activities are performed in a well-established and controlled fashion.” (5)

At Alenia, two separate production lines wereestablished for the Globalstar production, with eightislands each dedicated for spacecraft assembly andantenna manufacturing and integration.

Satellite Islands

• Propulsion Assembly Integration and Testing• Bus and Payload Integration• Alignment verification• Satellite performance test and thermal cycling• Vibration test and solar array installation and

deployment• Mass properties• Packing and shipping• Troubleshooting

Antenna Islands

• Antenna pre-assembly• Hard line test• Final integration• Thermal cycling• Vibration testing• Near field test range• Far field test range• Troubleshooting

The eight production islands are arranged in theAlenia’s Small Satellite Centre manufacturing sitebased on the following parameters.

• The spacecraft is the only item that moves alongthe production process.

• No equipment transfer from one island to anotheris necessary.

• No personnel transfer from one island to anotheris necessary.

• No backward movement along the productionprocess is necessary.

The use of technological islands requires the need foran extensive and reliable data management system.This allows the support and monitoring of allproduction activities, while it is integrated withtesting, documentation and resource managementsoftware for complete process control.

The whole manufacturing process can be monitoredat any time and any bottlenecks can be quicklyinspected and corrected. The use of integrated

simulation and production flow methods, playeddecisive role in allocating resources and finance aswell as establishing the manufacturing practices ineach stage of the production and formulating anoptimal process and method of work for each islandallowing high efficiency levels.

Each island is responsible for its own production andquality control of the finished product, while theparallel method of manufacture ensures thatproduction flow is kept to the optimal rate throughthe use of JIT methods. The autonomy of the conceptallows for reduction of time delays or disruptions,while any source of error can be discovered andeliminated, if not in the early stages, then before itcan infiltrate the production chain.

The factory layout is shown in Figure 1. TheTechnological Islands Conceptual Organisation andFlow is shown in Figure 2

Figure -1 Technological Islands Layout. (5)

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Figure 2 Alenia Factory Island Interactions (5)

ISLAND 4

PerformanceTesting

PerformanceTesting

Thermal Testing

ISLAND 7

Final VerificationAnd

Shipment to Launch Site

ISLAND 2

Bus Off LineIntegration

Bus Integration

Payload Integration

ISLAND 3

Alignment Verifications

ISLAND 1

Propulsion OffLine

Propulsion A.I.T

ISLAND 5

Solar ArrayRelease

Vibration Testing Solar ArrayInstallation

ISLAND 6

MassProperties

Verification

ISLAND 8

Troubleshooting

Angelos Tsinas 15th Annual USU Conference on Small Satellites6

In order to control the production schedule Alenia hasimplemented a Data Management System (DMS),which assures compatibility and ease ofcommunications between Alenia and the otherGlobalstar, partners. DMS is an extensiveInformation System which enables the Small SatelliteCentre to automatically and in real time receive,categorise, store, retrieve, process, modify anddistribute data. Data that may be processed includesdrawings, parts lists, material lists, processes andprocedures, specifications and results etc. DMSfacilities exist for every island, allowing rapidcommunication, and data exchange within the limitsof the factory and external contractors.

SNAPSATSNAPSAT (Short Notice Accelerated Production forSatellites) is a relatively new technique devised byCOI (Composite Optics Incorporated), for themanufacture of the FORTE satellite. (6) Althoughlimited data has been published, it has emerged that aset of disciplines has been devised and implementedthroughout the process, emphasising simplicity andlow cost development.

The main structural components are produced from asingle laminate plate although the same plate maycontain different parts. No mould process is required,and components are cut to appropriate shape by theuse of waterjet, or potentially laser. Both optionshave potential benefits and drawbacks.

The implementation of computer-controlled processfor the design and manufacture allows a uniformquality output that can be easily controlled andadjusted according to the incoming needs. Toolingminimisation yields low cost due to the reduction inthe non-recurring costs, less testing needs due toreduced part variation and part count, and reduceddevelopment time due to automation.

Group Part Process (GPP) is an important parameterin the SNAPSAT method. GPP allows thesimultaneous treatment of different shaped partsoriginating from the same initial plate. Furtherprocesses include the adaptation of advancedadhesive dispensing methods allowing reliable andsimultaneous bond cure.

Grastataro (7), indicates that the implementation ofSNAPSAT principles during the design of FORTEsatellite allowed the manufacture of a bus structure at40% of the cost of an equivalent conventionalcomposite structure, and 25% more expensive than aequivalent metallic structure. Krumweide (8)estimates the actual price at $160,000, $400,000 and$133,000 respectively. Respective weights are 42.6

kg, 42.6 kg and 64.4 kg. Cost in terms of $/kg are3750, 9400, 2060 respectively. Another importantissue addressed is that of manufacturing lead times,which stand at 10 weeks for FORTE, 16 weeks forthe aluminium structure and 30 weeks for theconventional composite structure.

The benefits behind SNAPSAT are based on thesimplicity of the manufacture of the structuralcomponents and the reduction in manufacturing leadtimes and overall weight savings, which result insignificant cost savings. The potential of theapplication of the SNAPSAT technology to themanufacture of the next generation of volumeproduced spacecraft is great, although COI had noplans of initiating a production run.

The shape restriction imposed by the single flat-stockpanels could be a potential disadvantage, but it couldbe argued that most of contemporary volumeproduced satellites exhibit a simple structuralconfiguration based on reducing part count andcomplex geometry. On both accounts, the applicationof SNAPSAT is compatible to the desired properties.The possible drawbacks of this process are shaperestrictions. Only flat surfaces are possible, duringmachining and cutting delamination of the compositelayers may occur as a result of improper waterjet orlaser beam use. The cutting area may be affectedeither by the water moisture of the waterjet or theheat of the laser as well as by improper drill speeds.Another disadvantage is an excessive amount ofwaste material as a result of the cutting technique;therefore, there is comparatively low ratio ofutilised/raw material.

A single plate provides less material cost in terms ofmanufacture or acquisition as well as testing. Wholestructural components are made as one-piececomponents, permitting fewer and well-controlledcomponent parts. Mass production quality isguaranteed through repeatability and fast productionrate is achieved. Tooling costs are kept to aminimum, as there is no need for unique moulds andequipment is needed, in fact, the main processing toolis a cutting device, either laser or waterjet to a simpleGerber knife and drills.

However due to the fact that the price of the materialin plain plate form is low compared with the price todevelop a fully processed component by followingmore efficient and thus complex methodscompensates for material loss. Single plate cutpermits the manufacture of whole structuralcomponents as one piece, thus saving tooling costsand processing time but also is damage prone in thesense that damage in one member may lead to

Angelos Tsinas 15th Annual USU Conference on Small Satellites7

scrapping of the whole component.

The benefits of a single plate make it quite attractivefor volume production of satellites and it is one of thestrongest candidates. The fact that in a truss shapestructure load path analysis is quite simple andstraightforward reduces the need for additional loadtests if the truss has been used before. The problem isthen changed if the material can withstand theseloads, which is simply as iteration/optimisationprocess using a typical FEA package.

Structural Options

TrussA typical truss consists of a large number of differentstructural members. It is of rectangular, triangular orpolygon shape, although the efficiency of thestructure drops as the number of corners increases.For many years, the truss has featured very heavily insatellite bus design.

A typical truss is constructed from extruded tubes ofvariable or uniform cross-section, usually square orcylindrical or separately machined members of anopen section such as I, Z shapes. The main materialsused for the construction of truss members are lightmetal alloys, primarily aluminium alloys orcomposites which combine the need of high stiffnessand low density.

The main attachment methods are fasteners, welding,and adhesives. If composite material is to be used formembers then the use of metallic, mainly titanium,inserts/ end-fittings is required. Joints in the form offasteners result in an increase of total structuralweight, large part count, and high cost in terms oflabour and manufacturing time. It is the aim of thedesigner, therefore, to find ways of minimising partcount and simplifying the manufacturing process byutilising uniform or standard components. Quiteoften, the truss members are manufactured as a singleblock in order to minimise the total part count.

The truss method is well-established practice in thespace industry, and carries much aircraft designheritage, although it has evolved to its own specificstandards and practices. The truss option was selectedfor all major constellation satellite systems, due to itssimplicity in manufacture and straightforward loadpath analysis and testing.

HoneycombHoneycomb is a well-established material in theaerospace industry. It offers great advantages in

weight savings, thermal properties, load distributionand high specific strength.

Applications include:• Panels either for component installation or solar

arrays.

• Thrust tubes, which act as the main load bearingstructure in many satellites. In this case thehoneycomb is attached to a combination oflongerons and stringers, which forms a frame.The resulting structure is called a monocoque.

Corrugated SheetsCorrugated sheets are another option for a sandwichstructure. In general, the corrugated structure consistsof thin isotropic facings, which have negligibleflexural rigidity about their own centroidal axes and ahighly orthotropic core.

Assumptions made for the cores are:

• Core shear modulus is far greater in the planelongitudinal to the corrugations than transverseto them.

• The core is orthotropic.• Bending rigidity for the core is negligible in the

transverse direction.• Shear distortions are admissible only in the plane

transverse to the corrugations.

IsogridIsogrid is a pattern of equilateral triangles integrallymachined from a flat surface, with or withoutmaterial left between the ribs to act as skin. It isisotropic in nature for in-plane loading. Isogridstructures are lightweight and offer high specificstrength, bending and buckling strength.

However, isogrid has not been extensively used as aprimary structure. Its application is more as acomponent-mounting platform since the hardpointsavailable offer a grid platform where components canbe arranged.

Such example of isogrid utilisation appears in theHS-601 bus.

The use of isogrid as main structural form wasdemonstrated in an experimental satellite projectcalled ISOSAT of the Texas University. (8) HoweverISOSAT is a nanosatellite class therefore no safeassumptions can be made for the isogrid suitabilityfor larger satellites.

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Structural Materials

IntroductionIn the past decade the choice for the primary structurematerials has shifted from almost exclusively metallicstructures to full composite utilisation. The metalsused were primarily aluminium with steel; titanium ormagnesium used at high stress points. Although thepractice was well-established and the properties ofthe materials well known, it had the drawback of highdead weight. However since the numbers of satelliteswere quite low and the production rate limited theissue of weight was not considered of paramountimportance.

With the advent of satellite constellations, though,(where the needs of a production rate reached manydozens), the issue of mass reduction was highlighted.The natural choice was the use of composites.Composite technology had reached an acceptablematurity and therefore it was suited for spacequalified applications.

Consequently, the modern designer is blessed with awide choice of spacecraft materials for any givenapplication, although the criteria of choice are stilldetermined by the operational characteristics andbudget. The main considerations focused on are:

• Stiffness (deflection levels)• Strength (Stress level limit)• Mass density (mass determination)• Thermal properties (Expansion, Conductivity)• Outgassing (material deterioration)• Corrosion (material deterioration)• UV-resistance (material deterioration)• Creep resistance (integrity)• Availability (choice)• Legality (choice)• Cost (budget)• Development time (timescale)• Tooling and plant development

(budget/timescale)

MetalsMetals were the first materials to be used in theaerospace industry and their use in the satellitemanufacture was a natural outcome of this. In termsof engineering, metals have an affordable cost foralmost all applications, well establishedmanufacturing and processing methods, knownproperties and behaviour which, when coupled withtheir isotropic properties, make stress analysis andsizing calculations a relatively simple task.

Spacecraft metallic structures are developed usingeither the prototype or the protoflight approach. Theminimum design and test factors of safety formetallic structures, excluding fasteners, are specifiedin the following table.

Table 1 Metal Structures Factors (9)

Ver

ific

atio

n

App

roac

h

Ulti

mat

e

Des

ign

Fact

or

Yie

ld D

esig

n

Fact

or

Qua

lific

atio

n

Tes

t Fac

tor

Acc

epta

nce

or P

roof

Tes

t

Fact

or

Prototype 1.4 1.0* 1.4 NA or

1.05**

Protoflight 1.4 1.25 NA 1.2

NOTES:

* Structure must be assessed to prevent detrimental yielding

during flight, acceptance, or proof testing.

** Propellant tanks and solid rocket motor cases only.

CompositesComposites are unique materials, combining two ormore materials, to utilise the respective advantages ofthe participating materials. The two main parts are thematrix and the fibre or filament. The fibre is of highstrength, high modulus, and low density. The mostcommon are boron, boric (silicon carbide coatedboron), and graphite. Boron filaments aremanufactured by vapour deposition of boron on a finetungsten wire. Graphite filaments are made bygraphitising tows or bundles of organic filaments.The most common matrix materials are epoxy resinand aluminium.

Although they offer many advantages over themetals, especially low density and high strength, theirwidespread use is held back by the complexmanufacturing procedures, which lead to lengthy andcostly development time.

In addition, the anisotropic nature of the materialrequires special care when load paths are calculatedand special methods of attachment have to be utilised(adhesives or fasteners). However this disadvantagecan be reduced if a modular construction approach isimplemented and advanced manufacturing methodsare applied. Furthermore, the overall manufacturingcost increase is a relatively small fraction of the totaldevelopment cost and the cost increase may be offset

Angelos Tsinas 15th Annual USU Conference on Small Satellites9

by the launch fee reduction due to smaller mass or theextra payload weight allocation which is freed.

Composite/bonded structures, excluding glass,developed for NASA spaceflight missions shall, as aminimum, use the design and test factors specified inthe following table.

Table 2 Composite Structures Factors (9)

Ver

ific

atio

nA

ppro

ach

Geo

met

ry o

fSt

ruct

ure

Ulti

mat

eD

esig

n Fa

ctor

Qua

lific

atio

nT

est F

acto

r

Acc

epta

nce

or P

roof

Tes

tFa

ctor

Prototype Discontinuities 2.0* 1.4 1.05

UniformMaterial

1.4 1.4 1.05

Protoflight Discontinuities 2.0* NA 1.2

UniformMaterial

1.5 NA 1.2

NOTE:

Factor applies to concentrated stresses. For non-safety criticalapplications, this factor may be reduced to 1.4 for prototypestructures and 1.5 for protoflight structures.

Material Choice for Structural Components1. Frame members: The frame consists of struts and

tubes, which are generally designed for buckling.Mass reduction can be achieved by usingberyllium, boron/epoxy, boron/aluminium, andgraphite epoxy. However, due to themanufacturing difficulties of the boron, specialdiamond based cutting tools, graphite epoxy ismore common.

2. Panels: The panels support the subsystemcomponents and enclose the structure. The mostcommon design option is aluminium corehoneycomb structure, with either metallic orcomposite facesheets. Composite facesheetshowever lack the electric and thermalconductivity of the metals; therefore, metallicinserts are used to compensate for the transferpaths.

3. Thrust cones: The thrust cone is the main load-bearing component of the spacecraft, whichutilise one. It is designed for axial compressiveloads and bending moments. The main failure

mode is shell buckling. There is a requirementfor lightweight high modulus materials, thereforeberyllium and advanced composites are used inthis section. Thrust cones can be eithermonocoque, semi-monocoque or sandwichdesign. The two latter produce more complexdesigns but are more lightweight.

The use of composite materials in satellite structureswas initiated in the design of GEO satellites due tothe significant savings in weight growth and thereforethe respective reduction in launch fees. However, theutilisation of composite materials for LEO satelliteswas restricted until recently despite the benefitsderived from the weight savings. The reasons werethat a series of considerations had to be addressed,most notable being the increased temperatures due tothe increase of the power output of the payload, andthe limited heat dissipation capability. Weight/volume and power restrictions constitute theutilisation of active heat systems a non-viable or non-attractive solution; therefore, the use of passivesystems is almost the standard. EMI shieldingavailable due to the nature of the composites wasquite low and the need for increased protection meantthat additional metal shielding was needed or thedevelopment of special Radiation Hardenedelectronics resulting in prohibitive cost levels.

The emergence of the LEO constellations howeverinitiated an extensive study at the behaviour of thecomposites, and efforts were made to enhance theirproperties. The need for a significant number ofsatellites allowed the absorption of the cost over alarge production run therefore limiting the unitResearch /Development cost.

Composite construction offers significant benefits interms of weight and cost compared to a relativemetallic structure. The total weight of the structure isin inverse analogy of the material’s stiffness i.e.K1100 < P 120 < M60J < M40J < 6061 Al.

Even with the requirement of EMI shielding theresulted weight penalty does not significantly affectthe weight. The procurement cost of the compositematerials in terms of (£ / Kg) is significantly higherthan the metal alloys utilised in satellite structures.However material procurement cost is comparativelylow with respect to the processing and machiningcosts where the composites have a significantadvantage, especially if lean production approachesto manufacture are implemented. The relationshipbetween material procurement cost, manufacturingand engineering cost is shown in the following figure.

Angelos Tsinas 15th Annual USU Conference on Small Satellites10

0

5

10

15

20

25

Alu

min

ium

6061 GR

/EP

rela

tive

co

st£/

kg

engineering

manufacture

material

Graph 1 Relative Cost of Materials

The overall composite cost breakdown can be shownin the following graph, which shows the part cost as apercentage of the total cost. It is interesting to notethat the purchase of the composites as raw material isa very small part of the total cost. However due to thenature of the composite properties, the amount oftesting / inspection required is a very significant partof the total cost.

Fabrication (27.00%)

Inspection (27.00%)

Assembly (28.00%)

Tooling (7.00%)

Design (8.00%)Materials (3.00%)

Composite Utilisation Costs

Graph 2 Total Composite Cost

The structure’s costs may be controlled further if thefollowing considerations are carefully considered andapplied.

• Design for reduction of part count.

One piece modular components, provide thekey answer to reduce weight and cost, due to the lackfor connectors and the increased toughness associatedwith a one piece component as no stressconcentrations occur which may otherwise have beena point of concern. The limitations in the size of themanufacture of one-piece components lie with thesize limit of curing ovens or mould considerations.Other point of concern could be the potential loss ofthe whole component if catastrophic damage hasoccurred in a limited region of the component.However the rejection rate can be minimised ifcareful processing procedures are implemented,furthermore the actual cost savings may constitute asmall amount of rejections as acceptable.

Co-cured structures, permit the simultaneoustreatment of the components, thus the resultingstructure has similar properties throughout. Time andenergy savings are possible thus resulting in reducedlabour and overhead cost.

• Design for reduction of labour hours.Batch processing.Reduction of ply count triax or lightly filled fabrics.

• Design for automation.Automated process permit an abrupt productionoutput in terms of unit rate and quality as theoccurrence

• Design for Integrated Assembly.The benefits of integrated assembly are mainly sizeand weight reductions, which result in reduced launchcosts. The implemented technologies may requiresubstantial investment and in terms of capital aresignificantly higher than the conventionalmethodologies, however this trend is about to changeas technology in this field become maturer. Anexample of integrated assembly is the multifunctionalstructures, which are reviewed in detail at the chapterof novel technologies.

• Design for low cost tooling approach.Tooling utilisation covers a substantial percentage ofthe non-recurring costs during the development of aproject. Tooling investment should be judged uponactual amount spent and production output in termsof cost or unit production cost. The manufacture ofvolume production satellites has enabled theutilisation of manufacturing techniques which were

Angelos Tsinas 15th Annual USU Conference on Small Satellites11

non-profitable for low unit output projects, especiallyat the composite material manufacturing andprocessing level. However this attitude was to betaken a few steps further when it was realised thatfurther cost reductions are needed in order to achievea high yield return of the investment.

C.O.I utilised the SNAPSAT technologywhich requires the manufacture of simple flat plateswhich are simultaneously treated and cured, appliesWatergate methods in order to cut the pieces in thedesired shape and titanium endfittings to connect thestructural elements together. This approach enablesthe manufacture of the structural components as onepiece, thus reducing the need for a large amount offasteners or end fittings resulting in reduced weight,requires a limited amount of stock resources, notablytwo different thickness panels and limited tooling interms of moulds. Iridium utilises a limited part countapproach that results in a limited amount of differentmoulds, thus reducing the overall cost. Thesimultaneous treatment of the structure during curingprovides additional cost savings.

Globalstar and ORBCOMM have decided tosolve the problem of low cost tooling andmanufacture by opting for aluminium structure. Thetooling requirements are metal treatment machines inorder to extrude and give the aluminium bars thedesired shape, drills, and cutters. The economicpotential is based on the relatively better knowledgeof metalworking combined with the lower materialacquisition costs.

Manufacturing costs may be contained if theoption of a dedicated subcontractor is preferred, insome cases the subcontractor is willing to invest onthe new technologies that are needed The benefits ofintegrating a dedicated partner in the project besidethe cost savings are the potential headstart in case ofthe use of patented and propertiary technology, andthe reduced risk involved as each part isconcentrating on the field of its expertise. Thepotential drawbacks of the subcontracting are that theproject potential is influenced by the ability of thesubcontractor to honour the agreement in terms ofschedule and quality. Although clauses for delays orpoor quality should be catered for, it is quite possiblethat any delays could be rather fatal for the viabilityof the project.

The mains cost drivers are:

1. Time: Time parameters include factors like leadtime, manufacturing time, processing time and totaltime. If time is a luxurious commodity, it is likelythat the occurring costs will be significant.

In cases where time is the driving parameter it iswiser to opt for proven technologies of low risk andlow cost, compared with more modern options,although it is quite possible that a loss ineffectiveness and reliability is statistically increased.

2. Mission requirements: Mission requirement is avery broad concept, as it can include the descriptionof the mission scope in both quantitative andqualitative terms. Between those two, there is afundamental difference, which must be wellunderstood and accepted before the finalisation of anindividual mission, or project may commence. To usean example from the satellite industry; the scope of acommunications satellite is to deliver and receivesignals of a certain frequency range so to permitcommunication and data exchange. However, thechoice of the frequency range, number of channels,quality of reception etc is a different issue. It can beargued that the choice of these parameters willdirectly influence the design of the satellite systemsbut still the dilemma is faced once the initialrequirements are set. Cost considerations, ortechnology restrictions will compromise the qualityor the capability of the potential design, so it is thetask of the project management team to come up withthe most effective option available.

3. Process development: It includes all the necessarysteps that are needed to be undertaken in order toachieve the final product. Therefore, processdevelopment includes both non-recurring andrecurring costs.• Raw material and parts procurement• Equipment and factory overheads• Labour costs:• Risk.• Testing and Inspection:• Specific/ Interrelated

Finite Element Analysis

Introduction to Finite Element Analysis (FEA)Finite element analysis is achieved by dividing thestructure under examination into small units called“elements”. The element can be given various shapes,although rectangles and triangles are by far the mostcommon. The elements are connected through the“node” points.

The accuracy of the calculations therefore dependsheavily on the quality of the modelling of thestructure, any assumptions made for boundaryconditions, the level of knowledge regarding loadingconditions and environment, and of course thenumber and the kind of elements used to represent the

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structure. Although the finer the quality of theelements, the more accurate the calculation gets, thishas its toll on computation time, which translates intoincrease of costs and extension of the projectedtimetable, for this reason the model should be only asaccurate as needed for the task in hand. The finiteelement analysis involves three major processes.

• Pre-processing• Solution• Post-Processing

FEA ModelThe bus to be modelled was chosen to be the LoralLS-400, as used for the Globalstar project. Sincecontact efforts to provide with accurate data failed,the structure was modelled using all availableinformation and extrapolating using appropriateequations in order to solve for unknown properties.The choice of LS-400 was based on the fact that it isa simple structure to model for FEA, thereforereducing the model complexity. Component layout iswell documented and available. Finally the samestructure was short-listed by NASA in Phase 1 oftheir Rapid Acquisition Plan, which is a goodmeasure of its engineering merit. (10) The structureof the LS-400 / Globalstar is shown in figure 3.

The satellite's trapezoidal shape, fabricated of a rigidaluminium honeycomb, is designed to conservevolume and facilitate the mounting of multiplesatellites within the fairing of a space launch vehicle.Separation of the satellite is achieved by explosivemechanisms, which free the satellite away from acentral core dispenser on the launch vehicle. Whenthe satellite is mounted within the fairing of thelaunch vehicle, the Earth face is oriented outward,and the anti-Earth face is mated to the control, andcommunications subsystems. The actual Globalstarsatellite is strongly based on the Loral LS-400 seriesand is manufactured under license by Alenia Spazzioof Italy and a series of subcontractors includingDASA and Alcatel. At the centre of the satellite ahoneycomb platform houses most of the onboardsystems while the rest are mounted to the otherpanels. However for reasons of modularity the Earthfacing panel which houses the main communicationantenna is manufactured separately and assembled ata later stage.

Figure 3 LS-400 Bus Structure

MethodologyBefore the actual finite element analysis, a proceduremust be followed. The first step is the analysis of theproblem, that is the definition of the objectives of theanalysis, definition of the environment and forcesacting or the boundary conditions and setting theappropriate number of constraints including cost andavailable hardware and software. The desired outputof the analysis was assessed, that is the format anddetail of presentation of the results of the analysis.Once the requirements are set, the level of detail forthe model can be determined.

All steps and assumptions during the solution processwere checked and monitored, clearly documented.This can easily trace way the possible mistakes andcorrected, thus saving time and cost. On the otherhand though, excessive documentation should also beavoided, especially at the preliminary design stages,as they consume precious time and resources andgive an unnecessary level of detail.

The output of the analysis was then checked. Usingcommon sense, it is possible to identify the areas oferrors and the reason. A second analysis, using adifferent technique, will in most cases either verifythe validity of the results or give directions to correctthe errors. This fact is an inherent flaw of the finite

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element methodology; it requires the appropriateinput information and instructions in order to processthe data and derive the correct solution. The effect ofthe initial assumptions should be checked. In somecases assumptions have, no detrimental effect on theoutput accuracy, but the over-simplification of themodel could generate a significant percentage oferror.

It should be noted that due to the sensitive nature ofthe space industry it is impossible to obtain accuratevalues for the required properties. However, it couldbe possible to derive vital clues by observing thegeneral layout and combine the available data withintuition. E.g., structural mass may be broken downto segments using statistical data depending on thematerials utilised and processes applied.

Once the FEA package has computed the output, thepresented solution set can be manipulated further, toproduce an optimised output through the process ofpost-processing and optimisation, which enable thedesigner to observe how the properties of the modelchange as certain features of the design parametersare varied. Optimisation is quite important as itallows obtaining the most effective solution not onlyin terms of performance but it can be expanded to acost saving process as well.

Discussion of the results

IntroductionFor the purpose of the analysis a truss based on theLS-400 was designed and a FEA model was made.Also a Flat Plate structure in order to model stressdistributions was created. The output for both casesgave useful insight regarding the force load path andthe effect of the structural layout in designing aneffective load bearing structure.

Truss like structureThe main problem associated with a truss-likestructure is the part count. Whether it is of metallic orcomposite structural members a large numbers offasteners or end-fittings is required. In this case theoption of a composite structure offers significantweight savings both directly, that is the density ofcomposites is lower compared to metals andindirectly that is adhesive bonding can besignificantly lighter than fasteners.

In terms of manufacturability, there is a greatflexibility. Members can be produced by extrusion,machining, formed out of sheet panels, and moulded.

The main problems associated with the truss structureoptions are the potential large part count especially inconnectors, end fittings, and fasteners, whichtranslates in increased costs due to the increasedmanufacture times, inspection and testing needed.

Additionally the existence of connectors eventuallycreates potential weak points in the structure, andlocations of probable failure whether being highstress concentrations around fastener holes alladhesive failure attributed to either high stresses orenvironmental effects.

Figures 4 to 12 show the stress distribution anddeflection of the LS-400 structure. It can be shownthat the connectors’ locations exhibit the higheststress concentrations while the lower frame of thepayload and antenna panels are deflected the most.

The problem of reducing part count can be addressedby implementing batch manufacturing likeSNAPSAT where large single pieces aremanufactured, co-curing of components and moulds.Moulds although it is an initially expensiveinvestment the benefits gained in the long run for avolume production run can offset that expense. Inaddition to weight savings this approach reduceswaste and lead times, making a composite approachcomparable to metallic structures therefore reducingproject cost considerably. Regarding metallic trusses,practices such as high speed machining can producelarge number of components at a minimal time withlow tooling costs.

“Flat designs”Flat designs include honeycomb, corrugated plate andisogrid. Regarding their use in the Small Satellitefield, all of their applications are focused primarily inthe use of these surfaces as mounting platforms forcomponent installation, although the panels theyformed could also act as load bearing components.

Their potential advantages for main structural use arequite attractive. They include issues like cheapmanufacturing, lightweight construction, reduced partcount and high strength. Additionally in terms ofassembly they offer a highly efficient area forcomponent mounting with the benefit of ease ofcomponent accessibility.

Their ease of integration with electronics in order toform multifunctional structures provides an additionaladvantage for their potential application in thesatellite manufacturing industry, since it can result toserious reductions in terms of mass and overall sizeof the satellite, thus offsetting the higher costs of suchapproach. Integration with the heat management and

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control system by the use of internal heat pipesreduces the overall mass of the satellite and the sizeof the protective enclosure. By careful selection of asun-synchronous orbit it is possible for appendages toprovide adequate shade for the components.

Considerations include issues regarding theirstructural efficiency at severe load cases such asignition and payload separation, where there is a highprobability of increased stress concentrations.Contrary to the isotropic properties of theconventional truss (especially the metallic ones),honeycomb, corrugations and isogrid structures are

orthotropic therefore additional structural analysis isneeded especially in the field of the possibility ofwrappage. Furthermore the proximity of thecomponents to the launch vehicle-payload clamp maycause failure if no vibration isolators or dampers areinstalled.

SummaryThe results and conclusions derived from theinvestigation of the alternative structural designs canbe shown on the next table.

TRUSS SINGLE PLATE/MULTI. HONEYCOMB

ADVANTAGES ADVANTAGES ADVANTAGESWell established analysis Offer reduction in weight and size. StiffnessGenerally ease of manufacture Offer systems integration. Thermal stabilityFlexibility and shape versatility Offer high accessibility LightweightChoice of materials Offer total savings by offsetting

investment and launch saving costs.Accessibility

Choice of manufacturing methods No special investment in tooling isneeded.

Low cost tooling

Ease of manufacture

DISADVANTAGES DISADVANTAGES DISADVANTAGESHigh tooling investment due todifferent parts. Specialised toolsmay be needed.

Unknown quantity

May produce large part count Initial investment high. Requirelarge production run to break even.(depends on method)

Isotropic load distribution requiresadditional testing

If metallic structure fasteners carrysignificant dead weight

Not attractive option for timepressing projects

Stiffening may be needed to protectagainst wrappage

If composite structure endfittingschoice and adhesive behaviour areconsiderations

CORRUGATION ISOGRIDADVANDAGES ADVANDAGES

Stiffness StiffnessThermal stability Thermal stabilityLightweight LightweightAccessibility Grid pattern of loading

DISADVANTAGES DISADVANTAGESIsotropic IsotropicStiffening may be needed toprotect against wrappage

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Figure 4 Deflection on structure

Figure 5 Effects of Diagonals on Structure

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Figure 6 Z/Y Stress

Figure 7 Effect of Diagonals on Z/Y Stress

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Figure 8 Von Misses Stress Layout

Figure 9 Effect of diagonals on Von Misses Stress

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Figure 10 Von Misses Stress Contours on panel

Figure 11 Deflections on panel. Lumped mass located at centre simulates payload.

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Figure 12 Von Misses on beams showing also the shell mesh. The stress is exaggerated by a factor of 2x inorder to show the effect on the end fittings.

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Acknowledgements

The authors would like to thank all those who provided with much needed information. Also credit should begiven to the technicians at Kingston University FEA labs for their support in maintaining the system in goodorder.

References

1. Sarafin Thomas, “Spacecraft Structures and Mechanisms, From Concept to Launch” , Microcosm Inc.and Kluwer Academic Publishers, 1995, Wiley J. Larson, ISBN 0-7923-3476-0

2. Serope Kalpakjian, “Manufacturing Process for Engineering Materials”, Addison Wesley Longman Inc.,ISBN 0-201-82370-5, 1997,pp891

3. Bendell A., ”Introduction to Taguchi Metghodology”, Taguchi Methods Proceedings of the 1998European Conference, Elsevier Applied Science, London England, 1988, pp 1-14

4. Swan Peter A. Manufacturing Technologies, the Key to a Small Satellite System. Technologies forEnhancing Small Satellite Systems, Motorola Symposium on Space Systems; 1994 ISBN: IAF-94-U.3.475.

5. Costabile V., Discepolli A., Fiorentino C., Morelli G., “New Spacecraft Assembly, Integration & TestingApproach for Globalstar Satellite Constellation” , AIAA-96-1149-CP, 1996

6. Thompson Timothy C.; Grastataro Cathleen; Smith Brian G.; Krumweide Garry, and Trembley Garry.,Development of an All Composite Spacecraft Structure for Small Satellite Programs

7. Krumweide, G. C. Affordeable Polymer Composite Structures for Various Spacecraft StructuralComponents. 9th Annual AIAA/Utah State University Conference on Small Satellites; 9/18/1995; Logan,UT, USA. AIAAISBN: A96-16830.

8. Greathouse J. et al , “The ISOSAT small satellite. A design in isogrid technology”, 6th Annual AIAA/ UtahState University conference on Small Satellites, Logan UT USA, ISBN A96-16830

9. NASA –STD –5001 “Structural Design and Test Factors of Safety for Spaceflight Hardware” , NASATechnical Standard.

10. http://rsdo.gsfc.nasa.gov/Rapidi/specs/loral/LS_400.ds.html