Structural Design of the MSX · PDF fileSTRUCTURAL DESIGN OF THE MSX SPACECRAFT JOHNS HOPKINS APL TECHNICAL DIGEST, VOLUME 17, NUMBER 1 (1996) 61 sides by aluminum honeycomb panels

STRUCTURAL DESIGN OF THE MSX SPACECRAFT

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tructural Design of the MSX Spacecraft

illiam E. Skullney, Harry M. Kreitz, Jr., Mark J. Harold, Steven R. Vernon, Teresa M.etenbaugh, Theodore J. Hartka, David F. Persons, and Edward D. Schaefer

he Midcourse Space Experiment (MSX) spacecraft represents the largeststructure ever assembled at the Applied Physics Laboratory. The design of the structureinvolves the application of conventional aluminum framework/honeycomb panelconstruction with composite materials technology to satisfy both the launch vehicleinduced loads and the dimensional stability requirements of the optical sensors andinstruments. The 1.59-m2 by 4.11-m-long structure is divided into three discreteelements consisting of (1) an electronics section, which provides an interface to theDelta II launch vehicle and a mounting platform for instrument and spacecraftelectronics; (2) a thermally stable graphite/epoxy composite truss structure formounting the Spatial Infrared Imaging Telescope III (SPIRIT III) instrument andproviding an interface between the instrument and electronics sections; and (3) atemperature-controlled instrument section with embedded heat pipes for mountingdelicate sensors and instruments. The instrument section also provides a thermallystable mounting structure for the optical bench and the beacon receiver bench. Thisarticle describes the features of the MSX structural design, the structural analysis, andtesting efforts associated with qualifying the design. A description of the effortsinvolved in the development of the graphite/epoxy composite optical bench andbeacon receiver bench is also presented.

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INTRODUCTIONThe Midcourse Space Experiment (MSX) spacecraft,

sponsored by the Ballistic Missile Defense Organization,is an observatory that provides a multisensor platformfor collecting background, target detection, and track-ing phenomenon data. Because of the precise pointing

and tracking requirements for the spacecraft, the phys-ical layout of the scientific instruments and sensors andof their electronics packages is critical to mission suc-cess. The 5.1-m-long (including optical sensors), 2812-kg spacecraft (Fig. 1) is divided into three discrete sec-

OHNS HOPKINS APL TECHNICAL DIGEST, VOLUME 17, NUMBER 1 (1996) 59

W. E. SKULLNEY ET AL.

tions: the instrument section, thetruss section, and the electronicssection (Fig. 2). The instrumentsection contains 11 sensitive opti-cal sensors that are precisely alignedrelative to one another so that tar-gets can be viewed and tracked si-multaneously by all sensors. Thestructure of the instrument sectioncontains honeycomb panels withembedded heat pipes to maintainall sensor mounting surfaces at auniform temperature regardless ofthe spacecraft’s orientation relativeto the Sun. Maintaining co-align-ment of the optical sensors is crit-ical during target tracking events.The center section of the spacecraftcontains the graphite/epoxy (G/E)composite truss structure that sup-ports the Spatial Infrared ImagingTelescope III (SPIRIT III). Thetruss structure connects the instru-ment section with the electronicssection while thermally isolatingthe sensors from their heat-generat-ing electronics packages. The elec-tronics section contains the sensorelectronics packages in addition toelectronics for the attitude control,power, command, and telemetrysubsystems necessary to operate thespacecraft. The aft end of the elec-tronics section structure containsthe cylindrical payload adapter formating the spacecraft to the DeltaII launch vehicle.

The instrument section alsocontains two ancillary structuresthat play an important role in thetracking and attitude control ofthe spacecraft. The optical benchis a precision inertial measurementplatform that is attached to the topof the instrument section structure (Fig. 2). The G/Ecomposite bench supports the ring laser gyroscopes anda star camera for establishing and maintaining space-craft attitude control. The beacon receiver bench is aprecision antenna platform that is attached to the 2yaxis honeycomb panel for launch (Fig. 2). Once orbitis achieved, the bench deploys into a position perpen-dicular to the 2y axis panel cantilevered off the topedge of the instrument section. The G/E compositebench contains four S-band receiving antennas thatcan acquire and lock onto a target signal beacon atfrequencies of 2219.5 and 2229.5 MHz during space-craft tracking activities.

Figure 1. The MSX

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spacecraft in APL’s Kershner Space Building.

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ELECTRONICS SECTIONTwo important functions in the mission of the MSX

spacecraft are performed by the electronics section.First, it is the major structural element in the spacecraftbus providing a direct load path for transferring launch-induced spacecraft loads into the Delta II launch ve-hicle. Second, it provides the mounting platform forelectronics packages associated with the spacecraftattitude control subsystem, spacecraft power and powerdistribution subsystem, spacecraft commanding sub-system, and spacecraft telemetry subsystem. Figure 2shows the MSX spacecraft with the electronics section

HNS HOPKINS APL TECHNICAL DIGEST, VOLUME 17, NUMBER 1 (1996)


+y

–y

+z –z

+x

–x

Top View

Opticalbench

Beacon receiver

bench

SPIRIT IIIinstrument

+z honeycombpanel with

solar array

–y honeycombpanel

–x honeycombpanel

Payloadadapter

ElevationView

Instrumentsection

Trusssection

Electronicssection

Figure 2. The MSX spacecraft configuration (stowed).

visible immediately beneath the truss structure. Theelectronics section is a hollow rectangular structurefitted with a cylindrical payload adapter to interface tothe payload attach fitting on the second stage of theDelta II launch vehicle. The overall dimensions of therectangular structure are 1.59 m2 by 1.12 m long, andthe overall dimensions for the payload adapter are 1.60m in diameter by 0.30 m long. The rectangular structureconsists of an internal skeletal frame covered on all

JOHNS HOPKINS APL TECHNICAL DIGEST, VOLUME 17, NUMBER 1 (1

sides by aluminum honeycomb panels. The interior andexterior of the honeycomb panels are densely populatedwith electronics packages used to power and operatethe spacecraft subsystems. In addition to providing thesquare-to-round transition necessary to mate the space-craft to the launch vehicle, the payload adapter pro-vides an overflow location for mounting electronicspackages and the electrical umbilical connections tothe spacecraft.

The skeletal frame of the electronics section is com-posed of several intricately machined aluminum struc-tural shapes and fittings that were precisely aligned andassembled using special high-tensile-strength fasteners.The assembled frame includes a unique triangularlyshaped bracket used to mount the attitude controlsubsystem’s reaction wheels in each inside corner. Thereaction wheel bracket was designed to precisely alignthe wheel’s axis of rotation at 0.44 rad to the y/z space-craft axes. The 0.44-rad inclination of the four reactionwheels provides redundant attitude and guidance con-trol of the spacecraft in the event of a single reactionwheel failure. The base of the frame includes a speciallymachined baseplate that initiates a smooth transitionfrom the square cross section of the electronics sectionto the round cross section of the Delta II launch ve-hicle. The transition is completed by attaching a cylin-drical payload adapter fitting to the underside of theframe baseplate. The skeletal frame was designed andfabricated at APL. Parts of the electronic section frame,the reaction wheel bracket, and the baseplate can beseen in Fig. 3.

The four sides (±y and ±z axes as shown in Fig. 2)and the bottom (2x axis) opening in the frame arecovered by 5.2-cm-thick aluminum honeycomb panelswith 0.6-mm-thick aluminum face sheets and integralmachined edge members. The honeycomb panels areindividually bolted to the faces of the frame to providethe lateral stiffness required for the entire assembly. Theelectronics section was designed so that the honeycombpanels on the ±z axes sides of the spacecraft can beremoved to provide access to the interior of the frame.The hollow center of the frame is used to house severalattitude control and data storage devices.

Figure 3 shows the interior of the electronics sectionframe as viewed with the +z honeycomb panel re-moved. Visible inside the frame are the followingdevices:

• Four each Reaction wheel assemblies mountedon triangularly shaped brackets in allfour lower corners.

• Three each Torque rod assemblies mounted on theinside surface of the 2x axis and +yaxis honeycomb panels.

• One each Reaction wheel control electronicspackage mounted on the inside surfaceof the 2x axis honeycomb panel.

996) 61

spacecraft battery and power distri-bution electronics packages.


• Two each Tape recorder assemblies (composedof separate electronics and recorderpackages) mounted on the inside sur-faces of the ±z axes honeycomb panels.

The exterior surfaces of the ±y, ±z, and 2x axeshoneycomb panels are used to mount an assortment ofelectronics packages, communication antennas, solararrays, batteries, and associated support brackets. Thehoneycomb panels contain threaded inserts, bondedintegral with the panel, used to attach the flight pack-ages. The +x axis side of the electronics section iscovered by two half panels, each 1.4 cm thick with 0.6-mm-thick aluminum face sheets and integral spools forattachment to the panels. This structure is primarilyrequired to provide the necessary in-place stiffness forthe top of the electronics section structure.

The cylindrical payload adapter was designed jointlyby APL and McDonnell Douglas Aerospace and fabri-cated by McDonnell Douglas Aerospace. The adapter,which completes the transition of the frame from thesquare-to-round cross section, is 1.60 m in diameter by0.30 m long with a skin thickness of 4.3 mm. The upperflange of the adapter contains a 48-hole bolt circle thatattaches to the baseplate of the frame with special high-tensile-strength fasteners. The lower flange of theadapter is machined to mate with the McDonnell Dou-glas 6306 payload attach fitting and separation clamp-band assembly. The hollow center section of the adapt-er surrounds the electronics packages mounted to theexterior surface of the 2x axis honeycomb panel. Theprimary hardware mounted in this area includes the

The solar arrayaxes honeycomb panel surface (0.1provide sufficientmounting electrofolded panels andfiguration. Two acare used to restrahoneycomb paneactive latches consemblies that alloorbit has been acto guide the foldedeployed solar arrsprings, compressiators is used to safrom the stowed toing of the solar acontrol electronichoneycomb panelSun sensors, locathe array control the boom to tracka solar wing thro

TRUSS STRUThe truss struc

Composite Opticsvides the primary and the instrumen

Tape recorder

Tape recorderelectronics unit

Torque rods(3 places)

Electronicssectionbaseplate

Reaction wheel andmounting bracket

(4 places)

Electronics sectionstructural frame

Figure 3. Interior of the electronics section.

62 JOH

SOLAR ARRAYSUBSYSTEM

The MSX solar array subsystemconsists of two solar array wingsdeployed along the z axis of thespacecraft. Each wing consists offour panels in a bifold configura-tion attached to a drive boom thatwill acquire and point the solararray toward the Sun. Figure 4 illus-trates the deployed solar arrays andtheir attachment to the electronicssection ±z axes honeycomb panels.The entire surface of each1.73 3 0.78 m panel is coveredwith 3 3 6 cm silicon solar cells.The complete solar array subsystemuses 4992 solar cells to generate abeginning of life power of approx-imately 1250 W at 34 V DC.

s are located in the center of the ±zpanels. The panels are spaced off the0 m) by a solar array drive adapter to room behind the stowed panels fornics packages. Figure 5 shows the the drive boom in the stowed con-tive latches and four passive latches

in the solar panels and boom to thel during prelaunch and launch. The

tain the pyrotechnic pin puller as-w the folded panels to deploy oncehieved. The passive latches are usedd panels as they unfurl to form the

ay configuration. A system of torsionon springs, and liquid dashpot actu-fely drive the solar boom and panels the deployed position. Active point-rray is controlled by the solar arrays unit, which is located on the 2z of the electronics section. Two pitchted on each wing, provide inputs toelectronics unit, which in turn drives the Sun. The drive system can rotateugh a full 360°.

CTUREture of the MSX spacecraft, built by, Inc., of San Diego, California, pro-support for the SPIRIT III instrumentt section. The critical requirements

NS HOPKINS APL TECHNICAL DIGEST, VOLUME 17, NUMBER 1 (1996)

+y, –z quad

–z

–y, –z quad

Pitch sensorsBoom

–z, +yOTBD

–z, +yINBD

–z, –yINBD

–z, –yOTBD

+z, +yOTBD

+z, +yINBD

+z, –yINBD

+z, –yOTBD

+y –y

+y, +z quad –y, +z quad

Boom deployment mechanismSolar array drive

Junction box

Drive adapter

Controlelectronicsunit

Spacecraft

Shunt plate(not shown)

+z

Panelrotationalactuators

Struts

Figure 4. MSX solar arrays: deployed configuration. Arrays arerotated with the panel “cell side” shown. INBD, inboard; OTBD,outboard; Quad, quadrant.

that drive the truss structure design are stiffness (space-craft frequency of 12 Hz lateral and 35 Hz axial), ro-tational stability of the SPIRIT III mounting plane(0.01°), mass (less than 54.4 kg), and strength (main-tain positive margins of safety). A low coefficient ofthermal expansion (CTE) (0.54 3 1026 in./in. per °C)and control of dimensional changes caused by absorp-tion and desorption of moisture were required toachieve the stability requirement. A low thermal con-ductivity (less than 5.0 W/m·°C) was necessary tominimize the heat transfer between the electronics andinstrument sections. Additional requirements consid-ered in the development of this unique part of the MSXstructure included stability performance when exposedto temperature gradients as a result of the colder thannormal temperature environment (+y side, 278°C; 2yside, 267°C; +z side, 271°C; 2z side, 267°C), surviv-al as a result of temperature extremes (290°C to+90°C), and a grounding impedance of ≤50 mV. Thiscombined set of requirements necessitated the use of



Panel rotational actuator(not visible)

Boom

Junction box

+z (–z)

Forward active latch

Pin puller

Solar array drive

Pin puller

Aft active latch

Boom & boomdeploymentmechanism

Side View

Inboard panel

Boomrotationalactuator

Passivelatch

Boom

Outboard panel

Passive latch

Solar arraydrive adapter

Solar arraydrive

Active latch pinpuller assembly

Boom deploymentmechanism

Hinge

Top View

+z

+y

+x

–x

Figure 5. MSX solar arrays: stowed configuration.

composite materials technology in lieu of more conven-tional materials such as aluminum for the fabricationof the truss section members.

The truss structure spans the full width of the struc-ture and is 1.97 m high. It interfaces at four corners tothe top of the electronics section and the bottom of theinstrument section (Fig. 2). The assembly uses 20 G/EI-beams connected together with titanium fittings orG/E plates. Three different sizes of I-beams were usedto construct the truss; they are classified as lower diag-onals, upper diagonals, and verticals. Unique titaniumfittings were designed to join the I-beams together andto establish the required interface for the SPIRIT IIIinstrument. The overall configuration for the trussstructure (Fig. 6) indicates the typical locations of these

96) 63


SPIRIT IIIinterface

Upperdiagonal

Instrumentsection

interfacejoint

Column

Electronicssectioninterfacejoint

Lowerdiagonal

1.97 m

1.45 m

1.49 m

Figure 6. MSX truss structure.

member groups. Figures 7 and 8 depict the differenttitanium fitting assemblies and illustrate how the de-tailed parts are interconnected. Special blind fasteners(Composi-loks) were used in these bolted joints toprovide a large blind-side footprint that distributes thebearing loads over a large area, thus preventing com-posite fiber delamination. High-strength titanium fas-teners were used at the SPIRIT III instrument, elec-tronics section, and instrument section interfaces.Bushings were used in the electronics section and in-strument section mounting holes to provide a closetolerance fit, thus minimizing the possibility of struc-tural movement during the launch environment. Har-ness clips were bonded to the truss members to providea means for securing the harness, which spanned theinstrument and electronics sections.

The truss elements were made from a T50/ERL1962G/E composite material configured in a pseudoisotropiclaminate stacking sequence of (02°/60°/260°)S3. T50graphite fibers were chosen because of their minimalthermal conductivity and high stiffness characteristics.Amoco’s ERL1962 resin was selected because it is atoughened resin system that has excellent mechanicalproperties and exhibits good resistance to microcrack-ing. Composite materials can be tailored to enhance

their thickness witThese elements weorder to maintainments. Prior to asseto individual G/E etion and desorptiowith Chemglaze ZThe eutectic coatia Composite Optica precise combinatexact thickness (0truss assembly was54.4 kg.

INSTRUMENThe MSX instr

mally stable platforof MSX instrumenminimize Sun-indvary the co-alignmious spacecraft obstrument section Visible Imagers anment (nine spectr

64 JOHN

properties in specific directions ofthe material, i.e., the material is“directionalized.” For the truss, theT50/ERL1962 combination is di-rectionalized to provide enhancedstiffness. Testing was performed tocharacterize the stiffness, strength,bolt bearing capacity, CTE, andshear strength. The material prop-erties are shown in Table 1. Thestructure used joint configurationsfeaturing double lap shear. A spec-imen test was performed to accu-rately evaluate the load transfer ofthe G/E I-beam section to the ti-tanium corner fitting. In-plane andthru-thickness thermal conductiv-ity measurements showed that thethermal conductivity was approxi-mately 4 to 5 times greater thanthe targeted thermal conductivity.As a result of these tests, otherpotential fiber candidates were re-evaluated; however, it was con-cluded that the T50 fiber selectedwas the best for this application.

The titanium fittings and G/Eplates were machined precisely toensure proper assembly with themating I-beams. The I-beams re-quired special tooling to control

hin tolerance during the G/E lay-up.re assembled using special fixtures in the overall dimensional require-mbly, a eutectic coating was appliedlements to control moisture absorp-n. Each member was then painted306 flat black polyurethane paint.ng, or moisture barrier, consisted ofs, Inc., proprietary process that usedion of tin and indium applied to an.0127 mm). The final weight of the 47.1 kg, well below the budgeted

T SECTIONument section must provide a ther-m to accurately support the majorityts. This structure was designed to

uced thermal gradients that wouldents of the instruments during var-

servations and maneuvers. The in-suite includes the Ultraviolet andd Spectrographic Imagers experi-

ographs and imagers), two horizon

S HOPKINS APL TECHNICAL DIGEST, VOLUME 17, NUMBER 1 (1996)

RUCTURAL DESIGN OF THE MSX SPACECRAFT
ST
sensors, three dilamp, a total pretemperature-conkrypton lamp, threference object ages. The instrustalled in the M

The instrumemachined aluminer to form a largthe electronics structure at four cfor the spacecraftwork consist of shaped center mchined corner members are useditional framewoment honeycombstructure measurused for attachineral SPIRIT IIIbench.

Very precise required on the tolerance builduparts of the strucstructure was ascustom-fabricateimize the flatnesall sides. After alassembled and drilled and permaers were installeextremely precisinstrument hone

The honeycoframe structure spacecraft, preserily related to thpanels is a 3.3-cm

Table 1. Material properties of T50/ERL1962.

Property MeanTensile strength (0°/90°) 4.34/3.46 3 108 N/m2

Tensile modulus (0°/90°) 8.69/9.28 3 1010 N/m2

Tensile Poisson ratio (0°/90°) 0.30/0.34Compression strength (0°/90°) 3.29/3.34 3 108 N/m2

Compression modulus (0°/90°) 8.07/8.83 3 1010 N/m2

Compression Poisson ratio (0°/90°) 0.32/0.33Flatwise tensile 1.17 3 107 N/m2

Two-rail shear 2.25 3 108 N/m2

Short beam shear 7.34 3 107 N/m2

CTE (0°/90°) 0.56/0.45 3 1026 in./in. per °C

Instrumentinterfaceend cap(titanium)

Upper diagonalinterface fitting(titanium)

Doublerplate

Column flangestrimmed atfitting location

Doublerplate

Upper diagonal I-beam

Figure 8. Upper diagonal interface fitting. All components areG/E unless noted.

SPIRIT III interfacefitting (titanium)

Fillerplate

Upper diagonalI-beam

Doublerplate

Gussetplate

Backingplate

Doubler plate

Lower diagonalI-beam

Figure 7. SPIRIT III interface fitting. All components are G/Eunless noted.


gital solar attitude detectors, a xenonssure sensor, a mass spectrometer, fourtrolled quartz crystal microbalances, ae Space Based Visible Telescope, six

canisters, and several electronics pack-ment section assembly is shown in-SX spacecraft at the top of Fig. 1.nt section structural design consists ofum structural members joined togeth-e “box”-style frame similar to that ofsection. It interfaces with the trussorners and provides the four lift points. The structural elements of the frame-“L”-shaped edge members and “U”-embers joined with intricately ma-

fittings and fasteners. The centerd on the y and z sides to provide ad-rk rigidity and support for the instru- panels. The instrument section framees 1.59 3 1.59 3 0.72 m and is alsog secondary structures that mount sev- electronic chasses and the optical

controls on tolerances and size wereinternal frame assembly to minimizeps as panels, instruments, and otherture were attached at later dates. Thesembled using a large granite table,d squares, and tooling in order to max-s, parallelism, and perpendicularity ofl the structural parts of the frame wereadjusted, the assembly was matchednent shear and tension Hi-Lok fasten-

d. The assembly process produced ane finished structure for mounting theycomb panels.

mb panels, which are attached to theon the four sides (±y and ±z) of thented several design challenges prima-e thermal requirements. Each of these

-thick rectangular-shaped aluminumhoneycomb panel with 0.5-mmaluminum face sheets, and eachcontains three embedded heatpipes for transferring heat from theinstruments to the radiator assem-blies. Many of the threaded spoolinserts used to mount the instru-ments are attached directly to theheat pipes, thus providing a con-ductive heat path to the radiatorassemblies. In addition, the +y sidepanel heat pipes extend beyondboth edges of the panel in order tointerface with the radiator assem-blies mounted on the +z and 2z

996) 65


panels. The 2y side panel containsfew packages; thus heat is trans-ferred through the frame via con-duction. Smaller packages thatwere not well defined when thepanels were fabricated are to be at-tached to the panels with surface-mounted rivenuts installed at alater date. The side-panel internalsurfaces were masked and paintedwith Z306 Chemglaze paint to augment radiating ca-pabilities, thus enhancing isothermal performance. Thepanels and frame were then coated with thermal greaseto maximize heat transfer between the panels andframe. Special “Milson” close-tolerance, high-strengthfasteners are used to attach the panels to the framestructure. These fasteners prevent movement of thepanels relative to the frame under load. They alsoprovide accurate repositioning of the panels shouldthey need to be removed and reinstalled.

Four “half-moon”-shaped honeycomb panels closeout the top and bottom of the instrument section struc-ture and provide additional rigidity and in-plane stiff-ness to the structure. All panels are made from 2-cm-thick aluminum honeycomb with 0.5-mm-thick facesheets. They are designed to be removed to facilitatethe installation of the SPIRIT III instrument and pro-vide adequate clearance around the instrument, wheninstalled. In addition, the top panels (+x side) containembedded structural members that are used for attach-ing the six reference object canisters.

SPACECRAFT STRUCTURALANALYSIS AND TESTING

The MSX primary structure was designed to providea safe mounting platform for the spacecraft bus andinstruments during the launch scenario and to maintainon-orbit co-alignment requirements of the instruments.To satisfy these requirements, spacecraft structurestrength, stiffness, and stability issues were addressed.

The McDonnell Douglas Delta II Payload Planner’sGuide recommends spacecraft structure fundamentalfrequencies above 35 Hz in the thrust axis and 12 Hzin the lateral axes to avoid dynamic coupling betweenlow-frequency vehicle and spacecraft modes. Dominantlaunch vehicle induced loads result from the liftoff andmain engine cutoff flight events. The maximum ex-pected flight loads, or limit loads, are calculated bymultiplying the spacecraft weight by a limit load factor(in g). The spacecraft limit load factors specified in thePlanner’s Guide are listed in Table 2. These two re-quirements, although independently defined, are relat-ed in the sense that failure to achieve the stiffnessrequirement could have a detrimental impact on theload conditions. Thus, the proper combination of these

Table 2. MSX

ParameterThrust (x) axisLateral (y, z) axNote: Loads are asion to launch ve

66

two requirements is necessary for equipment to survivethe launch environment.

To meet instrument co-alignment requirements, thetotal structural and on-orbit distortion for all of thestructure above the SPIRIT III mounting interface wasnot to exceed 0.042°. This requirement represents acombined mechanical and thermal loads problem thatdirectly affected the design of the truss and instrumentsection structures.

To address the structural requirements, a finite ele-ment model (FEM) representing the characteristics andproperties of the primary spacecraft structural elementswas developed. Portions of the 1839-element modelshown in Fig. 9 are based on models developed byexternal organizations supplying hardware, e.g., Engi-neering Consulting Resources Associates supplied theSPIRIT III model, Rockwell supplied a solar arraymodel, McDonnell Douglas supplied the spacecraftadapter model, and Composite Optics supplied the trussand beacon receiver bench models. Individual FEMswere also developed for the instrument and electronicssections, the truss, and the adapter to better evaluatelocal detailed stresses as a result of the applied loadconditions. The NASTRAN software developed byMacNeal Schwendler Corporation was used to analyzethe model when it was subjected to the quasi-staticloads in Table 2 and to determine the fundamentaldynamic characteristics (normal modes) of the space-craft. For quasi-static loads analysis, factors of safety(FS) were applied to the limit loads to account for testload margins and to ensure that an adequate designmargin exists. The FS used on the MSX program areas follows:

Material yield FS = 1.4Material ultimate FS = 1.8Material buckling FS = 2.0

Margin of safety (MS) calculations were made toindicate the degree of strength capabilities beyond thedesired factors, i.e., MS ≥ 0 is acceptable, and is definedas

MS= material allowable strength FS applied stress×

− 1 0. ,

where applied stress is the material stress resulting fromthe applied limit loads. A summary of the minimum

spacecraft limit load factors (Delta II).

Liftoff Main engine cutoff22.4/+0.2 g 27.1 g

es ±2.5 g —pplicable at the spacecraft’s center of gravity. The dash implies compres-hicle interface.

JOHNS HOPKINS APL TECHNICAL DIGEST, VOLUME 17, NUMBER 1 (1996)

Concentratedmasselements

Instrumentsection

Trussstructure

Solararrays

Separation planeAdapter

Electronicssection

SPIRIT III

Beaconreceiver

x

z y

Figure 9. MSX launch configuration finite element model.

Table 3. Minim

Structural comInstrument secInstrument secElectronics secElectronics secTruss structureTruss fastenersAdapter structu

Table 4. MSX

DescriptionSpacecraft benSolar array benSPIRIT III (y a

(z axis) bendSpacecraft torsSpacecraft axia

margins of safety for the quasi-static analysis of theMSX primary structure is presented in Table 3. Thefundamental dynamic characteristics are defined inTable 4.

To address the on-orbit align-ment issues, worst-case thermaltemperature distributions and 1 geffects were examined. The resultsfrom the finite element analysisindicated that maximum distor-tions were less than 0.02°.

The spacecraft FEM was analyt-ically coupled to a Delta II launchvehicle model to predict spacecraftresponses as a result of launchevents. Although the primaryspacecraft modes (8.4 and 29.9 Hz)do not meet the minimum Delta IIrecommendations (12 and 35 Hz),results from the coupled loads anal-ysis performed by McDonnell Dou-glas verified that the limit load fac-tors presented in Table 2 did notchange.

Once fabricated, the primarystructural elements were tested todemonstrate acceptable structuralperformance. Because the size of

JOHNS HOPKINS APL TECHNICAL DIGEST, VOLU


the MSX spacecraft precluded structural testing at thesystem level, static load tests were performed on eachof the spacecraft’s major structural sections. Thesestructures were proof tested to verify that they werecapable of surviving the worst-case predicted flightloads plus a 25% margin. Test scenarios were based ongenerating worst-case load conditions in critical struc-tural members. Strain gauges and deflection indicatorswere used in each test sequence to monitor the stressesin key structural elements to ensure that the designstress levels were not exceeded. The data collected wereused to verify and calibrate the analytical model of thespacecraft used in previous loads analysis.

Two separate tests were performed for each of theinstrument and electronics sections. The first, depictedin Fig. 10, was performed to verify the adequacy of thecorner fittings and attachments as a result of flight orlifting load situations. Tensile loads of 4535 kg and1135 kg were applied at each of the four corners for theelectronics and instrument sections, respectively. Theoverall structural integrity of each section was verifiedby application of a lateral load in combination with acompression load as shown in Fig. 11. Lateral loads of9070 kg and 1815 kg and compression loads of 2040 kgand 726 kg for the electronics and instrument sections,respectively, were applied in four orthogonal directionsto thoroughly check all structural elements and all pos-sible loading scenarios. Results of the testing indicatedthat both structural sections exhibited more than ad-equate strength to withstand the maximum anticipatedflight loads. Although stresses were significantly lowerand deflections somewhat higher than anticipated, this

um margins of safety: MSX spacecraft primary structure.

ponent Margin of safetytion structure 0.33tion fasteners 0.17tion structure 0.14tion fasteners 0.18 elements 0.31

0.23re 0.11

fundamental modes.

Frequency (Hz)ding modes: y and z axes 8.4ding modes: z axis 19.6nd z axes) and solar array 19.8ing modesion mode 19.9l mode 29.9

ME 17, NUMBER 1 (1996) 67


Figure 10. Static load test set-up: corner fittings.

Corner tensile load

Electronics orinstrument section

Adapter simulator (electronics)or mounting feet (instrument)

Mounting plate

Strong back

Lateral load(electronicsand instrument)

Load fixture

Electronics orinstrumentsection

Adapter simulator (electronics)or mounting feet (instrument)

Mounting plate

Strong back

Compressive load

(instrument)

Compressive load

(electronics)

Figure 11. Static load test set-up: overall structural integrity.

was attributed to joint slippage that occurred betweenthe structural frame and the honeycomb panels.

The truss section test, depicted in Fig. 12, was de-signed to test the lower diagonal member of the truss,the most critical member as defined by previous anal-yses, to a compressive load of 4322 kg. Although properloading was achieved, deflection results indicated alateral deflection 2.5 times the predicted value. Thisdiscrepancy was attributed to inaccuracies in the modelused for predicting test results, movement of the trussrelative to the base fixture/floor, and rigid body rotationof the truss/base fixture assembly. Relevant findingswere incorporated into the analysis model, which wassubsequently used to verify that there was no impact onthe spacecraft loading conditions.

68 JOH

OPTICAL BENCHDuring science missions, alignment shifts between

the optical bench components (two ring laser gyros,one star camera, and two autocollimators) should bemaintained within 2-arcsec relative rotation and0.0254-mm relative displacement in order to accuratelydetermine the SPIRIT III boresight direction. Any un-known or unmeasured rotations in the bench structurewill affect the component measurements and would betransformed into an apparent error in the SPIRIT IIIboresight direction. These requirements dictated theneed for a very precise and stable platform to supportthese components through the various on-orbit condi-tions. Use of conventional materials such as aluminumfor the bench structure would not provide the desiredstable features, thus dictating the use of a G/E compos-ite material. Because of the stringent requirements, theon-orbit operating temperature of the structure will bemaintained at 20°C and thermal gradients will bemaintained within 1°C in the plane of the bench andbetween the top and bottom face sheets of the bench.Furthermore, the dimensional and rotational stabilityalso dictated that the CTE for the laminate be mini-mized (a goal of 0.0) and that the bench mountingsystem would not allow a rigid body rotation greaterthan 0.006° (22 arcsec) from its nominal position. Thelatter requirement coupled with the 2-arcsec require-ment dictated the use of a special, unique kinematicmounting system.

The optical bench structure is mounted on top of theinstrument section structure near the aperture for theSPIRIT III instrument. The system consists of a mount-ing platform with two skewed equipment mountingplanes and three kinematic mounts to provide support.The system weighs 4.45 kg and is shown in Fig. 13. Thedistinctive geometry for the 0.34-m-wide by 0.97-m-long by 0.038-m-thick structure was dictated by sensormounting requirements and the space restrictions inthe MSX configuration. The bench structural arrange-ment is a sandwich construction consisting of a discreterib core with top and bottom face sheets. Both the ribsand face sheets were made from a unidirectional tapewith graphite fibers embedded in Amoco’s ERL1962resin. The P75S fibers (5.17 3 1011 N/m2 modulus)were selected because of the demanding thermal stabil-ity and stiffness requirements for the system; theERL1962 resin was selected for the same reasons asspecified for the truss structure. An added benefit ofusing this G/E system is the reasonable cost of thematerials and the substantial materials database thatalready exists as a result of previous use. EA 9394adhesive was selected for bonding the G/E parts togeth-er because it is less ductile and more stable for opticalbench applications. Components were mounted to thebench structure via inserts that were directly attachedto the ribs.


TURAL DESIGN OF THE MSX SPACECRAFT
STRUC
Instrumentsection

interfacejoint

Column

Electronicssection

interfacejoint

Lowerdiagonal

P1

Applied Loads P1 = 2313 kg P2 = 2667 kg P3 = 2667 kg

x

yz

P2

P3

Figure 12. Truss structure static load test set-up.

fundbenanddyn(bening (benquaexcexhditiwhistreA sof tTab

Tstabthe FEMcon

The structure constituents (ribs, face sheets, andgusset plates) were constructed from multiple layers ofthe unidirectional tape oriented in a (0°, 30°, 60°, 90°,120°, 150°)SN lay-up to yield the pseudoisotropic con-dition. The core ribs were 12-ply and the face sheetswere 24-ply laminates. Material properties for this con-figuration are given in Table 5. A tin-indium moisturebarrier similar to that used on the truss was applied topreclude the effects of moisture on the stability of theoptical bench. This barrier also helps to minimizemoisture contamination of the sensor optics located onthe forward part of the structure.

The configuration for the kinematic mount systemuses a ball mount concept that provides nonredundanttranslational support in one, two, and three directionsfor the three mounts. The attachment of the mountsto the optical bench is located at the central plane ofthe bench thickness. This method of attachment limitsthe moment transfer between the spacecraft and opti-cal bench, thus minimizing the distortion of themounted structure. The mounts were made from tita-

JOHNS HOPKINS APL TECHNICAL DIGEST, VOLUME 17, NUMBER 1 (1996)

nium to provide good thermalcompatibility (low CTE) with thecomposite bench; in addition, tita-nium is a relatively poor conduc-tor, thus providing thermal isola-tion between the bench and theinstrument section.

The optical bench structuremust meet all requirements afterexposure to temperature extremesof 240°C and +60°C. Quasistaticload strength requirements of212.5 g/+6.5 g in the spacecraftthrust direction (x) and ±6.5 g inthe spacecraft lateral direction (yand z) were imposed to ensure thatthe hardware will survive thelaunch environment. In addition,load constraints were placed oncomponent attachments (±35 gnormal and ±20 g parallel to thecomponent’s mounting plane) andlifting attachments (±3 g in the liftdirection [x]) to guarantee the in-tegrity of these attachments. Astiffness requirement of 100 Hz wasdesired to preclude dynamic inter-actions with other structural com-ponents. Given all these require-ments, the system’s mass was not toexceed 6.8 kg.

To predict the capability of theoptical bench system, an FEM wasgenerated to examine quasi-staticloading conditions, to determine

amental dynamic characteristics, and to evaluatech distortions. After examining numerous concepts iterating to the final version, the fundamentalamic characteristics were determined to be 111 Hzch bending in the thrust direction), 138 Hz (twist-of the star camera mounting plane), and 203 Hzch bending in the lateral direction). Results of the

si-static loading analysis indicated that, in general,ept for certain local areas, the structural designibits ample structural margins under the given con-ons. For components such as the optical bench,ch are driven by stiffness requirements, moderatess levels as a result of mechanical loads are typical.ummary of the minimum margins of safety for eachhe bench structural components is presented inle 6.o determine compliance with the alignment andility requirements, a thermal distortion analysis foroperational conditions was performed using the. This included not only the thermal gradient

ditions but also the effects of the breakaway torque

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Star camera

Thermalblanket

GyroGyro

Planebench

Figure 13. MSX optical bench (structure shown in lower right).

Table 5. P75S/ERL1962 material property summary.

Property MeanTensile strength (0°) 2.951 3 108 N/m2

Tensile modulus (0°/90°) 1.05/1.03 3 1011 N/m2

Tensile Poisson ratio (0°/90°) 0.32/0.32Compression strength (0°/90°) 1.68/1.78 3 108 N/m2

Compression modulus (0°/90°) 0.95/1.04 3 1011 N/m2

Compression Poisson ratio (0°/90°) 0.32/0.32Flatwise tensile 1.17 3 107 N/m2

In-plane shear strength 1.81 3 108 N/m2

CTE (0°/90°) 20.22 3 1026 in./in. per °C

of the kinematic mounts. The results indicate a max-imum relative displacement of 0.00635 mm and a max-imum relative distortion of 1.35 arcsec for the operat-ing conditions. This maximum occurs between the starcamera and autocollimators. For each inch-pound ofbreakaway torque, an additional 0.1 arcsec of rotationaldeformation would occur. To minimize the impact ofthis, bearings with breakaway torques not to exceed

0.17 kg·m were seanticipated distorarcsec. Breakawaydisplacement.

All G/E parts fofrom pseudoisotrofabricated that alcrete core details

70 JO

lected for this application. The totaltion for this configuration is thus 1.65 torque has a negligible impact on

r the bench structure were machinedpic flat stock. Special fixtures werelowed accurate assembly of the dis-in relation to the equipment attach-

ment fittings and the kinematicmount interface details. All G/Ecomponents were cleaned, platedwith the moisture barrier, andpainted before assembly with thekinematic mount metal parts. Oncethe bench structure was assembled,the optical bench was subjected tothe thermal extremes, modal sur-vey, static loads, and thermal distor-tion tests. Although not a specificrequirement, a ±1 g distortion as-sessment test was also conducted toevaluate the change in bench dis-



Table 6. Structural margin of safety summary.

Component Minimum margin of safetyG/E facesheets 1.5Ribs 0.7Adhesive bond line 0.02Kinematic mounts 0.4Hardware attachments 0.01

tortion between the ground and orbit environments. Abrief summary of each of these tests is presented next.

The thermal extremes and static loads tests wereconducted primarily to verify the integrity of the benchstructure. First the bench was subjected to 15 thermalcycles between 240°C and +60°C to verify bond lineintegrity. Subsequently, two static load test cases rep-resenting the worst-case tensile and compressive loadcases were performed. The bench was mounted to athick aluminum plate, and mass simulators were usedto allow the loads to be applied at the componentcenters of gravity. A test factor of 1.25 was applied toall loads to demonstrate margin over the worst-caseexpected loads. Post-test inspections revealed no dam-age to the bench structure or any bond lines.

Prior to static loads testing, a modal survey test wasconducted to verify the dynamic characteristics of theconfiguration. The configuration consisted of thebench assembly, without components, mounted to athick aluminum plate. A post-static loads test was alsoperformed to verify that no changes occurred as a resultof the structural testing. The results of 352 Hz and 360Hz (pre- and post-static test, respectively) comparedfairly well with the predicted value of 402 Hz. Exam-ination of the analytical model indicated the differencewas due to artificial rigidity introduced in interelementconnections. When the model was adjusted for morerealistic flexibilities, a natural frequency of approxi-mately 370 Hz was obtained.

Bench stability was determined by examining thethermal distortion between the star camera and theautocollimator mounting planes when a thermal gradi-ent was applied through the thickness of the benchstructure. A laser, beamsplitter, and mirrors were usedto measure the relative difference between the twolocations. Results of 14 test cases indicate an averagedistortion of 2.48 arcsec/°C. Although this measure-ment exceeds the goal of 2 arcsec/°C, it was determinedthat the slightly larger error would not have a signif-icant impact on scientific measurements.

BEACON RECEIVER BENCHThe beacon receiver subsystem must acquire coop-

erative targets with an initial pointing uncertainty of±5° 3 ±5° at a range of up to 8000 km and track them


with a ±0.1° 3 ±0.1° residual rms error during the5-year orbital mission of the MSX spacecraft. To ac-complish this objective, the deployed beacon receiverstructure must contribute no more than 0.5° to theinitial pointing error of and provide stability within0.03° for the system’s four antenna arrays. These re-quirements must be satisfied over a fairly wide temper-ature range (265°C to 215°C) and when the structureis subjected to large thermal gradients (up to 30°C). Inaddition, stiffness requirements dictate that the funda-mental frequency of the stowed bench structure mustbe greater than 50 Hz and less than 65 Hz in thespacecraft ±y direction to preclude any dynamic cou-pling with the spacecraft primary structure. Long-termcreep in the bench as a result of the 362.8-kg prelaunchcenter release preload must be minimized. Whenviewed as a complete package, the demanding andsometimes conflicting requirements of accurate initialpointing after deployment, long-term pointing stability,passive thermal management, and very low operationaltemperature with room temperature calibration allcombine to produce a very challenging mechanicaldesign problem.

The beacon receiver bench assembly consists of aG/E composite bench, four aluminum dish antennaarrays, a magnetometer, two hinge and damper assem-blies, electronics, and a latch and release assembly. Theapproximate size of the assembly is 1.397 3 1.143 30.305 m and it weighs approximately 36.3 kg. Fivetitanium brackets attach the receiver structure to the2y experiment section honeycomb panel and providesupport, hinging, and latching functions. This system,shown in Fig. 14, will be stowed during launch anddeployed within 4 hours of separation of MSX from thelaunch vehicle.

The G/E bench structure is designed to provide amechanically and thermally stable platform to supportthe receiver antenna array and associated electronicsafter deployment from the stowed or launch configu-ration. G/E composite materials were chosen in lieu ofmore conventional materials because of the combina-tion of stringent requirements for alignment, stability,distortion, and stiffness. The bench structure, built byComposite Optics, Inc., was chosen to be similar inconstruction to the MSX optical bench, using G/E faceskins separated by a G/E discrete rib core. The sameG/E material system (0.1-mm-thick P75S/ERL1962unidirectional tape) was also used to fabricate the facesheets and ribs because of the similarity in require-ments, although a different pseudoisotropic lay-up[(0°/36°/72°/108°/144°)S] was used to provide a mini-mal in-plane CTE. Material properties for the pseudo-isotropic lay-up are given in Table 5. Various additionalG/E angle clips and doublers are used in areas of highstress. Threaded titanium fittings, bonded onto the ribsand face sheets, were used for bolted attachments. As

96) 71


Antenna array(shown with radomeand thermal shield)

Hinge and damperassembly

G/Ecomposite

bench

Results of this afrequency of the structure caused shown to meet thG/E material allcreep effects. Mstructural load caand are shown inture. Table 8 givethermal distortio

In addition, aformed on an FElytically determinmance of the CTantenna dishes anthis analysis1 sho

To evaluate sying hinge deploystructure was fabfrom Invar to appstructure. It also service loops andAn aluminum sufive support poindeployment andmounted on thepyrotechnic releaperformed to deployed position. gaged for three tmeasurement of tlocked position. the hinge would cbly to within 0.11the 0.5° requirem

Figure 14. MSX beacon receiver subsystem.

with the truss and optical bench structures, a tin-indium metallic coating was applied to internal andexternal surfaces after assembly to ensure long-termdimensional stability against the effects of moisture.

The concept for the hinge design evolved fromsmaller hinges used successfully on previous APL mis-sions. The nonmagnetic beacon deployment hingeassembly uses beryllium copper torsion springs, titani-um supports and shafts, bronze bushings, and stainlesssteel dampers and functions much like an old-fashionedsaloon door. The hinge is designed to deploy the struc-ture such that the antenna array’s mechanical boresightis initially aligned within ±0.5° of the spacecraft’s x axisand is maintained within 0.01°. Stored energy in thepreloaded torsion springs deploys the bench when re-leased and positions the bench when fully deployed. Atorsion damper assembly consisting of 3.175-mm-thickstainless steel strips was used in the system to preventexcessive overtravel during deployment, thus avoidinginterference between the beacon receiver and otherspacecraft hardware. During deployment, the damperassembly absorbs the kinetic energy of the system whenit travels past its deployed position, and the torsionsprings wind back up to provide a resisting torque thatforces the bench to seek the final, deployed position.Clearances internal to the hinge compensate for differ-ent CTEs between the G/E composite bench and thealuminum instrument section. The center release andpin puller assemblies function to accurately preload thebench assembly during launch as well as to release itfor deployment after launch. The 362.8-kg preload isprovided by a beryllium copper bolt that is instrument-ed with a strain gauge to ensure that the desired tensionis set accurately. The joint design allows liberal toler-ance accumulations to be accounted for between the

72 JOH

bench and instrument sectionstructure during fabrication andassembly through the use of a rodend bearing, which provides rota-tion and pivot. Upon commandfrom the spacecraft, the pyrotech-nic activated pin puller retracts thepin and the bench is released.After the beacon receiver has beendetermined to be in the deployedposition, the pyrotechnic activatedorbit latch is engaged to provideposition stability to the benchagainst spacecraft torques duringon-orbit maneuvers.

An FEM of the bench assemblywas constructed to analyze quasi-static launch loads, latch preloadeffects, system dynamic character-istics, random vibration loading,and thermal stress and distortion.

nalysis indicate that the first naturalstructure is 55 Hz. All stresses in theby the center release preload weree goal of being less than 10% of the

owables, thus minimizing long-terminimum margins of safety for otherses were determined to be acceptable Table 7 for various parts of the struc-s the boresight pointing results for then cases analyzed. thermal distortion analysis was per-M of a single aluminum dish to ana-e the effect on radio-frequency perfor-E mismatch between the aluminumd the G/E bench structure. Results of

w dish distortion to be within budget.stem positioning and accuracies dur-

ment, a mass simulator of the receiverricated. The simulator was fabricatedroximate the CTE of the G/E bench

served as a mockup for cable and coax thermal blanket design and tie downs.pport frame was fabricated to providets and a g-negation system for hinge

stability testing. Optical flats were frame and a series of manual andse deployments of the assembly were

termine the repeatability of the de-In addition, the orbit latch was en-ests to determine an average angularhe deviation from the deployed to theResults from this testing indicate thatonsistently deploy the receiver assem-5° of the proper position, well within

ent. After the orbit latch was engaged,


RUCTURAL DESIGN OF THE MSX SPACECRAFT
ST
of any other sptested in APL’s Ksize and weightground support eand transportatsignificant pieceshipping contai

The high denecessitated theaccommodate plargest package strument is the lon the 2x honunderside of thehydraulic table quirement dictavate the spacecrunderneath for draulic table), fixture (shown “elephant standthe floor, givingfor all required support the spacprovide access fothe spacecraft fostiff spacecraft mthe need to remphant stand wafigurations to detimes during inmove the spaceconfiguration wphant stand wathe base plate tcomplish this taspacecraft and efor example, betFlight Center (Gtesting.

A special shiwas designed tobuilding limitatof transportatio

Table 7. Beacon receiver margin of safety summary.

Component Minimum margin of safetyG/E facesheets 0.17Ribs 0.21Titanium fittings 1.41Adhesive bond line 0.01Inserts 0.56

Table 8. Beacon receiver dimensional stability results.

Requirement Actual AllowableArray boresight alignment with spacecraft x axis 0.0000471 rad 0.00524 radDish-to-dish boresight relative angles 0.000259 rad 0.00175 radDish-to-dish relative translations in the dish

mounting plane 0.02245 mm 0.2032 mmDish-to-dish relative translations normal to

the dish mounting plane 0.02972 mm 0.0762 mmDish rotation about a dish boresight axis 0.0000029 rad 0.00262 rad

the hinge was stable in the presence of disturbancetorques to 0.003°, well within the 0.01° requirement forhinge play. The simulator and frame assembly werethen placed in the large APL thermal vacuum chamberand a cold deployment was successfully accomplished.The orbit lock was successfully engaged during the sub-sequent cold soak to complete hinge, latch, and orbitlock verification testing.

Prior to assembly, the G/E structure was successfullysubjected to a static loads test to 1.25 times the qua-sistatic design loads, a temperature extremes test, anda moisture loss test, thus verifying performance. Thebench-mounted electronics as well as the orbit latchwere subjected to sine and random vibration at thecomponent levels. The beacon receiver assembly, in-cluding the composite material structure, dishes, ther-mal shield, multilayer insulation, hinge, support brack-ets, and center release, was subjected to three-axis sineand random vibration testing to verify system-levelcapability. The first natural frequency of the receiverassembly was determined to be 49 Hz, slightly below theinitial 50-Hz requirement but nevertheless satisfactoryfrom a stiffness viewpoint. Before and after vibrationtesting, the center release mechanism was actuated andthe unit was walked out through its full range of motionto verify proper operation of the deployment mecha-nism. All testing was completed satisfactorily with noproblems encountered.

GROUND SUPPORT EQUIPMENTThe MSX spacecraft, at 2812 kg, 5.1 m in height,

and 3.3 m in diameter, is 2 to 3 times the size and weight


acecraft that has been integrated andershner Space Building. Because of its

, several unique pieces of mechanicalquipment were required for integrationion of the spacecraft. The two mosts of hardware, the “elephant stand” andner, are described here.nsity of packages on the spacecraft use of the space inside the adapter tolacement of the battery, the secondon the spacecraft (the SPIRIT III in-argest). The 90.7-kg battery is mountedeycomb panel surface (essentially the spacecraft) and requires the use of afor mechanical integration. This re-

tes a special fixture to sufficiently ele-aft above the ground and provide accessmechanical operations (using the hy-wiring, and thermal blanketing. Thisat the bottom of Fig. 1), named the,” supports the spacecraft 0.9 m above adequate room beneath the spacecrafttasks. One of the six gusseted legs thatecraft mounting surface is removable tor use of the hydraulic table underneathr installation of the battery. An extra-ounting surface is required because ofove one of the leg supports. The ele-

s proof tested to 8163 kg in both con-monstrate adequate design margin. Attegration and test, it was necessary tocraft and elephant stand together, aeighing in excess of 3447 kg. The ele-s designed with special pockets undero allow air bearings to be used to ac-sk. With a mere 35.6-N (8-lb) force, thelephant stand could be moved together,ween test chambers at Goddard SpaceSFC) during spacecraft environmental

pping container and transporter system move the assembled spacecraft. Size,ions, contamination issues, and modesn were important issues considered in

the hardware development. Aswith typical containers, shock andvibration isolation design featureswere incorporated to protect thespacecraft during transportation.The spacecraft was designed to belifted vertically and installed on thecontainer base by an overheadcrane. Four container sections, re-quired because of limited hookheight in the Kershner SpaceBuilding clean room, are lowered

96) 73


over the spacecraft to completely enclose it. Each sec-tion joint has an O-ring seal to provide an airtightconfiguration during transportation. The rotation fix-ture attaches to the enclosure and is designed to lift theloaded container, then rotate it 90° so that it mateswith the transporter (Fig. 15). Rotation to the horizon-tal position is necessary so that the hardware can betransported in the C5 aircraft (to Vandenberg Air ForceBase) or air-ride lowboy truck (to GSFC). Pressureinside the container is regulated within 1723.7 N/m2

of ambient pressure by banks of breather valves locatedin the base of the container. The sensitive opticalinstruments are protected from contamination duringchanges in external pressure while on the C5 aircraftby special HEPA-filters mounted inside the breathervalves to clean the incoming air. As with all groundsupport equipment, the shipping container, transporter,and rotation fixture were load tested before they wereused to verify that their capabilities were adequate. Thecontainer/transporter assembly was also subjected to aninstrumented 0.15-m drop test to verify the shock iso-lation system.

SUMMARYThe configuration developed for the MSX spacecraft

represents the largest structure ever assembled by theApplied Physics Laboratory. The design of each of thethree major structural components—the electronicssection, the truss structure, and the instrument sec-tion—was based on somewhat unique sets of require-ments involving packaging, structural, thermal, andstability constraints. These constraints resulted in theapplication of technologies in the areas of compositematerials and heat transfer. Two thermally stable, pre-cision alignment platforms—the optical bench and thebeacon receiver bench—were developed to providesupport for critical attitude control and radio-frequencysystem components. All structural components weresuccessfully tested and are currently integrated with theinstruments and spacecraft electronics awaiting launch.

74 JO

Rotationfixture

Shippingcontainer

Transporter

Figure 15. MSX shipping container and lift/rotation fixture.

REFERENCE1Jablon, A. R., and Persons, D. F., “Spacecraft Reflector Antenna ThermalDistortion Analysis,” in RF Expo West Proc. (18–20 Mar 1992).

ACKNOWLEDGMENTS: The MSX structural design and analysis efforts havebeen a major effort involving many APL staff and external participants. We wishto acknowledge the support provided by staff members in the Technical ServicesDepartment; Kelly Dodson, Helmuth Dorth and the team at Composite Optics,Inc., in San Diego, California; and Ted Sholar and other contributors from Swalesand Associates in Beltsville, Maryland. The MSX mission is sponsored by theBallistic Missile Defense Organization. This work was supported under contractN00039-94-C-0001.

THE AUTHORS

WILLIAM E. SKULLNEY is Supervisor of the Mechanical Engineering Group(S3M) in the APL Space Department. He joined APL in 1978 after receiving hisB.S. and M.S. degrees in engineering mechanics from The Pennsylvania StateUniversity. He is a member of the Principal Professional Staff and has been thelead structural engineer for the Delta 180, 181, 183, and MSX programs. He hasbeen a lecturer in the Space Systems Course offered by the G.W.C. WhitingSchool of Engineering since 1990 and has authored a chapter on structuraldesign and analysis in the textbook Fundamentals of Space Systems. His e-mailaddress is [email protected].



HARRY M. KREITZ, Jr., is a Senior Staff member of the Mechanical DesignSection of the APL Space Department’s Mechanical Engineering Group (S3M).In 1966, he received a B.S. degree in mechanical engineering from theUniversity of Maryland. Since joining APL in 1991, he has worked on the MSXprogram as the payload mechanical engineer. In addition to MSX, he is currentlyassigned to the Active Geophysical Rocket Experiment (AGRE) as the payloadmechanical engineer. His e-mail address is [email protected].

MARK J. HAROLD is a Senior Staff member of the Mechanical EngineeringGroup in the APL Space Department. He received a B.S. in civil engineeringfrom the Pennsylvania State University in 1979 and a master’s degree in businessadministration from the University of Baltimore in 1991. Before joining APL in1989, he was employed as a mechanical design engineer by Boeing MilitaryAircraft working on the 757, 767, and B52 aircraft, Grumman FlaxibleCorporation working on the 870 mass transit bus, and Martin Marietta Aero andNaval Systems working on the U.S. Navy’s Vertical Launching Systems andWide Aperture Array programs. Since joining APL, he has been the problemsponsor for the MSX truss structure and has worked on proposals involving theTACSAT, VIP, SPFE, GAMES, and CONTOUR spacecraft. He is currentlyworking as the assistant payload mechanical engineer on MSX and assisting onthe ACE spacecraft. His e-mail address is [email protected].

STEVEN R. VERNON is an Associate Staff engineer in the APL SpaceDepartment. He received his B.S. in mechanical engineering from The JohnsHopkins University in 1992. He is the MSX mechanical engineer responsible forthe instrument section, beacon receiver, and UVISI mechanisms and sunshades.His current duties are the lead mechanical engineer for TIMED, the CassiniINCA Sensor, the NEARS and MIDEX spacecraft proposals, and the develop-ment of several geophysical sampling devices. His e-mail address [email protected].

TERESA M. BETENBAUGH is a Senior Staff member of the Analysis and TestSection of the APL Space Department’s Mechanical Engineering Group (S3M).She received a B.S. in civil engineering from the University of Virginia in 1981and an M.E. in mechanical engineering from The Johns Hopkins University in1991. Since joining APL in 1989, Ms. Betenbaugh has performed structuralstrength and dynamic analysis and testing of the MSX spacecraft and itscomponents. She is currently assigned to the Advanced Composition Explorer(ACE) program as lead engineer for structural analysis and testing. Her e-mailaddress is [email protected].

JOHNS HOPKINS APL TECHNICAL DIGEST, VOLUME 17, NUMBER 1 (1996) 75


THEODORE J. HARTKA is an Associate Staff member of the MechanicalDesign Section of the APL Space Department’s Mechanical Engineering Group(S3M). In 1987, he received a B.M.E. degree in mechanical engineering fromThe Johns Hopkins University. Since joining APL in 1988, Mr. Hartka hasbeen involved in numerous spacecraft studies and proposals and wasresponsible for mechanical ground support equipment on the MSX program.He is currently assigned to the Near Earth Asteroid Rendezvous (NEAR)program as lead spacecraft mechanical engineer. His e-mail address [email protected].

DAVID F. PERSONS is a Senior Staff member of the Analysis and Test Sectionof the APL Space Department’s Mechanical Engineering Group (S3M). In 1977,he received a B.S. degree in mechanical engineering from the University ofCincinnati, and in 1979, an M.S. degree in the same discipline from theUniversity of Michigan in affiliation with the Chrysler Institute of Engineering.Since joining APL in 1982, Mr. Persons has performed structural analysis andtesting on a variety of spacecraft, spacecraft subsystems, and instruments. Hisactivities have included modeling thermal distortion of radar, ultraviolet, andvisible sensors; performing spacecraft, instrument, and subsystem stress anddynamics analysis; and vibration and strength testing. He is currently assigned tothe Near Earth Asteroid Rendezvous (NEAR) program as lead engineer forstructural analysis and testing. His e-mail address is [email protected].

EDWARD D. SCHAEFER is a Senior Staff member of APL’s Space Depart-ment. He holds M.S. degrees in mechanical engineering from NorthwesternUniversity and engineering management from George Washington University.Mr. Schaefer is a structural engineer with over 25 years of design and analysisexperience with marine, aircraft, and spacecraft structures and has served as theenvironmental test engineer for the MSX spacecraft. His e-mail address [email protected].

76 JOHNS HOPKINS APL TECHNICAL DIGEST, VOLUME 17, NUMBER 1 (1996)

Documents

Structural Design of the MSX · PDF fileSTRUCTURAL DESIGN OF THE MSX SPACECRAFT JOHNS HOPKINS APL TECHNICAL DIGEST, VOLUME 17, NUMBER 1 (1996) 61 sides by aluminum honeycomb panels