Structural Finite Element Analysis of Stiffened andHoneycomb Panels of the RASAT Satellite
S. Onta,', S. Dag2, MI. Gokler3'TUBITAK Space Technology and Research Institute, suat.ontacguzay.tubitak.gov.tr
METU Campus, 06531, Ankara, TURKEY2 Mechanical Engineering Department
Middle East Technical University, sdaggmetu.edu.tr0653 1, Ankara, TURKEY
3 Mechanical Engineering DepartmentMiddle East Technical University, goklergmetu.edu.tr
0653 1, Ankara, TURKEY
Abstract SFF and EFF are meshed with hex elements and theThis paper describes the structural analysis carried out on honeycomb panels are meshed with solid brick and shell
the main stiffened and honeycomb panels of the RASAT elements. For the calculated gRMS value the static analysis hadsatellite. The analysis here supports the design process and been conducted in each axis of the panel assembly. For theaims to ensure that the panels survive structural qualification dynamic case, the same finite element mesh and materialtesting. This analysis therefore forms part of the overall properties had been used. In this case, the boundary conditionsqualification process. The stiffened and honeycomb panels are applied in such a way to determine the mode shapes and thebeing considered in this document form the outer box structure resonance frequencies. Furthermore, the stress values had beenof the satellite. These panels consist of the space-facing facet determined with respect to the applied static and dynamic(SFF), solar panels including solar cells and earth facing facet loading cases. They had been compared with the allowable(EFF). All these panels are key parts of the satellite's structure stress values of the materials. In this paper the complete finiteand are critical to mission safety. The separation panel is element analyses procedures are described and the results ofparticularly highly loaded, since it supports the battery pack, the analyses are presented. According to the computed results,reaction wheels, gyro module, magnetorquer rods and sun some conclusions are drawn in order to guide experimentalsensors. The separation panel also supports the solar panel qualification tests.assembly. The solar panels are also of critical importance, theirintegrity maintaining the required power supply to operate thesatellite's electronic systems. As being different from the SFF I. INTRODUCTIONand EFF, the solar panels are made of aluminum honeycomb The first step in the analysis is the preparation of solidpanels. The solar panels are particularly sensitive, as they carry models of the honeycomb and stiffened panels. For thisarrays of delicate ceramic solar cells together with their wiring. purpose, CAD models of the panels are developed using aThroughout all loading conditions experienced during the CAD software. Models for each of panels are created and atmission, the solar panels must continue to support the solar the end all of these panel models are assembled in order tocells without cell failures or wiring disconnections. The EFF is demonstrate the whole satellite outer skin. In the modelling ofperhaps the least critical of the stiffened panels but still must honeycomb panels, the outer and inner skins of thesupport the top of the solar panel assembly and must carry honeycombs are modelled as a sheet and the core part isvarious antennae. modeled as a solid. The space-facing facet and the earth-facingThe main objective of this study is to assess the strength and facet are modeled as solids. The whole assembly model is
vibration response properties of the stiffened and honeycomb shown in the Figure 1.panels by conducting static stress and modal analyses. For the After completing solid modelling, the model was convertedcase of static loading, the reliability can be estimated with great to a parasolid file and transferred to the finite element analysisefficiency, whereas for dynamic loading the performance software.depends on the considered frequency range. The obtainedresults are very significant in that, they illustrate the feasibilityof a full scale analysis for structural reliability in a designcontext for large-scale structures. The analyses are conductedby means of the finite element method. For the static case, the
* 700 x 700 mm square panel* 24 mm overall thickness* Some parts of the inner side are emptied in order to
decrease the weight and keep the stiffness
EFF specifications are as follows:
F 700 x 700 mmmsquarepanel* 20 mm overall thickness
p Some parts of the inner side are emptied in order todecrease the weight and keep the stiffness
m cut-outs for Hi-Res Camera and main stack coolingpurpose
ThegpropertiAssemblysloft honeycombps ar end a Figure 2 shows a view of the FE model of the closed cube
Fiurl1oAs:em ly od o ne ma structure formed by the 4 main honeycomb panels and 2stiffenedpanels of RASAT stiffened panels. The panels shown in the figure represent onlypart of the whole-spacecraft FE model.
II. FINITE ELEMENT MODEL
Each of the aluminum honeycomb and stiffened panels wasinvestigated by analysing either all or part of a whole-spacecraft model. This model included a stack configuration orrepresentation that has been validated for various previousspacecraft FE models. It is important to characterise thestiffness of the stack assembly accurately as it governs thefundamental behaviour of the spacecraft as a whole.
The modelling methodology used for all honeycomb panelsis as follows: The core material was represented by tetrahedron3-D elements. For each panel, the two main outersurfaces(inner and outer skins) are meshed with 2-D shells. Allnodes of the shell elements (representing the skins) and the tetelements (representing the core) were equivalenced. Tocharacterise the behaviour of the aluminum honeycomb corematerial, a 3-D orthotropic material definition was used, asdetailed later.
The properties of the honeycomb panels are given as Figure 2. Finite element model ofhoneycombfollows: and stiffened panels ofRASAT
III. MATERIALS* all panels have same outer envelope Grade 2014A aluminum alloy is used for all honeycomb* 700 mm top (EFF) edge length, 700 mm bottom panel skins. Elastic properties of this Aluminum alloy are given
(SFF) edge length, 554 mm vertical edge length as* 15 mm overall panel thickness* 0.4 mm inner skin; 0.7 mm outer skin (solar cell elastic modulus, E =70000 MPa
surface) thickness Poisson's ratio, v= 0.33* ±Y Solar Panels have elliptical cut-out (for Star density,;+1,pIr Q1 )1
280 k/mm311
benchmark FE model tests have shown that, as long as thematerial model parameters give an appropriate disparity in thedifferent axes, the precise value of each parameter makesrelatively little difference to the analysis results. Thus, theindividual parameter values are nominal. Given that axis 1 runsin the ribbon direction, axis 2 runs in the tangential directionand axis 3 is normal to the plane of the sheet of core material, ,S i ;the material parameters used are as follows:
v3I = 0shear modulus, G12 1 MPaG23 = 220 MPaG3 1 = 440 MPa Figure 3. Point masses on the SFF (Some solarp = 83.3 kg/mm3 panels and the EFF removed for clarity)max allowable shear strain = 2.8 MPa (from data sheet)
The SFF and EFF panel materials are grade 7075-T6aluminum alloy. Its elastic isotropic properties are as follows:
elastic modulus, E =72000 MPaPoisson's ratio, v= 0.33density, p 2810 kg/mm3minUTS ;570MPa
IV. LOADS
To investigate strength, a general loading approach has beenfollowed in all FE modelling. The aim is to design to allow for mj9 Wiamfifthe highest reasonably likely instantaneous inertial loadexperienced by the spacecraft, factored to qualification levels.For the launchers likely to be used, the qualification-levelrandom vibration loading is the most severe, this being definedby a spectrum of 12 gRMS. A statistical study will show thatfive times this integrated level gives the highest likelyinstantaneous applied acceleration level (at the spacecraft, orsubsystem, base) to 98% certainty. Therefore, all FE models Figure 4. Point masses on the EFFare assessed for strength by applying a static inertial load of60g in each axis separately. This value thus includes aneffective margin of safety by virtue of the original random v. ANALYSIS AND RESULTSspectrum being factored for qualification levels. Also, The FE model of the spacecraft was analysed in differentconsidering that real failures take finite time to develop, ways in order to understand the behaviour of the honeycombapplying a static inertial load to an FE model is more severe and stiffened panels. The panels were analysed for modalthan the actual situation of occasional transient, momentary frequencies and stress. With regard to the Solar Panels, resultsacceleration peaks. from the normal modes analysis were used to estimate likely
In order to represent the masses on the stiffened panels some peak strains in the outer skins.point masses are located on the SFF and EFF. For instance onthe SFF there are battery pack, torque bars, reaction wheels and The behaviour of the SFF and EFF were determined bysome electronic module boxes. Similarly, on EFF there are analysing the whole spacecraft model. Both normal modes andsome antennas which have some masses. All these point mass static stress analyses are carried out. Normal modes and staticrepresentations are shown in Figure 3 and Figure 4 for the SFF stress analyses were also carried out for the Solar Panels. Inand the EFF respectively, this case only the cube structure formed by the four main
honeycomb panels are analysed.
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1. Analysis of Stiffened Panels (EFF and SFF) B. Stress Analysis
The SFF and EFF were assessed by analysing the whole The strength of the SFF was then investigated. Again, thespacecraft model (including Separation System, Stack, Shear whole spacecraft FE model was analysed with a static 60gPanels and other equipment). Both normal modes and inertial inertial load in each of the x, y and z axes separately. For eachstress were of interest. The SFF and EFF together must support of these three load cases, the Separation System fixingthe Solar Panel with enough stiffness to give a frequency which boundary conditions used above were retained, together withdoes not couple with the whole spacecraft. Then, in terms of an inertial load in the appropriate direction.strength, the SFF is the most critical in that it carries severalheavy components, including the Battery Pack, Reaction The results of these three static stress analyses suggested aWheels, Torque Bars and some electronic modules. This panel maximum local effective stress in the SFF of 161 Mpa asmust be strong enough not to fail under inertial loads arising shown in Figure 6. This is much less than the available UTS offrom vibration and other loads. 570 MPa of the 7075-T6 aluminum material. As expected, this
was near the Battery Pack and corresponds to the flexing of theA. Normal Modes Analysis corner of the panel predicted by the normal modes analysis.
The EFF was considered next. For this panel, maximumFixed boundary conditions were applied at nodes around the effective stress for any of the three inertial load cases was
lower flange of the Separation System (launcher interface) at predicted to be 60 MPa.. This value suggests a clear strengthtwelve bolt locations, as shown in Figure 5. The whole margin for the EFF.spacecraft model was then run in normal modes. This analysisshowed that the first mode involving the SFF was predicted to EquivaIleht (von-Mis6s)5tressbe at a frequency of 86 Hz. This mode showed the Battery Max L6126+002Pack causing that corner of the Separation Panel to flex. As the MihX 4:5236-002corners of the Separation Panel are not constrained to the rest 161,241of the Spacecraft (i.e. the Solar Panels) and are thus free to 143330bend like a cantilever, this mode was expected. However, the 125419predicted 86 Hz is not likely to adversely couple with the first 107509axial mode of the spacecraft as a whole, thus it is adequately 8 9
separated in terms of frequency. No other SFF modes werecauses for concern. Moving onto the EFF, this panel is not as
-93f777highly mass-loaded as the SFF and does not contribute so much
to the stiffness, and thus to the normal modes, of the spacecraft. 35,866
0,045 . _
Figure 6. Von Mises Stress Distribution on theSFF
2. Analysis ofthe Honeycomb Panels
The Honeycomb Panels were also analysed for both normalmodes and static stress. For the purposes of these Solar Panelinvestigations, only part of the whole spacecraft model wasexamined. The closed box structure formed by the four mainhoneycomb panels, as shown in Figure 2, was considered. Mostcritical for the Solar Panels is the integrity of the solar cellsbonded to the outer skins of all four panels. No loading
Figure 5. Fixed Boundary Conditions at environment should lead to the failure of any of the adhesiveSeperation System Connection on the SFF side bonds, cause fracture of a solar cell or result in the damage to
the associated wiring. Thus, neither should the panels
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themselves fail and nor should solar cells or wiring become conditions as above were used, together with a 60g inertia loaddebonded or fail. These conditions are ensured by controlling separately in each of the main axis directions. Experience atvibration response and thus preventing excessive coupling specraft design suggests that strength itself isn't a key issuewhich might otherwise cause large panel amplitudes and with solar panels, this being supported by the FE results here.accelerations and thus displace solar cells. Solar cell security is Maximum effective stress in the Solar Panel skins wasalso ensured by controlling maximum strain in the outer panel predicted to be 34 MPa, well below the UTS of the 2014Askins during vibration. Finally, stress in the Solar Panels aluminium alloy material.arising from peak inertial forces should also be low enough notto cause failure of the panels. VI. CONCLUSIONS
A. Normal Modes Analysis The analyses carried out by considering the four mainaluminum honeycomb panels and two stiffened panels suggest
The normal modes analysis was used both to determine the that all panels are stiff enough to ensure adequately high modalnatural modes and frequencies of the solar panels and to frequencies to meet good design practice and also to avoidprovide results for the estimation of Solar Panel skin strain. For vibration coupling between the panels and the adjoiningthis analysis of the honeycomb box structure, fixed boundary spacecraft structure. The SFF and EFF contribute to a suitablyconditions were applied at locations on the SFF and EFF. high first axial modal frequency of the BILSAT of 129 Hz,These locations correspond to the points where the structure is which may be useful for predicting RASAT first axial modalconnected to the rest of the spacecraft. Thus, twelve equally frequency. The results of the analyses presented in this paperspaced groups of nodes around the upper flange of the also show that the first SFF mode is 86 Hz, where the heavySeparation System PCD on the SFF were constrained in all six Battery Pack causes the free corner of the panel to flex.degrees of freedom. In addition, nodes on the EFF However, this mode will not adversely couple with the abovecorresponding to the bolt locations around the square footprint first axial mode of the spacecraft.Of the shear panels were constrained.
The normal modes analyses also demonstrate that the SolarThe first mode given by the analysis was ignored as it Panels have a safe first mode of 189Hz, the+Y panel showing
showed the four Solar Panels moving up and down as an the largest deflections. Using the results from this mode it isassembly on flexing SFF and EFF. No significant distortion in possible to estimate the peak strain in the outer skin of this keythe Solar Panels themselves was shown. The second mode was panel.of more interest, showing the Solar Panels flexing. This modeoccurred at 189 Hz. Past experience of designing spacecraft Looking at the various stress analyses, all four honeycombsuggests that this is an adequately high first solar panel mode, and stiffened panels were predicted to experience sufficientlyavoiding the risk of dynamic coupling with the adjoining low effective stress and strain levels. Subjected to 60g inertialstructure of the craft. loads separately in the three axis directions, the SFF was
predicted to develop a maximum stress value of 161 MPa. TheNext, other data from the results of this normal modes margin in the applied qualification-level loads guarantees
analysis was used as a basis for estimating the peak strain additional factor of safety. The EFF is not as highly loaded andoccurring in the Solar Panel skins during launch vibration. is shown to develop a peak effective stress of 60 MPa in theInspecting the mode shape, the FE model showed the +Y Solar skins and a peak maximum shear stress in the core material ofPanel to experience the most pronounced distortions and thus 15 MPa. These values are again satisfactory and show goodstrains. This was as expected, as this panel has a Star Camera margi of strengthcut-out. Furthermore, the highest skin strain was shown to be atthe periphery of the Star Camera cut-out. Taking strain results Finally, the analyses showed that the strength of the Solarfrom the analysis and using the strain calculation method the Panels was adequate with a very good margin as well. A peakhigmthet lelysssrain uingtheskinwas calculateon be
theffective stress value anywhere in the skins was predicted to be
highest likely strain in the skin was calculated to be 475 ptc. 34 MPaThis compared to a similar value of 474 gt£ which has asuccessfully passed previously tested structural qualificationmodels, where a partially populated solar panel survived REFERENCESvibration testing with no problems. Therefore, that comparison
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