Strength Analysis - Finite Element Analysis for Satellite Structures - page 186

• Coefficient of thermal expansion (a) = 24.7 9 10 -6 m/m/ C
• Assign the locations of mounting the high precise equipment to calculate
thermal deformation, and hence mounting accuracy.
• Equipment support is designed to eliminate thermal loading as they allow
sliding of the support under thermal expansion. Therefore, in the simplified
thermal model, all satellite equipment are removed to avoid the appearance of
any artificial thermal stresses that may result from improper representation of
the sliding support.
• The basis unit block module is meshed according to the same criteria
implemented in Sect. 4.3 for fine meshing, while the rest of structural modules
are meshed with coarse meshing. ANSYS and SOLID92 elements are used in
the meshing process.
• The nodes located at connection areas between different structural modules
are coupled at all DOF.
2. Apply the loads and define boundary conditions: Thermal deformations are
evaluated for the worst cases, which are obtained from the on-orbit thermal
analysis for the satellite structure. Thermal satellite engineers should perform a
complete on-orbit thermal analysis of the satellite, and supply input data for
surface temperature gradients of different satellite structure modules. This data
are the input data needed to perform an on-orbit thermal deformation analysis.
For Small Sat, the satellite structure is solved thermally under on-orbit cyclic
thermal loads. The results of this data are obtained from the research project
titled ''Prediction of satellite structure life on-orbit under thermal fatigue effect'',
National Authority of Remote Sensing and Space Science, NARSS, Egypt, 2006
[ 6 ]. Input data for surface temperature gradients of different satellite structure
modules of Small Sat are listed in Table 4.26 . It contains both maximum and
minimum average temperatures for each structural module (Table 4.27 ).
Displacement boundary conditions are the DOF constraints usually specified at
model boundaries to define the structural support points. During on-orbit oper-
ation, the satellite is totally free without any fixation points. To conduct a ther-
moelastic analysis, displacement boundary conditions must be defined for the
related model. Therefore, to simulate satellite thermal deformation due to on-
orbit cycling, the satellite model must be constrained by rigid support points.
Selection of these points' locations and their fixation manner should provide
minimum effect on satellite deformation. For the Small Sat model represented at
Fig. 4.61 , the support points are selected at the plane of the star sensor mounting
with its bracket. This plane is selected, because the mounting accuracy of all
precise equipment installed at the basis unit block is measured relative to the star
sensor bracket. Displacement boundary conditions are applied by fixing one of
the four star sensor points of fixation (on bracket) in the ''X'' direction, the next
point in the ''Y'' and one of the other in the ''Z'' .
3. Set solution controls and solve the analysis: By reviewing the input data listed
in Table 4.26 , some of the satellite structural modules are divided into more
than one division according to their position. Each division has its maximum