ANALYSIS AND VALIDATON OF THE STRUCTURE OF PRATHAM, INDIAN INSTITUTE OF TECHNOLOGY BOMBAY'S FIRST STUDENT SATELLITEAnki t C hiplun kar*, Ramanath Pai**, Anir u dh S u bra m anyam*** * De partment of Aer ospace Engineering, Indian Institute of Technology Bombay, P owai, Mumbai 400076. (E-mail:[email protected]) **Department of Mechanical Engineering, Indian Institute of Technology Bombay, Powai, Mumbai 400076. (E-mail:[email protected]) ***Department of Chemical Engineering, Indian Institute of Technology Bombay, Powai, Mumbai 400076. (E-mail:[email protected]) ABSTRACT 'Pratham', is a nano-satellite built by the students of IIT Bombay and is slated for launch by the Indian Space Research Organization (ISRO) in the third quarter of 2011. This paper discusses the wor k done by the Struct u res Sub-s y stem of Pra tham. T h e objective of the s u b-s y stem is to e n sure t h e robustness of the satellite structure so that it survives launch loads. A finite element model of the satellite structure has been made and representative launch loads have been applied. Various static and dynamic analyses have been performed on the satellite structure to obtain the response. Finite Element Analyses of the printed circuit boards (PCBs) onboard the satellite have also been performed. The FEA results h a ve been v a lidated in 2 stages: the g eometr y was v a lid ated by comparing with theoretical results while the element types were validated by comparing with an a lyses of isol a ted in dividual str u c tural e lements. Th e results suggest that th e satellite will maintains its structural integrity during launch and that no component of the satellite will fail during launch. KEYWORDS : Finit e ele ment a n a lysis; Launch load vi brat i ons; Microsatellite; Va lidation. INTRODUCTION „Pratham‟ is a microsatellite with a payload to measure the total electron count of the ionosphere and perform the to m ograp hy of i onosp h ere. Dur ing l aunch, Pra tham is e x pected to withstand static l oads of about 9 times the gravitational force and dynamic loads leading to vibrations. Under such harsh conditions, it is imperative that the structure of a satellite withstand these loads with little or no deformation so that the internal circuitry and actuators are not damaged and the satellite can execute the complete life cycle it was designed for. This paper discusses the various analyses that were carried out to ensure that the structure of the satellite satisfies the above requirements. The design approac h h a s been br iefly expla in ed. A CAD model of the sate llite is prepared a n d it is then m es h ed to obtain the Finite Element Model. FE models of the components onboard, including that of a typical Printed Circuit Board (PCB) are made. Simulation conditions, including simulation of launch loads and physical constraints have been discussed. Validation of the analysis is done by comparing with theoretical results for the same geometry and validation of the elements used for analysis is don e b y comparing with res ults th at wo uld be obta in ed f or so me stan dard g eometry.
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'Pratham', is a nano-satellite built by the students of IIT Bombay and is slated for launch by the
Indian Space Research Organization (ISRO) in the third quarter of 2011. This paper discusses the
work done by the Structures Sub-system of Pratham. The objective of the sub-system is to ensure the
robustness of the satellite structure so that it survives launch loads. A finite element model of the
satellite structure has been made and representative launch loads have been applied. Various static
and dynamic analyses have been performed on the satellite structure to obtain the response. Finite
Element Analyses of the printed circuit boards (PCBs) onboard the satellite have also been
performed. The FEA results have been validated in 2 stages: the geometry was validated by
comparing with theoretical results while the element types were validated by comparing with
analyses of isolated individual structural elements. The results suggest that the satellite willmaintains its structural integrity during launch and that no component of the satellite will fail during
launch.
KEYWORDS :
Finite element analysis; Launch load vibrations; Microsatellite; Validation.
INTRODUCTION
„Pratham‟ is a microsatellite with a payload to measure the total electron count of the ionosphere and
perform the tomography of ionosphere. During launch, Pratham is expected to withstand static loadsof about 9 times the gravitational force and dynamic loads leading to vibrations. Under such harsh
conditions, it is imperative that the structure of a satellite withstand these loads with little or no
deformation so that the internal circuitry and actuators are not damaged and the satellite can execute
the complete life cycle it was designed for. This paper discusses the various analyses that were
carried out to ensure that the structure of the satellite satisfies the above requirements. The design
approach has been briefly explained. A CAD model of the satellite is prepared and it is then meshed
to obtain the Finite Element Model. FE models of the components onboard, including that of a
typical Printed Circuit Board (PCB) are made. Simulation conditions, including simulation of launch
loads and physical constraints have been discussed. Validation of the analysis is done by comparing
with theoretical results for the same geometry and validation of the elements used for analysis is
done by comparing with results that would be obtained for some standard geometry.
Pratham is expected to satisfy the following requirements :
1. Launch Vehicle Placement: The satellite is launched into Low Earth Orbit by the Polar
Satellite Launch Vehicle. The launch vehicle interface to be used is the IBL230V2, to be
provided by VSSC.
Launch vehicle interface requires 8 M6x1, 9mm long helicoil inserts at 230mm PCD on
bottom deck of the satellite.
There may be no interference in the joint from the satellite to the launch vehicle body.
2. Launch Loads : The satellite is carried to its orbit by a launch vehicle in a flight lasting
about 17 minutes. The high levels of acceleration, vibrations and shocks experienced by the
vehicle during this period are transmitted to the payloads attached to the flight decks of thevehicle and impose strict requirements on the overall structure. The loading specification for
which the launch vehicle interface is tested is assumed to be the loading data for the satellite
during launch and has been specified in the IBL230V2 documentation.
Static Loads: Lateral loads are considered to act simultaneously with longitudinal loads.
Earth‟s gravity is also included in the levels given below. All loads apply at the centre of
gravity of the satellite as body forces. The longitudinal and lateral axes are defined later.
Direction Loading Longitudinal ± 11g
Lateral ± 6g
Table 1: Static loading levels
Harmonic Loads: The qualification loading levels are used for analysis of the satellite
Quadratic quadrilateral target Targe170 3D target segment
Quadratic triangular contact Conta174 3D 8 node surface to surface contact
Quadrat ic triangular target Targe170 3D target segment
Table 4: Elements used to mesh the satellite structure
Boundary conditions: The part of launch vehicle interface attached to the satellite is constrained in
space. This is a correct assumption as in actual launch it will be constrained using ball lock
mechanism constraining all the degrees of freedom.
1. Static Analysis of QM:
Loads applied: As specified in Table 1.
Results:
Maximum von Mises stress (occurs on a washer) 20.51MPa (4.6% of failure
stress)
Maximum total deformation (at the tip of the
middle antenna)
0.214mm
Table 5: Results of Static Analysis
A buckling analysis was also performed on the satellite. A buckling load factor of more
than 100 was obtained for the solar panels and this immediately ruled out the possibility
of buckling as a cause of failure.
The maximum stress obtained in the analysis is far less than the yield strength. Also, total
deformation is far too less to for contact between 2 surfaces to occur.
After being loaded with maximum static loading levels possible during actual launch, wecan say that the structure will maintain its integrity. Hence the structure will not fail under
static loads even in such a worst case scenario.
2. Modal Analysis of QM:
Loads applied: NIL
Results: The first 10 natural frequencies are presented below:
Table 10: Comparison of Experimental modes and simulation modes
The first mode cannot be observed experimentally due to inaccurate positioning of the accelerometer.
We may conclude that the overall satellite model is correct and there is litt le or no error in the
analysis.
Validation of the satellite model
In this geometry, since the antenna is major contributing factor in stiffness characteristics of the
model. The modal frequencies of the QM are compared to those obtained from antennae alone. Themonopole holder was fixed and first two resonant frequencies were obtained. Since the first 6 modes
of the structure are antenna modes, the 2 modes extracted from this model serve as a good
representative of the first 2 QM modes. From a structural point of view, the monopole antenna
attached to the satellite is similar to a cantilever beam. The formula for the natural frequency of a
cantilever beam is given by:
Where f is the natural frequency of the cantilever beam, E is the modulus of elasticity, I is the
bending moment of inertia, L is the length of the cantilever beam, g is the gravitational constant, w is
the weight of the cantilever beam. The result obtained using this formula is used to validate the
element.
Mode QM frequency (Hz) Antenna model frequency (Hz) Theoretical
1 152.46 164.26 166.85
2 154.45 164.56 -
Table 11: Comparison of QM and antenna model
The close matches in the modal frequencies indicate that the entire satellite model is correct.