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CHAPTER 1
INTRODUCTION
1.1
ANUSAT (ANna University SATellite)
During the year 2002, Indian Space Research Organization (ISRO)
threw an open challenge to the educational institutions in India to take up a
project on the development of microsatellite mainly to bring in awareness in
team spirit and work force development towards satellite building in the
country. In response to ISRO’s challenge Anna University, Chennai was the
first university in the country to prepare and submit a proposal towards
development of a microsatellite in university environment. After careful
scrutiny of the proposal by ISRO, the proposal of the project has approved for
a total cost of Rs 5.44 crores. Objectives of this Satellite development in Anna
University, Chennai are (i) to establish team of expertise in the development
and usage of microsatellite, (ii) to provide qualified and trained work force for
India’s space programme and (iii) to initiate research activities towards
development of microsatellite.
After getting a formal approval from ISRO, task teams were formed
to take up various activities viz. (i) Structure, (ii) Thermal, (iii) Orbital
Mechanics, (iv) Attitude control, (v) Bus electronics, (vi) Payload, (vii)
Power, (ix) Integration, (x) Mission, (xi) Communication, (xii) Telemetry and
Telecommand, (xiii) Antenna, and (xiv) Ground station. Task team members
looked after these activities and the Project Director did the co-ordination of
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all the activities. This development has been completed successfully and the
microsatellite (ANUSAT) developed by Anna University Chennai is in orbit
since April 20, 2009 and the health of the satellite is being monitored from the
Ground Station established at MIT. All the subsystems of the satellite are
functioning as per the prediction, which has seen from the website
http://www.annauniv.edu/anusat. Without the dedicated efforts by the task
team members and the people associated with the development, the satellite
would not have gone to the space. All the task team members, project staff
and students participated and contributed to the development of this satellite
have benefited immensely and the university would be in a position to
develop similar microsatellites with specific mission in the future as it has
acquired in depth knowledge and experience in the development of
microsatellite.
The structural design is one of the important components of the
microsatellite design and this thesis mainly deals with structural design.
1.2
MOTIVATION
TOWARDS
MICRO-SATELLITE
STRUCTURE
Small satellites have literally been around since the dawn of the
space age.
However, the success of trunk communications via satellite,
coupled with manned space has forced the space industry towards ever larger
and more expensive missions. Small, cheap satellites used to be the exclusive
domain of scientific and amateur groups.
They provide cost-effective
solutions to traditional problems at a time when space budget is decreasing.
Interest
in
small
satellites
is
growing
fast
worldwide.
Business
Establishments, Governments, Universities and other organizations around
the world are starting their own small satellite programmes.
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Micro-Satellites (Spacecraft) fall into the 10-100 kg category of
small satellites. A reliable structural design with sufficient space to carry all
necessary components is essential for the success of a spacecraft’s mission. It
is also essential to limit the mass and size of the spacecraft in order to reduce
the cost for sending spacecraft into space. These major requirements drive the
overall structural design of every spacecraft. The structure has designed to
withstand, the static, dynamic and thermal stresses that occur during launch,
deployment and service without failure or excessive distortion. Besides that,
the structure also has to secure the payload and the most sensitive electronic
parts against excessive distortions, vibrations, temperature changes and
undesirable radiations.
1.3
DEFINTION OF SATELLITE
Literally, a satellite is a celestial object orbiting a larger body in
periodic fashion because of gravitational attraction. Since the beginning of the
space adventure, a distinction has made between natural satellites like the
Moon and artificial satellites. However, the term “satellite” which is used now
commonly describes the man made variety.
Satellites have built for a broad variety of missions, so each one is
different. However, just as any motor-driven vehicle has a chassis, an engine,
fuel tanks and steering systems, all satellites share the same basic structure
and organization. Depending on its mission, a satellite may carry a number of
instruments, to acquire images, record data, and to transmit and receive radio
signals. It generates its own power from solar panels.
An Earth satellite generally orbits at an altitude of between 450 and
36,000 km. Its motion has sustained naturally by the gravitational attraction.
However, it also has its own propulsion system for orbital maneuvers, which
in some cases may include boosting the satellite to its final orbit. An onboard
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computer that manages all the equipment and communicates with the ground
stations controls all these systems.
Satellites
today
provide
society
with
everything
from
environmental scientific data to global telecommunications services, resulting
in a multibillion-dollar industry. The field of satellite design plays a direct
role in shaping the small satellite industry. A satellite structure must fulfill
various requirements. First, it must resist the loads induced by the launch
environment (acceleration, acoustics, thermal), meet all the functional
performances required on orbit such as dimensional stability. For example, it
must also interface with some other subsystems such as thermal control,
optical components, electronic equipment, mechanisms, etc. In addition, the
structure will be the skeleton used during the assembly process of these
subsystems into the satellite and then it must provide very clean interfaces to
each individual element in order to simplify the sequence of integration.
Finally, the concept must be compatible with the standard manufacturing
process and use standard components (sheet-iron, tube) every time it is
possible.
The earliest satellites were small but as time went on, the satellites
that flown were developed to serve several different missions and they
became larger and more expensive and took a long time to design, build and
launch. A failure of the whole system meant the death of many different
projects. The future is likely to see more small satellites, each of which is
dedicated to a particular mission objective and carries a single instrument.
In 1957, the former Soviet Union was the first country to launch a
man made object into Earth orbit. Limited launcher capability was the main
constraint on satellite size and mass at that time. In the years that followed,
the size, complexity and cost of satellites grew, as did the capabilities of the
launch vehicles.
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The first microsatellites have built by enthusiasts of the amateur
radio community and launched in the early 1960s. The invention/introduction
of the microprocessor in the 1970s represented a quantum jump for the
onboard capabilities of a spacecraft. This technology introduction represented
a prime catalyst in the development of microsatellites since it enabled small
physical structures in support of sophisticated data handling applications. The
engineering of microsatellites, which emerged in the early 1980s, took a
radical change of approach from the custom design of traditional spacecraft,
namely a design-to-capability scheme to achieve cost reductions by focusing
on available and existing technologies using a general-purpose bus and ‘offthe-shelf’ components and instruments.
1.4
TYPES OF SATELLITES
1.4.1
Based on the size
Table 1.1 Types of satellite based on mass
S. No.
Name
Mass
1
Large satellite
>1000kg
2
Medium sized satellite
500-1000kg
3
Mini satellite
100-500kg
4
Micro satellite
10-100kg
5
Nano satellite
1-10kg
6
Pico satellite
0.1-1kg
7
Femto satellite
<100g
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1.4.2
Based on the mission
The satellites have classified based on their missions also as below,
1.4.3
1.4.4
1.
Astronomy satellites
2.
Atmospheric studies satellites
3.
Communication satellites
4.
Navigation satellites
5.
Reconaissance satellites
6.
Remote sensing satellites
7.
Search and rescue satellites
8.
Space exploration satellites
9.
Weather satellites
Based on the orbit on which it is placed
1.
Geosynchronous satellites
2.
Geostationary satellites
3.
Polar Satellite
4.
Sun-synchronous satellites
Based on the stabilization
1.
Three axis stabilized
2.
Spin stabilized (single spin and dual spin stabilized)
3.
Long booms
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1.5
SPACECRAFT STRUCTURES
Spacecraft structures have mainly divided into three main
categories and shown in Figure 1.1 of the model satellite.
1.5.1
Primary structure
The purpose of the primary structure or main structure is to transmit
loads to the base of the satellite through specifically designed components
(central tube, honeycomb platform, bar, truss, etc.). This structure provides
the attachments points to the launch vehicle for the payload and the associated
equipments. Failure of the primary structure leads to a complete collapse of
the satellite.
1.5.2
Secondary structure
The secondary structures such as baffle, thermal blanket support
and solar panel must only support themselves and has attached to the primary
structure, which guaranties the overall structural integrity. A secondary
structure failure is not a problem for the structural integrity, but it could have
some important impacts on the mission of the spacecraft if it alters the
thermal control, the electrical continuity, the mechanism or if it crosses an
optical path.
1.5.3
Tertiary structure
The interconnecting elements, brackets and the electronic boxes
form the tertiary structure of any satellite.
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Figure 1.1 Model satellite-showing categories of structures
1.6
LAUNCHING SYSTEMS FOR SMALL SPACECRAFT
In addition to the large range of launch vehicles available, several
different methods exist for the actual integration of spacecraft to the launch
vehicles. Present day launch vehicles are capable of launching big and small
satellites simultaneously. Small satellites has integrated to and subsequently
deployed from the launch vehicle in many different ways. This is an inherent
ability due to the small size in terms of mass and physical dimension of the
small satellite. There are essentially four launch configurations available to
the small spacecraft operator.
1.
Piggy-Back (Auxiliary) Payload
The ‘Piggy-Back’ or ‘Auxiliary’ launch is probably the most
commonly used method. The spacecraft is located alongside of the primary
payload in the excess volume that, in most launchers would contain a dummy
payload to create the appropriate mass distribution within the launcher. This
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means that such a launch is generally very cost effective. However, the major
drawback of this type of launch is that the primary payload often determines
the orbital injection point of the small satellite sitting in the piggyback
position.
2.
Multiple Spacecraft Dispenser
In mission specific dispenser, several spacecraft has integrated and
launched on a single launch vehicle.
3.
Secondary Payload
A third method involves two payloads, which are typically stacked
one on top of the other within the launcher fairing. This ‘stacking’ can take
several forms.
4.
A fourth, typically financially unrealistic, option is that of being the
primary or sole launch vehicle customer.
1.6.1
The Launch Vehicle: PSLV
The Polar Satellite Launch Vehicle (PSLV) has proposed to be the
launch vehicle for the microsatellite. PSLV is a four-stage launch vehicle
primarily designed to inject 1000 kg class spacecraft into a 900 km SunSynchronous Polar Orbit (SSPO), when launched at a nominal azimuth of 140
degree from Sriharikota (SHAR) located 80 km north of the city of Chennai,
India. In addition, the vehicle can also launch passenger payloads in 50-150
kg class into the orbit with main payload. The vehicle has designed and
developed by Indian Space Research Organization (ISRO) with the
participation of Indian industries and institutions. The first developmental
flight took place in September 1993 followed by two successful
developmental flights in October 1994 and March 1996. The first operational
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flight carrying 1200 kg remote sensing satellite (IRS-1D) took place in
September 1997. Subsequently provisions have made on the Vehicle
Equipment Bay (VEB) to carry two Auxiliary satellites along with the main
satellite. The first flight carrying Auxiliary satellites, Kisat-3 of South Korea
and DLR-Tubsat of Germany, along with main satellite, Oceansat-1 of India
has successfully launched in May 1999. The PSLV also undertake launches to
Low Earth Orbits (LEO) as well as Geosynchronous Transfer Orbits (GTO).
The Figure 1.2 shows the PSLV C-12 launcher configuration showing all the
stages and the satellite locations.
1.6.2
Vehicle Axes Definition
The PSLV sign convention in flight has shown in Figure 1.3. Pitch,
Yaw and Roll motions has indicated in their respective positive directions.
The vehicle axes system at the time of the launch has defined as follows.
Origin
:
Centre of gravity of the vehicle
Positive yaw axis
:
Along the direction of the launch azimuth (X-axis)
Positive pitch axis :
Points towards the left when looking along the
direction of the launch azimuth (Y-axis)
Positive roll axis
:
Towards vehicle nose, (Z-axis) Rotation about the
positive vehicle axes is positive attitudes.
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Figure 1.2 PSLV-C12 Configurations
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Figure 1.3 PSLV axes definition
1.7
SPACE MISSION ENVIRONMENTAL LOADS
1.7.1
Static or quasi-static external loads
These loads have related to the acceleration forces applied on the
structure and component during launch of the satellite in the supposed orbit.
1.7.2
Static or quasi-static internal loads
Some of these types of loads could be compressive in nature due to
driver components such as magnetometer, pre-stresses in mechanical
components and thermo-elastic stresses.
1.7.3
Dynamic external loads
Some of these types of loads could be engine thrust, sonic forces
caused by air turbulence effect on the launcher (gust), impulse thrust for
orbital adjustment and loads caused by transportation from manufacturing site
to the launching site.
1.7.4
Dynamic internal loads
Forces caused by internal motion for satellite attitude control like
momentum wheel, magnetic torquer or any other type of electric servomotors.
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1.8
GENERAL REQUIREMENTS AND CONSTRAINTS
The main objective of the spacecraft structural design is to provide
better and safe housing for all the spacecraft components from ground
transportation to space transportation. The configuration design and layout for
the spacecraft depends on the following requirements and constraints.
The micro-satellites normally carried as piggy-bag of the large
satellites by the launch vehicle. The overall dimension of the microsatellite has limited by the launch vehicle requirements.
Sufficient area and volume required for mounting all the subsystems of
the spacecraft.
To provide necessary interface between the Microsatellite and the
launcher interface ring.
Need to provide adequate space between the subsystems for the wire
harnessing.
Functional requirements of the subsysems in the spacecraft have to
maintain. For example, the torquers have kept away from the
magnetometers to avoid the electro-magnetic field coupling; one
torquer has aligned along the spin axis and the other torquer
perpendicular to it.
To provide adequate radiation pattern, the UHF and VHF antennae are
to be mounted on the top of the spacecraft.
The twin-slit sun-sensor is to be located out side the spacecraft and free
from any shadows, determines the spin rate and the Sun Aspect Angle.
To meet the integration requirements, the components like umbilical
connector, interface ring and ground checkout connectors have to keep
at the bottom of the spacecraft and closer to the vehicle interface.
To meet the mass, Centre of Gravity and Moment of Inertia
requirements.
To provide necessary stiffness and strength values to resist static and
dynamic loads.
Sufficient area for the solar cells is necessary to generate the required
power.
Thermal management system is required to maintain the temperature of
all the subsystems in the satellite during operating and non-operating
conditions.
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1.9
SCOPE AND OBJECTIVE
The scope of this research work is to design, analyze and develop
the micro-satellite structure of mass less than 50kg and to accommodate 50
subsystems of various mass and dimensions. The preliminary conceptual
shape design of the satellite structure has done and the internal structural
configuration has selected by the analysis made for mass and frequency
values from among the 11 configurations. The selected configuration with the
subsystems went through the mass location analysis for satisfying the C.G and
Moment of Inertia (M.I) constraints. For a spin stabilized satellite the M.I
about the spin axis should be greater than the other two axes by at least 11%.
The Finite Element Analysis of the micro-satellite was generated using MSC
PATRAN with all the subsystems modeled as point mass elements and
connected to the structure through 1D-beam elements. Figure 1.4 shows the
process involved in developing a micro-satellite structure. The following
analyses have been made using MSC NASTRAN.
Modal identification, to verify the lateral mode above 45 Hz
and the longitudinal mode above 90 Hz in order to avoid
coupling between the launcher excitation modes and the
natural vibration modes of the satellite.
Optimization of the satellite structure based on the mass and
stiffness constraints.
Structural stress and buckling analysis is to ensure that the
structure will fulfill its intended function in a given loads
environment.
Sinusoidal analysis is to investigate the trend of response
characteristics with the qualification test data.
Random analysis is to determine the accelerations of the
subsystems as much as displacements and stresses on the
structure subjected to random excitations.
Thermal analysis is to predict the temperature in all the
components for all the seasons.
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Tests were conducted with the prototype of the satellite to
determine its natural frequencies using the Electro Dynamic Shaker. The
sinusoidal vibration tests and random vibration tests on the micro-satellite
were carried out in the lateral direction and longitudinal direction for the
given qualification data using Electro Dynamic Shaker. The purpose of the
tests is to determine deficiencies in design and in the process of manufacture.
Figure 1.4 Process of developing spacecraft structure
1.10
ORGANISATION OF THE THESIS
Chapter 1 focuses on the motivation towards the development of
micro-satellite structure, types of satellite, types of micro-satellite structure,
launcher specifications, space mission environmental loads, general
requirements and constraints in developing the satellite structure and the
process of developing spacecraft structure.
Chapter 2 discusses the literatures on design, analysis and
development of various satellites that has been done earlier.
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Chapter 3 deals with the types of satellite design, design of internal
configuration, design of external configuration and solar panel mounting of
the satellite structure.
Chapter 4 includes the mass location analysis of the micro-satellite
and the sensitivity analysis of the mass location values.
Chapter 5 presents the optimization of the micro-satellite structure
based on the mass and stiffness constraints and the free vibration analysis of
micro-satellite.
Chapter 6 discusses the structural stress analysis of micro-satellite
and the determination of factor of safety values of the various structural
components in the satellite for different load cases. It also deals with the
buckling analysis of the micro-satellite structure and identifies whether the
structure will undergo buckling instability or not for different critical load
cases.
Chapter 7 presents the dynamic responses of the micro-satellite for
the sine and random excitations.
Chapter 8 presents the experimental results of vibration of microsatellite namely the sinusoidal vibration test and the random vibration test of
the micro-satellite, comparison of theoretical and experimental results, test
observation and spacecraft performance.
Chapter 9 focuses on the temperature prediction of all the
components in the micro-satellite and the method to control the temperatures
of the components.
Chapter 10 presents the summary of results of all the chapters.
Chapter 11 gives conclusion of the entire research work and the
scope for the future work.