CHAPTER 1 INTRODUCTION

Transcription

1 1 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

2 2 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 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.

3 3 Micro-Satellites (Spacecraft) fall into the 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

4 4 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.

5 5 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 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 kg 3 Mini satellite kg 4 Micro satellite kg 5 Nano satellite 1-10kg 6 Pico satellite 0.1-1kg 7 Femto satellite <100g

6 Based on the mission The satellites have classified based on their missions also as below, 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

7 7 1.5 SPACECRAFT STRUCTURES Spacecraft structures have mainly divided into three main categories and shown in Figure 1.1 of the model satellite 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 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 Tertiary structure The interconnecting elements, brackets and the electronic boxes form the tertiary structure of any satellite.

8 8 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

9 9 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 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 Sun- Synchronous 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 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 The first operational

10 10 flight carrying 1200 kg remote sensing satellite (IRS-1D) took place in September 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 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 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.

11 Figure 1.2 PSLV-C12 Configurations 11

12 12 Figure 1.3 PSLV axes definition 1.7 SPACE MISSION ENVIRONMENTAL LOADS 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 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 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 Dynamic internal loads Forces caused by internal motion for satellite attitude control like momentum wheel, magnetic torquer or any other type of electric servomotors.

13 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.

14 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.

15 15 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.

16 16 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.