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STUD SAT: A REPORT Student Sate!ite
Chandrayaan 1: PSLV - C11 Anshuman Mansingh April 2011
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NITTE MAHALINGA ADYANTHAYA MEMORIAL INSTITUTE OF TECHNOLOGY (A UNIT OF NITTE EDUCATION TRUST)
NITTE - 574110 UDUPI DISTRICT, KARNATAKA AUTONOMOUS COLLEGE UNDER VISVESVARAYA TECHNOLOGICAL UNIVERSITY RECOGNISED BY - ALL INDIA COUNCIL FOR TECHNICAL EDUCATION, NEW DELHI ACCREDITED BY NATIONAL BOARD OF ACCREDITATION ISO - 9001-2000 CERTIFIED
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TABLE OF CONTENTS Introduction!
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In the era of development in space technology, miniaturization plays a vital role to perform cost effective space missions. Pico-sate$ites require short development time and low cost. These compe$ing factors make them feasible candidates for a variety of ambitious space missions, especia$y the experimental missions. These space missions provide a platform to co$aborate many academic institutions and research organizations; bridging the gap between academia, research and industry.! 8
OBJECTIVES AND GOALS!
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Mission Objectives!
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Mission Goals!
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SMALL/EXPERIMENTAL SATELLITES!
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THE STUDSAT 1 MISSION!
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STUDSAT 1: INTRODUCTION! Description!
STRUCTURE!
19 20
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Objectives!
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Material!
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Specifications!
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Stages of development!
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Structure Dimensions!
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Catia Images!
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COMMUNICATION SYSTEM!
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Objective!
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A look into STUDSAT 1!
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Specification!
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Beacon specification!
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Telemetry format!
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Antenna system!
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ADCS!
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Attitude Determination & Control System!
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Objective!
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A Look Into STUDSAT 1!
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Specification!
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Coils specification!
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ON BOARD COMMAND AND DATA HANDLING!
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Antenna Deployment!
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Stabilize the sate$ite in the orbit!
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Establish a communication link with ground station!
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Monitor the health of the sate$ite!
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To capture images of the earth!
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ELECTRICAL POWER SYSTEM! Objectives!
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EJECTION SYSTEM!
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NASTRAC!
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NASTRAC-Nitte Amateur Sate$ite Tracking Center!
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Introduction!
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Data Processing!
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Telemetry Display!
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BEACON!
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OVERALL ARCHITECTURE!
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STUDSAT 2A/2B!
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TECHNICAL COMPONENTS!
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Subsystems!
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Proposed Objectives!
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Scientific Experiments!
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Test Bed for Cutting Edge Technologies developed for Sma$ Sate$ite Missions!
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REFERENCE(S)!
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∞!
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Anshuman Mansingh!
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I N TRODUCTION STUDSAT Programme Anshuman Mansingh April 2011 STUDSAT 2A/2B
Images courtesy Google, Some images related to the Indian space programme
In the era of development in space technology, miniaturization plays a vital role to perform cost effective space missions. Pico-satellites require short development time and low cost. These compelling factors make them feasible candidates for a variety of ambitious space missions, especially the experimental missions. These space missions provide a platform to collaborate many academic institutions and research organizations; bridging the gap between academia, research and industry. In this regard and in an Indian context, Project STUDSAT 2A/2B makes an attempt to test some of the very essential technologies such as constellation flying of satellites and inter satellite communication, along with various scientific experiments carried on board of the two satellites STUDSAT 2A & STUDSAT 2B. Also STUDSAT 2A/2B being a student oriented satellite educational program will provide an excellent opportunity for students and the faculty involved in the project. In developing satellites such as STUDSAT 2A & STUDSAT 2B, the approach is to design the entire architecture in order to serve the objectives by selecting a suitable payload. Depending upon the requirement of the payload specifications of other subsystems like Structure, On Board Communication System, Electronic Power System, On Board Command & Data Handling, Attitude Determination & Control System will be finalized. During the phase of design & development the hardware & software of subsystems will be developed according to the respective specifications. The project STUDSAT 2A/2B will adopt Cubesat standards to develop it satellites STUDSAT 2A and STUDSAT 2B. Both the satellites will be similar in their dimensions and mass, but with different payloads and power handling capabilities.
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OB JECTI VES AND GOALS Mission Objectives and Goals Project Studsat 2A/2B
Mission Objectives 1.
Mission Objectives of Project STUDSAT 2A/2B: Objectives:
To provide students a hands on experience on satellite design and development. • •# To provide a platform for research and academic community to qualify their work products / technological innovations for space applications. •# To demonstrate co development and integration of systems from various disciplines of science, technology and engineering. To demonstrate inter institution and student collaboration. • • To venture into high end technologies such as Inter-satellite communication, Formation Flying etc. •# To develop cost effective technologies such as GPS, EPS etc for use by academic institutions and research laboratories of the region. To enable undergraduate students the opportunity to track and communicate with satel• lites, space stations.
Mission Goals 2.
Mission Goals of STUDSAT Project:
To obtain a communication link with the satellite. • To take the image of a space object. • To take the image of the Earth. • •# To demonstrate the techniques of advanced compression algorithms onboard. •# To demonstrate accurate calculations done by students in calculating the orbit and path consideration which are involved while taking the picture of the other space object and Earth.
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S M A LL/EX P ER I MENTAL SATELLITE S A report on sma!/experimental sate!ites launched by ISRO Indian Space Research Organization April 2011
A RYA B H ATA Launch Date 19.04.1975
The First Indigenously built Indian Satellites
Mission Scientific/ Experimental Weight 360 kg On board power 46 Watts Communication VHF band Stabilization Spinstabilize Payload X-ray Astronomy Aeronomy & Solar Physics Launch date April 19,1975 Launch site Volgograd Launch Station (presently in Russia) Launch vehicle C-1 Intercosmos Orbit 563 x 619 km Inclination 50.7 deg Mission life 6 months(nominal), Spacecraft mainframe active till March,1981 Orbital Life Nearly seventeen years (Re-entered on February 10,1992)
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RT P Launch Date 10.08.1979
Mission Experimental Weight 35 kg onboard power 3 Watts Communication VHF band Stabilization Spin stabilized (spin axis controlled) Payload Launch vehicle monitoring instruments Launch date August 10,1979 Launch site SHAR Centre, Sriharikota, India Launch vehicle SLV-3 Orbit Not achieved
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APPLE Launch Date 19.06.1981 APPLE was used for nearly two years to carry out extensive experiments on time, frequency and code division multiple access systems, radio networking computer inter connect, random access and pockets witching experiments. Mission Experimental geostationary communication Weight 670 kg Onboard Power 210 Watts Communication VHF and C-band Stabilization Three axis stabilized (biased momentum) with Momentum Wheels, Torquers & Hydrazine based Reaction control system Payload C - band transponders (Two) Launch Date June19,1981 Launch Site Kourou (CSG), French Guyana Launch Vehicle Ariane -1(V-3) Orbit Geosynchronous (102 deg. E longitude, over Indonesia) Inclination Near zero Mission life Two years
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RS-1
Launch Date
18.07.1980
First satellite successfully launched by the indigenous launch vehicle SLV Mission Experimental Weight 35 kg Onboard power 16 Watts Communication VHF band Stabilization Spin stabilized Payload Launch vehicle monitoring instruments Launch date July 18,1980 Launch site SHAR Centre, Sriharikota, India Launch vehicle SLV-3 Orbit 305 x 919 km Inclination 44.7 deg. Mission life 1.2 years Orbital life 20 months
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SRE-1
Launch Date 10.01.2007 Space Capsule Recovery Experiment (SRE – 1) is a 550 kg capsule intended to demonstrate the technology of an orbiting platform for performing experiments in micro gravity conditions. After completion of the experiments, the capsule was de-orbited and recovered. SRE – 1 mission provided a valuable experience in fields like navigation, guidance and control during the re-entry phase, hypersonic aero thermodynamic, development of reusable thermal protection system (TPS), recovery through deceleration and flotation, besides acquisition of basic technology for reusable launch vehicles. SRE – 1 carries two experiments, an Isothermal Heating Furnace (IHF) and a Bio-mimeic experiment. SRE was launched into a 635 km polar SSO in January 2007 as a co-passenger with CARTOSAT -2 and stayed in orbit for 10 days during which its payloads performed the operations they are intended to. The SRE capsule was de-boosted and recovered successfully back on earth on 22nd January 2007.
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A N U S AT Launch Date 20.04.2009 ANUSAT (Anna University Satellite) is the first satellite built by an Indian University under the over all guidance of ISRO and will demonstrate the technologies related to message store and forward operations. Altitude 550 km Inclination 41 deg Orbit Period 90 minutes Mass 40 kg
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S T U D S AT Launch Date 12.07.2010 Student Satellite (STUDSAT) is the first pico-satellite developed in the country by a consortium of seven engineering colleges from Karnataka and Andhra Pradesh. STUDSAT weighing less than 1 kg, has the primary objective of promoting space technology in educational institutions and encourage research and development in miniaturized satellites, establishing a communication link between the satellite and ground station, capturing the image of earth with a resolution of 90 meters and transmitting the payload and telemetry data to the earth station.
Mission Experimental / Small Satellite Weight Less than 1 kg Altitude 630 km Orbit Polar Sun Synchronous
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YO U T H S AT Launch Date 20.04.2011 YOUTHSAT is a joint Indo-Russian stellar and atmospheric satellite mission with the participation of students from Universities at graduate, post graduate and research scholar level. With a lift-off mass of 92 kg, Youthsat is a mini satellite and the second in the Indian Mini Satellite (IMS) series. Youthsat mission intends to investigate the relationship between solar variability and thermosphere-Ionosphere changes. The satellite carries three payloads, of which two are Indian and one Russian. Together, they form a unique and comprehensive package of experiments for the investigation of the composition, energetics and dynamics of earth's upper atmosphere. The Indian payloads are: 1. RaBIT (Radio Beacon for Ionospheric Tomography)- For mapping Total Electron Content (TEC) of the Ionosphere. 2.
LiVHySI (Limb Viewing Hyper Spectral Imager) - To perform airglow measurements of the Earth's upper atmosphere (80- 600 km) in 450-950 nm.
The Russian payload: 1.
SOLRAD - For monitoring the solar X- and gamma ray fluxes and to study solar cosmic ray flux parameters and conditions of their penetration in the Earth's magnetosphere.
Lift-off Mass 92 kg Orbit Period 101.35 minDimension 1020 (Pitch) x 604 (Roll) x 1340 (Yaw) mm3 Attitude and Orbit Control 3-axis body stabilised using Sun and Star Sensors, Miniature Magnetometer, Miniature Gyros, Micro Reaction Wheels and Magnetic Torquers Power Solar Array generating 230 W, one 10.5 AH Li-ion battery Mechanisms Paraffin Actuator based Solar Panel Hold Down and Release Mechanism Launch date April 20, 2011 Launch site SHAR Centre Sriharikota India Launch vehicle PSLV- C16 Orbit Circular Polar Sun Synchronous Orbit altitude at injection 822 km + 20 km (3 Sigma) Orbit Inclination 98.731 º + 0.2 º Mission life 2 years
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TH E STUDSAT 1 MISSION The STUDSAT (Student Sate!ite) programme
Anshuman Mansingh April 2011
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STUDSAT 1: INTRODUCTION About the mission, Team Studsat Anshuman Mansingh April 2011 Studsat (Student Satellite) is a unique satellite technology endeavor undertaken by Undergraduate students in India. Studsat will be the first Pico-Satellite being launched by India, and more so special, as it is a project undertaken by Under-graduate students from seven different Academic Institutions, from different regions of India, under the guidance of the Indian Space Research Organization (ISRO). When realized successfully Studsat will be the smallest operational satellite launched by ISRO. The mission of the satellite is experimental in nature, but the major objective of this mission is to enable students to have hands-on experience on Space technology by involving themselves in the design, fabrication and implementation of a full-fledged space mission at a nominal budget. The idea for a project of this kind crystallized during the International Astronautical Congress, 2007, between four students from different academic institutions, from Hyderabad and Bengaluru, after their epochal interaction with the Project Director of Small Satellites, ISRO Satellite Centre. Starting then, the team has expanded slowly by involving more like minded students, to complete the entire conceptual design of the satellite. By the virtue of its collaborative nature, the students approached each of their academic institutions, persuading them to contribute funds for the realization of this ambitious project. An accurate and convincing budget analysis was performed by the students, presenting the same to the managements seeking their much needed support, in terms of resources, sponsorship and finance pooling. The institutions were overwhelmed by the Team's enthusiasm, and wholeheartedly supported the team. The team then approached ISRO along with the Academic institutions for a Preliminary Review of the Student Satellite project. After detailed review of the project, and elaborate presentations from the team, ISRO approved our project. Subsequently, the team grew, and today comprises of around 45 students from ten different academic institutions. Of the ten participant colleges, seven formed a consortium in order to sponsor the project. An internal MoU has been signed between the colleges, and have chosen Nitte Meenakshi Institute of Technology,
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Bengauru, one of the participant institutions as their representative, to sign the official MoU with ISRO on their behalf.
DESCRIPTION
The satellite is close to being a cube, of miniature size, when compared with the common satellites, with dimensions of 10 cm x 10 cm x 13.5 cm; it weighs just above 850 gm and has a volume of 1.1 litre, and hence falls into the category of "Pico Satellites". The satellite is intended to be launched in a 700 km solar-synchronous orbit. The functional objective of the satellite is to perform remote sensing, and capture images of the surface of the earth using it's camera of resolution 90 m; The best resolution hitherto achieved by any Pico Satellite in the world. * Communication sub-system. * Power generation and distribution sub-system. * Attitude Determination and Control sub-system. * On-board computer. * Payload(Camera). * Mechanical Structure. Apart from these constituent subsystems of the satellite, a fully operational amateur frequency Ground Station has been designed in order to communicate with the satellite. The Ground Station has been established in Nitte Meenakshi Institute of Technology, Bengaluru. All the above subsystems have been indigenously designed by the students.
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STRUCTURE Subsystem A Report STUDSAT 1 OBJECTIVES * To withstand the forces acting during launch and ejection * To protect the satellite electronics from thermal issues * To accommodate and protect the sub-systems during the entire mission. * To design a structure this is compatible with all the ejection systems. * To maintain CUBESAT standards.
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M AT E R I A L High grade Aluminum Alloy which is space qualified. S P E C I F I C AT I O N S * The CubeSat shall be 100.0+0.1 mm wide (X and Y) and 113.5+0.1 mm tall (Z). * Rails are smooth and edges must be rounded to a minimum radius of 1 mm. * No external components other then the rails shall be in contact with the internals of the Ejection system. * Each rail shall be a minimum of 8.5 mm wide. * At least 75% of the rail must be in contact with the Ejection system rails. 25% of the rails may be recessed and NO part of the rails may exceed the specification. * All rails must be anodized to prevent cold-welding, reduce wear, and provide electrical isolation between the CubeSats and the Ejection system. * Separation springs must be included at designated contact points. STAGES OF DEVELOPMENT * Initial design of the structure incorporating all the subsystems of the satellite. * Meshing of the design. * Analysis of the design. * Prototype fabrication. * Qualification and Acceptance tests. * Flight model fabrication. STRUCTURE DIMENSIONS
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C AT I A I M AG E S
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COMMUNICATION SYSTEM Subsystem A Report STUDSAT 1 OBJECTIVE The objective of this system is to establish communication link with ground station.The telemetry and payload data is transmitted to the ground station .A customized protocol is used for representing the data. The communication system should be developed within the constraints imposed with respect to power, size and mass.
A
LOOK INTO
S T U D S AT 1
S P E C I F I C AT I O N
BEACON
S P E C I F I C AT I O N
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T E L E M E T RY
F O R M AT
Modulation: Morse code will be used to modulate the signal. Table 1 in the appendix. Shows the morse code representation of various alphabets, numbers and punctuation marks by a combination of dots and dashes It should be noted that: * A DAH (-) is 3 times longer than a DIT (.). * The space between a DIT and a DAH is one DIT. * The space between two alphabets is a DAH. * The space between two words is seven DITs. The beacon message is divided into 4 parts. Each part is transmitted at an interval of 30sec. hence, it takes 2 minutes to transmit the complete telemetry. The contents of the telemetry are described below. The numbers transmitted follow the octal representation. Part 0: Part 0 contains the satellite ID. The format of part 0 is as shown below: Header-Space-Satellite ID-Space The Header contains the Part number. In this case Header contains '0'. The header is followed by a space then the satellite ID. The satellite ID contains 6 characters in alpha-numeric form and the corresponding morse code is transmitted.
Part 1: Part 1 contains some of the status parameters of the satellite and battery 1 voltage. The format of Part 1 is as shown below: Header-Space-Status Parameters-Space-Battery 1 Voltage Part 1 starts with a header which contains the part number, In this case it is '1'. The header is followed by a space then the Status parameters of the satellite. This field provides information about the status(ON/OFF) various subsystems of the satellite. The different status parameters are as follows: * Payload ON/OFF * EPS ON/OFF * ADCS ON/OFF
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* Communication ON/OFF * Antenna Deploy. ON/OFF * Image ready to Tx. Yes/No There are six parameters which can be represented by one bit each. Say 0 representing OFF or NO and 1 representing ON or Yes. Consider an example where the "Status Parameter" field contains "111010", this in octal form is 72 and it means the following: - Payload : ON - EPS: ON - ADCS: ON - Communication :OFF - Antenna Deploy.: ON - Image ready to Tx.: NO. The status parameter field is followed by a space, followed by Battery1 Voltage. The battery voltage is represented using an ADC value represented by 10 bit number. To decode the actual value of the battery voltage, the following formula is used: V = (-0.00939 * ADC) + 9.791 V.
Part 2 Part 2 contains the coil currents and battery 2 voltage. The format of part 2 is as follows: Header-Space-Coil Current-Space-Battery 2 voltage Part 2 starts with a header which contains the part number, In this case it is '2'. The header is followed by a space, followed by the coil currents of the 3 coils placed on 3 different axis. The coil current varies from -20mA to +20mA with a step size of 1mA. This can be represented using 6 bits as follows. -20mA is represented as '0' (in binary 000000) and +20mA is represented as '40'(in binary 100000). Thus a total of 18bits to represent the current flowing in 3 coils. The coil Current field is followed by a space, followed by Battery 2 Voltage. The battery voltage is represented using an ADC value represented by 10 bit number. To decode the actual value of the battery voltage, the following formula is used: V = (-0.00939 * ADC) + 9.791 V.
Part 3:
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Part 3 contains panel current and RTC value. The format of Part 2 is as shown below: Header-Space-Panel current-space-RTC The different panels from which the current is measured and transmitted are X1, X2, Y1, Y2 and Z1. Each parameter has an ADC value represented using 10bits, associated with each of the panel currents. It is decoded using the following formula. I = ((-0.486*ADC) + 502.524) mA.
ANTENNA
SYSTEM
* Two Perpendicular monopoles configuration for data Transmit/Receive * Two perpendicular monopoles configuration for beacon Transmit This kind of configuration maintains stability of spacecraft after deployment and provides a good near omnidirectional pattern for successful communication link.
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ADCS Subsystem A Report STUDSAT 1
A T T I T U D E D E T E R M I N AT I O N & C O N T RO L S Y S T E M OBJECTIVE The attitude of a spacecraft is its orientation in space. The determination of attitude and controlling it to the requirement within the tolerance is a part of attitude determination and control. It projects the feasibility study and describes the preliminary analysis of both the position of the satellite centre of mass in three dimensional space and the methods to determine and control the orientation of the satellite with respect to its centre of mass in three dimensional spaces thus stabilising the satellite in the orbit. Thus works accordingly to point the camera of the satellite towards Nadir within tolerance.
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A L O O K I N T O S T U D S AT 1 S P E C I F I C AT I O N
COILS
S P E C I F I C AT I O N
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ON BOARD COMMAND AND DATA HANDLING Subsystem A Report STUDSAT 1 The command and data handler forms the brain of the STUDSAT. It is responsible for controlling the satellite from the time it is ejected into the space. The command and data handling unit STUDSAT consists of a central controller to control all the subsystems of the satellite. The controller that will be used for this purpose is a 32 bit RISC controller with features like; very low power consumption and DSP instruction set. Some of the major objectives of the OBC&DH are as listed below: ANTENNA DEPLOYMENT As the satellite is ejected into the orbit, the first major task is to deploy the antennas. This task is initiated the OBC&DH subsystem. The Onboard computer commands the antenna system to deploy, so that that satellite can successfully establish a link with the ground station.
STABILIZE
T H E SAT E L L I T E I N T H E O R B I T Once the satellite is ejected into the orbit, the satellite will be highly unstable due to the various forces acting on it. Hence the satellite's first task is to get stabilized and orient the antennas towards the earth. This process is called the attitude determination and correction (ADC) and it is carried out by the OBC.
ESTABLISH
A C O M M U N I C A T I O N L I N K W I T H G R O U N D S T ATION when the satellite is stable, it should be in a position to communicate with ground station. Hence after attaining stability, the satellite will wait till it reaches the communication window and start establishing a communication link with the ground station. The communication protocol used for communicating with the ground station is the modified version of the AX.25 protocol and frames used are the AX.25 UI frames. The OBC&DH system is responsible for the implementation of the complete communication protocol.
MONITOR
T H E H E A LT H O F T H E SAT E L L I T E It is very important to monitor the health of the satellite and coordinate the activities in the satellite so that no subsystems eats away all the power and the satellite is left with no power to perform any further tasks. So the OBC performs the task of monitoring the health of the satellite by acquiring the data from various sensors and comparing their values with the previously set thresholds. Based on the result of the comparison appropriate actions are taken.
TO
CAPTURE IMAGES OF THE EARTH
The main objective of the satellite is to capture the images of the earth. This function is performed by the camera placed in the satellite. The actions of the camera are completely controlled by the OBC. The captured image is stored onboard in a parallel FRAM and is transmitted to the ground station as the link is established.
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ELECTRICAL POWER SYSTEM Subsystem A Report STUDSAT 1
OBJECTIVES
The Solar Panel is suitably Designed and Implemented.A detailed study of the COTS Power Board is done with required testing and verification.The power budget analysis needs to be done meticulously and the power distribution architecture is to be identified.
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EJECTION SYSTEM Subsystem A Report STUDSAT 1
Ejection system is used to place a satellite into the relevant orbit.It acts as the interface between adapter of PSLV(rocket)and satellite. The ejection system will be mounted on the adapter of the rocket and the satellite will be placed inside the Ejection system,Once the rocket reaches the required orbit some commands will be issued to the Ejection system and it in turn ejects the sateelite into the orbitThe main objective of the system is to successfully eject the cubesatellite into the orbit from the launch vehicle.
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NASTRAC Subsystem, Nitte Amateur Sate!ite Tracking Center A Report STUDSAT 1
N A S T R AC - N I T T E A M AT E U R S AT E L L I T E T R AC K I N G CENTER
INTRODUCTION
The NASTRAC is the first and final terrestrial end of a communications link to an object in outer space. Wireless Communication is done with Satellites; hence the ground station serves as the access point on Earth. The satellite dumps the data to the ground station whenever it passes by that area. Main purpose of the ground stations is to track and receive telemetry and image data from satellite for data analysis, and also to control the satellite by commanding. Ground station consists of hardware and software to transmit and receive data reliably. The components include, a computer programmed with orbital-prediction software compatible with the hardware for auto-tracking, a transceiver to transmit and receive data and a TNC to decode the incoming data. The SGS(Studsat Ground Station) is a part of the student Pico-satellite program. SGS is a proposed satellite tracking facility to be built by undergraduate students. The STUDSAT team will develop a working prototype of Pico-satellite that conforms to the CUBESAT standard of total mass less than 1Kg, with dimensions of 10 cm x 10 cm x 11.3 cm. The main objective of the STUDSAT comprises of taking images of Earth's surface. The architecture of the SGS is based on its operation and requirement. The following diagram shows the architecture of the ground system. The gain of the antenna, gain and NF of the LNA, the affordable losses in cable and the EIRP were all calculated based on the link budget analysis. The received data undergoes various processes such as amplification, down conversion and demodulation before bifurcating into image data and telemetry tracking data. The image data is further processed based on the end user application and the telemetry data is used to
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study the health of the satellite. The up-command to be transmitted to the satellite is modulated, amplified and finally directed towards antenna systems for transmission. Downlink: The down link operating frequency is 437.5MHz. The link is initiated by the ground station. After the satellite receives the Link Establishment Frame (LEF), the satellite starts dumping telemetry data. The communication protocol used is AX.25 (using UI frame). The satellite then waits for a command from the ground station and on receiving the command it will start transmitting the image. Uplink: The up-link operating frequency is 437.5 MHz. It is a simplex mode of communication between the satellite and ground station. The satellite uplinks the LEF, Keplerian elements( for on-board orbit prediction algorithm) and up-commands.
D ATA P RO C E S S I N G
Primary Processing * Image Sharpening. * Image Smoothing. * Negative of an image.
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* Image zooming. * Histogram generation. * Histogram equalization. * Image scaling. Secondary image Processing * Map Synchronization * User defined imaging
T E L E M E T RY D I S P L AY
The different telemetry data received are: * Battery charge and discharge current * Mean anomaly * Right ascension of ascending node (RAAN) * Battery voltage level * Solar cell voltage * Temperature. * Modes of operation
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BEACON Subsystem A Report STUDSAT 1 Beacon Specification
Mode of communication Data rate MODE Amateur frequency
Simplex 18wpm downlink. CW(continuous wave) 437.861MHz.
T elemetry format Modulation: Morse code will be used to modulate the signal. Table 1 in the appendix. Shows the morse code representation of various alphabets, numbers and punctuation marks by a combination of dots and dashes It should be noted that: - A DAH (-) is 3 times longer than a DIT (.). - The space between a DIT and a DAH is one DIT. - The space between two alphabets is a DAH. - The space between two words is seven DITs.
The beacon message is divided into 4 parts. Each part is transmitted at an interval of 30sec. hence, it takes 2 minutes to transmit the complete telemetry. The contents of the telemetry are described below. The numbers transmitted follow the octal representation. Part 0: Part 0 contains the satellite ID. The format of part 0 is as shown below: Header-Space-Satellite ID-Space 7KH+HDGHUFRQWDLQVWKH3DUWQXPEHU,QWKLVFDVH+HDGHUFRQWDLQVÄ The header is followed by a space then the satellite ID. The satellite ID contains 6 characters in alpha-numeric form and the corresponding morse code is transmitted.
Part 1: Part 1 contains some of the status parameters of the satellite and battery 1 voltage. The format of Part 1 is as shown below:
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Header-Space-Status Parameters-Space-Battery 1 Voltage Part 1 starts with a header which contaLQVWKHSDUWQXPEHU,QWKLVFDVHLWLVÄ The header is followed by a space then the Status parameters of the satellite. This field provides information about the status(ON/OFF) various subsystems of the satellite. The different status parameters are as follows: - Payload ON/OFF - EPS ON/OFF - ADCS ON/OFF - Communication ON/OFF - Antenna Deploy. ON/OFF - Image ready to Tx. Yes/No. There are six parameters which can be represented by one bit each. Say 0 representing OFF or NO DQG UHSUHVHQWLQJ21 RU
V = (-0.00939 * ADC) + 9.791 V. Part 3: Part 3 contains panel current and RTC value. The format of Part 2 is as shown below: Header-Space-Panel current-space-RTC The different panels from which the current is measured and transmitted are X1, X2, Y1, Y2 and Z1. Each parameter has an ADC value represented using 10bits, associated with each of the panel currents. It is decoded using the following formula. I = ((-0.486*ADC) + 502.524) mA.
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OVERALL ARCHITECTURE STUDSAT 1 Student Satellite A Figure
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S TUDSAT 2A/2B Student Sate!ite: Proposed Mission
Anshuman Mansingh April 2011
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TECHNICAL COMPONENTS (Proposed) Student Satellite
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Subsystems STUDSAT 2A •) •) •) •) •) •) •) •
Payload 2 ( PMT) Attitude Determination & Control System. Electronic Power System. Additional Battery Pack. Communication System. GPS Module. Command & Data Handling. Structure.
21) STUDSAT 2B • • • • • • •
Payload 2A ( Nano IR Image Sensor) Attitude Determination & Control System. Electronic Power System. Communication System. GPS Module. Command & Data Handling. Structure.
Proposed Objectives SCIENTIFIC EXPERIMENTS Photomultiplier Tube (PMT) Application: Space Radiation Measurement, To study thunderstorm-induced phenomena in earth's atmosphere, Atmospheric Studies. To study the airglow layer of Earth’s upper atmosphere. Imaging and tracking of wave-like structures ranging from 20 to 200 km. Data will be very useful for the astronomers of world wide. Nano Infra Red Image Sensor Application: Remote Sensing, Vegetation Studies First IR camera developed with Nanotechnology to be flown to space. Study the behavior of nano materials in space. Quantum Dot Infrared Photo (QDIP) structure with pixel size 50 μm x 50 μm.
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TEST BED
FOR
VELOPED FOR
CUTTING EDGE TECHNOLOGIES S M A L L S AT E L L I T E M I S S I O N S
D E-
• MEMS based thrusters. • Micro Reaction Wheels. • Other possible subsystems within the constraints.
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*A real-time operating system (RTOS) is an operating system (OS) intended to serve realtime application requests. A key characteristic of an RTOS is the level of its consistency concerning the amount of time it takes to accept and complete an application's task; the variability is jitter. A hard real-time operating system has less jitter than a so* real-time operating system. The chief design goal is not high throughput, but rather a guarantee of a soft or hard performance category. An RTOS that can usually or genera!y meet a deadline is a soft real-time OS, but if it can meet a deadline deterministically it is a hard real-time OS. A real-time OS has an advanced algorithm for scheduling. Scheduler flexibility enables a wider, computer-system orchestration of process priorities, but a real-time OS is more frequently dedicated to a narrow set of applications. Key factors in a real-time OS are minimal interrupt latency and minimal thread switching latency, but a real-time OS is valued more for how quickly or how predictably it can respond than for the amount of work it can perform in a given period of time.
*A magnetometer is a scientific instrument used to measure the strength or direction of the magnetic field, produced either in the laboratory or existing in nature. The Earth's magnetic field (the magnetosphere) varies from place to place, for various reasons such as inhomogeneity of rocks and the interaction between charged particles from the Sun and the magnetosphere. Magnetometers are a frequent component instrument on spacecraft that explore planets.
*In satellite systems, a magnetorquer or magnetic torquer is a system for attitude control, detumbling and stabilization built from electromagnetic coils. The magnetorquer develops a magnetic field which interfaces with an ambient magnetic field, usually the Earth's, so that the counter-forces produced provide useful torque.
*AX.25 is a data link layer protocol derived from the X.25 protocol suite and designed for use by amateur radio operators.[1] It is used extensively on amateur packet radio networks. AX.25 occupies the first, second, and often the third layers of the OSI networking model, and is responsible for transferring data (encapsulated in packets) between nodes and detecting errors introduced by the communications channel. AX.25 is thus comparable to the combination of Ethernet and TCP in the services it provides. However, as AX.25 is a pre-OSI-model protocol, the specification was not written to cleanly separate into OSI layers. This is not necessarily a problem, but can have some advantages. In practice, it is not uncommon to find an AX.25 data link layer as the transport for some other network layer, such as IPv4, with TCP used on top of that. Note that, like Ethernet, AX.25 frames are not engineered to support switching. For this reason, AX.25 supports a
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somewhat limited form of source routing. Although possible to build AX.25 switches in a manner not unlike how Ethernet switches work, this has not yet been accomplished. AX.25 supports both virtual-circuit connected and datagram-style connectionless modes of operation. The latter is used to great effect by the Automatic Packet Reporting System.
*Global Positioning System receivers provide highly accurate time and position information, and use a constellation of 24 navigation satellites maintained by the US DOD in order to accurately determine their own position. Although designed for Terrestrial geolocation and navigation, the system can be employed in the highly dynamic environment in Low Earth Orbit. Landsat was the first civil spacecraft to carry a GPS receiver into orbit, but the receiver technology has advanced to a stage where GPS receivers can now routinely be carried on smaller platforms. Since Selective Availability has been switched off, accuracies in the order of 10m in all axis are achievable for simple position fixes. A space borne GPS receiver is different from a terrestrial receiver, as it travels at great velocity relative to the GPS constellation, and satellites tracked change constantly. As a consequence it has to cope with higher Doppler shifts and Doppler shift rates. In LEO this can be of the order of ±60kHz. Consequently, the receiver only has a short time available to perform a frequency/code search in order to lock onto the spread spectrum GPS signals. In order to predict GPS satellite visibility, the antenna geometry and orbital dynamics of the host satellite must be well known. For more precise applications, differential and double-differential filtering of GPS is proposed for formations of satellites. For greater precision, kinematic GPS employs the basic understanding of the dynamics to further improve accuracy. Orbit determination using GPS receivers on small satellites is now well demonstrated, and the next challenge now is to perform attitude determination on a small platform. Several small missions have recently been launched in order to demonstrate this. GPS can also be used in ionospheric research, and this is achieved by monitoring GPS signals as the satellite-to-satellite path skims the upper atmosphere. GPS can only be effectively used near the Earth's surface, but various experiments have demonstrated that sidelobe signals can be picked up above the GPS constellation at 20,000; e.g. from Geostationary heights.
*Microelectromechanical systems (MEMS) (also written as micro-electro-mechanical, MicroElectroMechanical[citation needed] or microelectronic and microelectromechanical systems) is the technology of very small mechanical devices driven by electricity; it merges at the nano-scale into nanoelectromechanical systems (NEMS) and nanotechnology. MEMS are also referred to as micromachines (in Japan), or micro systems technology - MST (in Europe).
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**Propulsion systems with the capability to deliver accurate micro- to milli-Newton thrust levels has been identified as mission critical for many advanced space systems such as Darwin, Gaia, LISA and Microscope, to mention a few, currently under development. Different technologies are being pursued, and one of the promising concepts is based on Micro Electro Mechanical System (MEMS) technology. (For instance: The Northrop Grumman-led Digital Micro-Propulsion Program is producing and demonstrating tiny thrusters to perform orbital insertion, station keeping and attitude control functions on micro-, nano- and pico-satellites. Individual micro-electromechanical system (MEMS) thrusters, each a poppy seed-sized cell, are placed in arrays made up as a three-layer silicon and glass sandwich, fabricated with the middle layer consisting of an array of small cells, or plenums, and sealed with a rupturable diaphragm on one side. Each cell represents a separate thruster, like a solid rocket motor. The cells are loaded with combustible propellants or an inert substance in gas, liquid or solid form. When ignited, each cell delivers one impulse bit. Propulsion is adjusted in discrete increments by igniting individual thrusters, several thrusters at once, or in controlled sequences to deliver precise amounts of impulse. The microthruster design offers several advantages over conventional thrusters: it has no moving parts, uses a variety of propellants, is scalable, and eliminates the need for tanks, fuel lines and valves.)
*A reaction wheel is a type of flywheel used primarily by spacecraft for attitude control without using fuel for rockets or other reaction devices. Reaction wheels are devices which aim a spacecraft in different directions without firing rockets or jets. They are particularly useful when the spacecraft must be rotated by very small amounts, such as keeping a telescope pointed at a star. They may also reduce the mass fraction needed for fuel. This is accomplished by equipping the spacecraft with an electric motor attached to a flywheel, which when rotated increasingly fast causes the spacecraft to spin the other way in a proportional amount by conservation of angular momentum. Reaction wheels can only rotate the spacecraft around its center of mass (see torque), they are not capable of moving the spacecraft from one place to another (see translational force). Reaction wheels work around a nominal zero rotation speed. However, external torques on the spacecraft may require a gradual buildup of reaction wheel rotation speed to maintain the spacecraft in a fixed orientation. Momentum wheels (used in the Hubble Space Telescope) are a different type of actuator, mainly used for gyroscopic stabilization of spacecraft: momentum wheels have high rotation speeds (around 6000 rpm) and mass. Reaction wheels are usually implemented as special electric motors, mounted along the x, y and z axes. Changes in speed rate (in either direction) are controlled electronically by computer controls. The strength of the materials of a reaction wheel determines the speed at which the wheel would come apart, and therefore how much angular momentum it can store.
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Since the reaction wheel is a small fraction of the spacecraft's total mass, easily-measurable changes in its speed provide very precise changes in angle. It therefore permits very precise changes in a spacecraft's attitude. For this reason, reaction wheels are often used to aim spacecraft with cameras or telescopes. Over time reaction wheels may build up stored momentum that needs to be cancelled. Designers therefore supplement reaction wheel systems with other attitude control mechanisms. In the presence of a magnetic field (as in low Earth orbit), a spacecraft can employ magnetorquers (better known as torque rods) to transfer angular momentum to the Earth through its magnetic field. In the absence of a magnetic field, the most efficient practice is to use highefficiency attitude jets such as ion thrusters, or small, lightweight solar sails on the ends of projecting masts or solar cell arrays. Most spacecraft, however, also need fast pointing, and cannot afford the extra mass of three attitude control systems. Designers therefore usually use conventional monopropellant vernier engines to cancel reaction wheels, as well as for fast pointing.
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R EF ERENCE(S) Bibliography and Webliography A Report April 2011 1. The World Wide Web teamstudsat.com/ wikipedia.org/ google.com/ ieeexplore.ieee.org/ onlinelibrary.wiley.com/ and others.
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∞ Anshuman Mansingh April 2011
NITTE MAHALINGA ADYANTHAYA MEMORIAL INSTITUTE OF TECHNOLOGY (A UNIT OF NITTE EDUCATION TRUST)
NITTE - 574110 UDUPI DISTRICT, KARNATAKA AUTONOMOUS COLLEGE UNDER VISVESVARAYA TECHNOLOGICAL UNIVERSITY RECOGNISED BY - ALL INDIA COUNCIL FOR TECHNICAL EDUCATION, NEW DELHI ACCREDITED BY NATIONAL BOARD OF ACCREDITATION ISO - 9001-2000 CERTIFIED
Anshuman Mansingh Student Bachelor of Engineering Department of Electronics and Communication Engineering
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White Space
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