VHF Beacon Development, Ground Segment, and Operations for Canx Missions

VHF Beacon Development, Ground Segment, and Operations for CanX Missions

Amee Shah

A thesis submitted in conformity with the requirements for the degree of Master’s of Applied Science Institute for Aerospace Studies University of Toronto © Copyright by Amee Shah 2009

Library and Archives Bibliothèque et Canada Archives Canada

Published Heritage Direction du Branch Patrimoine de l’édition

395 Wellington Street 395, rue Wellington Ottawa ON K1A 0N4 Ottawa ON K1A 0N4 Canada Canada

Your file Votre référence ISBN: 978-0-494-71380-8 Our file Notre référence ISBN: 978-0-494-71380-8

NOTICE: AVIS:

The author has granted a non- L’auteur a accordé une licence non exclusive exclusive license allowing Library and permettant à la Bibliothèque et Archives Archives Canada to reproduce, Canada de reproduire, publier, archiver, publish, archive, preserve, conserve, sauvegarder, conserver, transmettre au public communicate to the public by par télécommunication ou par l’Internet, prêter, telecommunication or on the Internet, distribuer et vendre des thèses partout dans le loan, distribute and sell theses monde, à des fins commerciales ou autres, sur worldwide, for commercial or non- support microforme, papier, électronique et/ou commercial purposes, in microform, autres formats. paper, electronic and/or any other formats. . The author retains copyright L’auteur conserve la propriété du droit d’auteur ownership and moral rights in this et des droits moraux qui protège cette thèse. Ni thesis. Neither the thesis nor la thèse ni des extraits substantiels de celle-ci substantial extracts from it may be ne doivent être imprimés ou autrement printed or otherwise reproduced reproduits sans son autorisation. without the author’s permission.

In compliance with the Canadian Conformément à la loi canadienne sur la Privacy Act some supporting forms protection de la vie privée, quelques may have been removed from this formulaires secondaires ont été enlevés de thesis. cette thèse.

While these forms may be included Bien que ces formulaires aient inclus dans in the document page count, their la pagination, il n’y aura aucun contenu removal does not represent any loss manquant. of content from the thesis.

VHF Beacon Development, Ground Segment, and Operations for CanX Missions

Amee Shah

Master’s of Applied Science

Institute for Aerospace Studies University of Toronto

2009 Abstract

The versatile Generic Nanosatellite Bus (GNB) technology fused with the Canadian Advanced

Nanospace eXperiment (CanX) missions is pioneered by the Space Flight Laboratory (SFL) at

University of Toronto Institute for Aerospace Studies. Microspace philosophy has been employed at SFL to develop low cost nanosatellites with emerging technologies for education and research.

This thesis provides an insight to the space systems engineering experience acquired by the author being a master’s student at SFL. This thesis describes the design, construction and testing of the VHF beacon transmitter for GNB. The umbilical Electrical Ground Support Equipment

(EGSE) designed and built by the author is presented in this thesis. The assembly, integration and testing of the SFL ground station for the CanX missions has been explored. The on-orbit spacecraft operations for CanX-2 and NTS, from the launch campaign to the commissioning phase to nominal operations along with the anomalies faced and the contingency operations carried out by the author to date have been explained in detail in this thesis.

Acknowledgments

First of all, I would like to express my gratitude to my supervisor, Dr. Robert Zee, for giving me such a golden opportunity to be a Master’s student at SFL and be part of the CanX adventure that has added considerably to my graduate experience. I highly appreciate his vision and passion to establish Space Flight Laboratory which has been a cradle for nanosatellites in Canada.

Being part of the SFL was a gratifying experience filled with motivation and encouragement from the students and the staff. I am grateful to the guidance, advice and training received from all the SFL staff members.

A very special thanks goes out to Mihail Barbu, Daniel Kekez and Milko Dimitrov without whose expertise, knowledge, skills and experience, this thesis wouldn’t have been possible. I would also like to acknowledge the spacecraft operations team – Karan Sarda, Daniel Kekez, Nathan Orr and Mirue Choi for assisting me and guiding me during contingencies and overall operations.

I would like to express my heartfelt gratitude to my parents for their unconditional love, encouragement and motivation. I would also like to thank my family, friends and teachers for the support they provided me through my entire life.

iii

Table of Contents

Acknowledgments...... iii

List of Figures ...... vi

List of Tables ...... x

List of Abbreviations & Acronyms...... xi

Chapter 1 Introduction...... 1

1.1 Introduction...... 1

1.2 Space Flight Laboratory...... 2

1.3 CanX-2...... 3

1.4 CanX-6/NTS ...... 4

1.5 CanX-3/BRITE ...... 5

1.6 CanX-4/-5 Formation Flying mission...... 6

1.7 AISSat-1...... 7

Chapter 2 Generic Nanosatellite Bus...... 8

2.1 Communication Subsystem ...... 8

2.2 VHF Beacon...... 9

2.3 Umbilical EGSE...... 30

Chapter 3 Ground Station...... 40

3.1 Introduction...... 40

3.2 UHF Ground Station...... 41

3.3 Limit Switch Assembly...... 47

3.4 Rotator controller...... 48

Chapter 4 Spacecraft Operations ...... 50

4.1 Introduction...... 50

4.2 CanX-2 On-Orbit Operations...... 60

4.3 NTS On-Orbit Operations...... 84

Chapter 5 Conclusion ...... 86

References...... 87

List of Figures

Figure 1: CanX-2 ...... 3

Figure 2: CanX-6/NTS...... 4

Figure 3: CanX-3/BRITE (Exploded View)...... 5

Figure 4: CanX-4/-5...... 6

Figure 5: AISSat-1 ...... 7

Figure 6: VHF Beacon block diagram ...... 13

Figure 7: VHF Beacon ...... 13

Figure 8: Voltage Regulator block diagram...... 15

Figure 9: Oscillator block diagram [13]...... 16

Figure 10: Crystal Equivalent Circuit [15] ...... 17

Figure 11: Frequency response of the crystal (a) Marker at 145.95 MHz with span of 5 MHz (b) Frequency response curve from 300 kHz to 500 MHz...... 18

Figure 12: Frequency response of the crystal compensating parasitic capacitance (a) Marker at 145.96 MHz with a span of 5 MHz (b) Frequency response from 300 kHz to 500 MHz ...... 19

Figure 13: Oscillator output as seen on the Spectrum analyzer...... 20

Figure 14: Simulation of the feedback, biasing, impedance matching for the power amplifier using GENESYS software ...... 21

Figure 15: Spectrum Analyzer display for the oscillator and power amplifier stage ...... 22

Figure 16: Beacon harmonic filter as modeled using GENESYS ...... 23

Figure 17: (a) Helical antenna mounted on the panel for far field test (b) helical antenna coated with heat shrink and kapton tape for the antenna vibe test ...... 24 vi

Figure 18: Antenna Vibration Test (a) X-axis setup (b) Y-axis setup (c) Z-axis setup [17] ...... 27

Figure 19: Umbilical EGSE...... 30

Figure 20: Front panel of the Umbilical Box...... 32

Figure 21: Rear panel of the Umbilical Box...... 34

Figure 22: Umbilical Cable...... 35

Figure 23: Isolated RS485 to USB Converter (B&B electronics) [21] ...... 36

Figure 24: RS485 to CMOS/TTL Converter ...... 36

Figure 25: Umbilical EGSE circuit board...... 37

Figure 26: Umbilical EGSE test setup...... 38

Figure 27: Umbilical Box (exploded view) ...... 39

Figure 28: CanX Ground Station Overview ...... 40

Figure 29: UHF Ground Station overview [11]...... 41

Figure 30: Setting up UHF ground station antennas (a) Antenna Rotator (b) Initial setup with two Yagi-Uda Antennas...... 42

Figure 31: Upgrading the ground station with an array of four Yagi-Uda antennas ...... 43

Figure 32: Cable getting tangled and resisting the motion of the antennas ...... 43

Figure 33: Antenna Rotator by AlfaSpid ...... 44

Figure 34: UHF ground station tracking...... 45

Figure 35: Cabling on UHF ground station ...... 46

Figure 36: Limit switch assembly (a) Limit switch open box (b) Limit switch mounted on the rotator...... 47

vii

Figure 37: Functional block diagram of the limit switch system...... 47

Figure 38: In-house built rotator controller with motors and lightning arrestors [23]...... 48

Figure 39: COTS Rotator Controller by AlfaSpid [24] ...... 49

Figure 40: SFL Ground Station Hardware [26] ...... 52

Figure 41: SFL Ground Station Software [26] ...... 54

Figure 42: Tracker software by Sumus technology [28] ...... 55

Figure 43: Terminal Interface Program (TIP)...... 56

Figure 44: (a) Snapshot (b) WODomatic...... 57

Figure 45: Nanosatellite Interface Control Environment (NICE)...... 58

Figure 46: (a) CanX-2 Status Report (b) NTS Status Report ...... 60

Figure 47: Battery voltage telemetry when CanX-2 is in orbit normal alignment mode [30]...... 62

Figure 48: Power generated versus consumed while CanX-2 was in Momentum Alignment mode [30]...... 62

Figure 49: Structural panel and battery temperatures when CanX-2 is in Orbit Normal Alignment [30]...... 63

Figure 50: ACS Auto-update plots ...... 64

Figure 51: Wheel speed plot [35]...... 65

Figure 52: EKF estimated body-rate plot during wheel spin-up [35]...... 66

Figure 53: CanX-2 B-dot rate controller damping rates from 50o/s [30]...... 67

Figure 54: CanX-2 Momentum Align Controller - alignment angle between spacecraft Y-axis and orbit normal approaching 0o [30] ...... 68

Figure 55: CanX-2 Wheel Pitch Controller - Aligning payload to point target...... 68 viii

Figure 56: CanX-2 Wheel Pitch controller - Aligning GPS antenna to zenith [30] ...... 69

Figure 57: GPS Position Estimation plotted relative to TLE estimated ground track using STK [30]...... 72

Figure 58: GPS Occultation Experiment - Attitude sphere plotting four GPS spacecraft tracked above the atmosphere, while one GPS satellite occults through the atmosphere (Figure provided by University of Calgary GPS occultation team) [32]...... 73

Figure 59: GPS Occultation Experiment [32] - Estimated position differences relative to NORAD estimated TLEs (Figure provided by University of Calgary GPS occultation team).... 74

Figure 60: Spectra of greenhouse gases taken over Ontario, Canada by Argus 1000 spectrometer (Figure provided by York University Argus Spectrometer team) [31]...... 75

Figure 61: First CanX-2 image of the Moon using the POBC imager [27]...... 76

Figure 62: CanX-2 flatsat testing in progress ...... 77

Figure 63: Body rates rising during ACS spin-up [35]...... 80

Figure 64: ACS spin-up scenario - body rates rising and then settling down [35]...... 80

Figure 65: CanX-2 magnetometer ripple ...... 82

Figure 66: Measurement of magnetic field vector in body frame during mode transitions ...... 83

Figure 67: Image showing location of sea vessels from the data provided by NTS using STK (Figure provided by COM DEV) [40] ...... 85

List of Tables

Table 1: Communications subsystem - allocated frequencies and bandwidths [8] ...... 8

Table 2: VHF Beacon Subsystem Requirements [10] ...... 10

Table 3: VHF Beacon Link Budget ...... 26

Table 4: Random vibration test profile ...... 28

Table 5: Requirements for Umbilical EGSE [18]...... 31

Table 6: CanX-2 Timeline ...... 51

List of Abbreviations & Acronyms

Nanosatellite Interface Control ACS Attitude Control Subsystem NICE Environment Attitude Determination and ADCC NLS Nanosatellite Launch System Control Computer Attitude Determination and NORth American Aerospace ADCS NORAD Control Subsystem Defense Command AIS Automatic Identification System NSP NanoSatellite Protocol

AWG American Wire Gauge NTS Nanosatellite Tracking Ships On-orbit Attitude System BL1 Bootloader 1 OASYS Software BL2 Bootloader 2 OBC On-Board Computer

BPSK Binary Phase Shift Keying PI Principal Investigator

BRITE BRIght-star Target Explorer POBC Payload On-Board Computer Canadian Advanced Nanospace CANOE PSLV Polar Satellite Launch Vehicle Operating Environment Canadian Advanced Nanospace CanX PTT Push-To-Talk eXperiment Complementary Metal Oxide CMOS PVC Polyvinyl Chloride Semi-conductor Canadian Nanosatellite Advanced CNAPS PWM Pulse-Width Modulation Propulsion System COTS Commercially Off The Shelf QPSK Quadrature Phase Shift Keying

CRC Cyclic Redundancy Check RAM Random Access Memory Energy per bit to noise power E /N RF Radio Frequency b o spectral density ratio EDAC Error Detection And Correction RMS Root Mean Square Electrical Ground Support Received Signal Strength EGSE RSSI Equipment Indicator EKF Extended Kalman Filter SAA South Atlantic Anomaly

FOV Field Of View SEL Single Event Latch-up

FSS Fine Sun Sensor SEU Single Event Upset

GNB Generic Nanosatellite Bus SFL Space Flight Laboratory

Static Random Access GPS Global Positioning System SRAM Memory Transmission Control Protocol/ GUI Graphical User Interface TCP/IP Internet Protocol HKC House Keeping Computer TIP Terminal Interface Program

IF Intermediate Frequency TLE Two-Line Elements

ISL Inter-Satellite Link TNC Terminal Node Controller

ISP Specific Impulse TVAC Thermal Vacuum Test

ISS Intersatellite Separation System UHF Ultra High Frequency Temps Universel Coordonné LEO Low Earth Orbit UTC (Coordinated Universal Time) University of Toronto Institute LFFT Long Form Functional Test UTIAS for Aerospace Studies MIB Minimum Impulse Bit VHF Very High Frequency

MOBC Main On-Board Computer VUV Vacuum Ultraviolet

MTP Mass Transfer Program WOD Whole Orbit Data eXperimental Push Out NANOPS NANOsatellite Propulsion System XPOD Deployer

xii 1

Chapter 1 Introduction

“The inspirational value of the space program is probably of far greater importance to education than any input of dollars. A whole generation is growing up which has been attracted to the hard disciplines of science and engineering by the romance of space.”

- Arthur C. Clarke, “First on the moon”, 1970.

Canadian Advanced Nanospace eXperiment program initiated by the Space Flight Laboratory, University of Toronto Institute for Aerospace Studies in 2001 has flourished into a world renowned nanosatellite space program for education and research.

1.1 Introduction

Fifty years back, with the launch of Sputnik – the first man-made satellite saw an advent of the new era called “The Space Age”. Since then, new technologies have emerged and great progress has been made in the field of space. Today, space technology is extensively used for communications, Earth observation, atmosphere monitoring, scientific experiments and educational purposes. The technological advances in the field of microsystems technology have made it possible to scale down the size of the satellites.

The author of this thesis has been grateful to be involved with CanX series of missions as part of her Master’s degree. This exclusively unique program established by Space Flight Laboratory has provided an invaluable hands-on experience to the author with overall design and development of the nanosatellite. This is reflected through the various projects carried out by the author of this thesis during her Master’s program. The author contributed to the Generic Nanosatellite Bus (GNB) with the VHF Beacon as part of the communication subsystem and with the Umbilical Electrical Ground Support Equipment (EGSE) developed to interface with the satellite for monitoring and testing purpose. The author was also involved with the ground segment development at SFL. Being a spacecraft operator for CanX-2 and NTS, the author acquired a deeper insight into each spacecraft subsystem and the mission on the whole. Her major contributions have been elaborated in the following chapters with the brief background information of the Space Flight Laboratory and the CanX missions.

1.2 Space Flight Laboratory

The UTIAS Space Flight Laboratory (UTIAS/SFL) has been working together with various business, government and academic institutions worldwide on spacecraft projects and the development of new space technologies. The principal objective of SFL is to develop and demonstrate new technologies in space. SFL, bridging a gap between academics and industry, also provide hands on experience to master’s students in the design and construction of an actual satellite thus generating skilled space systems engineers.

The microspace philosophy [2] has been utilized by SFL to achieve low-cost CanX missions with spacecraft development in less than two years. Hence, the spacecraft design can make use of the continuous development being made in the field of microsystems/MEMS technology. This is accomplished through small talented design teams and the use of commercial off the shelf (COTS) parts. This allows SFL to follow a disciplined approach for designing, building, testing and qualifying new space systems.

SFL was established in 1998 with the MOST mission. Later on, the Canadian Advanced Nanospace eXperiment (CanX) program was launched in 2001 at SFL, thus beginning many exciting adventures for students and staff at SFL. CanX missions provide cost-effective access to space for the research and development community in the world through the use of nanosatellites and picosatellites. In addition to micro and nano spacecraft projects, the Space Flight Laboratory engages in technology research to facilitate the microspace revolution. Nanosatellite Launch Service (NLS) Program, established at UTIAS/SFL, provides launch sharing arrangements for nanosatellites along with the XPOD deployment mechanism to various institutions around the world at reduced cost.

The Generic Nanosatellite Bus (GNB), a 20 x 20 x 20 cm spacecraft bus that can be customized to fit a range of missions, has been developed by SFL. This reduces cost for any mission, shortens development cycle and facilitates the use of existing technologies with space heritage. The GNB was originally developed for the BRITE and CanX-4/-5 missions, and is now being employed for other missions around the world. AISSat-1 represents first such mission and is a perfect example of the resilient GNB platform.

1.3 CanX-2

The CanX-2 Mission is the second nanosatellite within the CanX program at SFL. CanX-2, 3.5 kg in mass and 10 x 10 x 34 cm in dimension was launched on April 28, 2008 to demonstrate and evaluate novel technologies that will be used on the CanX-4/-5 formation flight mission [3]. The technologies to be tested include a novel nano-propulsion system, custom radios, innovative attitude sensors and actuators, and a commercial GPS receiver.

Magnetometer

GPS Antenna UHF Antennae (4)

Materials Science Experiment S-Band Patch Antennae (2)

Atmospheric Z Spectrometer X Propulsion Y Thruster (2)

Figure 1: CanX-2

In addition to evaluating these miniature technologies, CanX-2 carries scientific experiments for other university researchers across Canada. These include a GPS radio occultation experiment to characterize the upper atmosphere provided by University of Calgary, Argus atmospheric spectrometer to detect greenhouse gases provided by York University, and a space materials experiment that will measure the effects of atomic oxygen on advanced materials by University of Toronto [3].

1.4 CanX-6/NTS

Figure 2: CanX-6/NTS

The CanX-6 also known as NTS – the Nanosatellite for Tracking Ships was launched onboard PSLV-C9 along with CanX-2. NTS, a 6.5 kg nanosatellite, was built in less than a year by SFL in collaboration with COM DEV International Ltd. CanX-6 uses the Generic Nanosatellite Bus (GNB) developed by SFL to demonstrate the space-based Automatic Identification System (AIS) detection technology developed by COM DEV [4]. AIS, a self-organizing radio communication system, is designed to be used for identifying the location of ships. NTS tracks sea vehicles and collects AIS data from space which is then transmitted to the ground where it is analyzed for ship information [4].

1.5 CanX-3/BRITE

Figure 3: CanX-3/BRITE (Exploded View)

The CanX-3, also known as BRIght-star Target Explorer (BRITE) is a mission comprising of constellation of satellites. Each BRITE satellite is based on the GNB spacecraft bus developed by SFL. The main objective of the BRITE mission is to perform photometric observations of the brightest stars in the sky and examine these stars for variability [5]. The observations obtained will have a precision 10 times better than those achievable using ground-based observations. The secondary objective of the BRITE mission is to investigate stellar winds and their effects on future stellar life cycles. The mission payload for each BRITE spacecraft comprise of a telescope and a star tracker. The telescope detector has a large field of view of approximately 25o across in order to observe multiple stars at a time, and be able to obtain a differential photometry of the stars in brightness and colour [5]. Figure 3 shows an exploded view of BRITE. The Principal Investigator for the BRITE mission is Professor Anthony F. J. Moffat Département de physique, Université de Montréal. The Austrian Co-Principal Investigator is Professor Werner W. Weiss, University of Vienna. It is scheduled to launch in December 2010 [5].

1.6 CanX-4/-5 Formation Flying mission

S-Band Patch S-band Downlink S-band Downlink Magnetometer Antenna (1 of 2) Patch Antenna VHF Antenna Patch Antenna

Sun Sensor (1 of 6) Intersatellite S-band ISL (2 of 4) UHF Antennas Separation Patch Antenna (2 of 4) System (1 of 2)

Y X Z S-band ISL Imager CNAPS Patch Antenna Z X GPS Thrusters (4) Y Antenna

Figure 4: CanX-4/-5

CanX-4 and CanX-5, a pair of identical nanosatellites, will make use of the GNB spacecraft developed at SFL. The primary mission objective is to demonstrate on-orbit formation flying. On-orbit formation flying can be defined as two or more satellites controlling their position and orientation with respect to one another to achieve a predefined configuration necessary for coordinated operations [6].

CanX-4/-5 mission will use the technology proven aboard CanX-2 to achieve and maintain several controlled formations in orbit [6]. Formation will be controlled with the second generation NANOPS called CNAPS developed at SFL. This will evaluate the propellant usage in autonomous formation control strategies. CanX-4/-5 will also utilize innovative carrier-phase differential GPS techniques to obtain relative position measurements accurate to less than 10 cm. CanX-4 and CanX-5 will communicate using an inter-satellite communication link developed at SFL [6]. CanX-4/-5 mission is scheduled to launch in December 2010.

1.7 AISSat-1

Figure 5: AISSat-1

AISSat-1, a 6 kg Norwegian nanosatellite mission is being developed by SFL in collaboration with the Norwegian government. AISSat-1, based on the GNB spacecraft will serve as both research and development platform. The primary mission objective is to explore the viability and performance of a spacecraft-based Automatic Identification System (AIS) sensor in low-Earth polar orbit. The purpose of the mission is to track maritime assets. The space-based AIS data obtained through the mission will integrate into a national maritime tracking information system [7]. AISSat-1 is scheduled to launch in October 2009.

Chapter 2 Generic Nanosatellite Bus

Design, construction and testing are the major attributes that defines engineering. The author of this thesis designed, constructed and tested the VHF Beacon and the Umbilical Electrical Ground Support Equipment (EGSE) for the GNB missions. This involved research and simulations, but then again the real world is different from the computer simulations and various iterations were made in the design in order to meet the requirements and obtain successful results for the tests.

2.1 Communication Subsystem

The communication subsystem plays a crucial part in any space mission. The subsystem is required for communicating with the ground station about the health information of the satellite, the position of the satellite and the payload data [1]. The GNB missions have fully duplex communication system. The system will utilize the existing equipments and framework of CanX- 2 mission with required changes and improvements tailored specific to the mission. Maintaining heritage speeds up the design process, minimizes cost and allows the usage of existing allocated frequency bands thus maintaining the compatibility with the existing ground segment.

Table 1: Communications subsystem - allocated frequencies and bandwidths [8]

The GNB communication architecture is formed from the following major subsystems:

UHF Uplink: For the GNB uplink design, the commands and software will be sent to the spacecraft for all the mission functions via the amateur UHF band. The UHF uplink will be active at all times when power is applied by the spacecraft power system.

S-band Downlink: For the GNB satellites, the telemetry and payload data will be sent to the ground station via the SRS S-band, the band used for CanX-2 mission at present.

VHF Beacon: For the GNB missions, VHF beacon is optional equipment that will transmit in the VHF amateur satellite band, sending Morse code telemetry at 15 wpm containing basic information of spacecraft health.

2.2 VHF Beacon

Beacons have been widely used for air and sea navigation for years to track an airplane or a ship. Beacon transmitters are located on the object to be tracked and they transmit continuous or periodic signals carrying information like the object identification, location of the object and telemetry. This concept of beacon has been used on many satellites and space rovers. The satellite beacon is a low power carrier transmitted by a satellite which is at times modulated and provides ground station with monitoring telemetry data and tracking information of the satellite. The beacon is activated after launch of the spacecraft and provides a simple, reliable method of ascertaining spacecraft health in the event of loss of communication over the primary links.

Most amateur and educational satellites are employed with beacons. Of these, many beacons operate in the amateur radio frequency bands and can be received by the amateur radio stations around the world [9]. This feature makes it possible to track the spacecraft during loss of communication with the primary ground station confirming that the spacecraft is operational. Beacons can also be programmed to provide spacecraft telemetry and experimental data. For instance, GeneSat-1 beacon operating at 437.075 MHz FM, sends an AX.25 packet at 1200 baud every 5 seconds – the packet contains data about the spacecraft systems operation [9].

The beacon operating in the VHF band was proposed for various CanX and GNB missions after the trade-off of its advantages and disadvantages. The beacon is set to activate upon separation of the spacecraft from the deployment mechanism. Reception of the beacon will indicate that the satellite has separated. The beacon will assist in ensuring that the antennas are correctly pointed for communication. By observing the Acquisition of Signal (AOS) and Loss of Signal (LOS) time of the beacon signal one could determine which object it is in a closely spaced group. The

10 beacon’s activation will also indicate that the power system is operational and it will provide some basic telemetry (battery voltage, beacon temperature, and operating time) [11].

Addition of beacon to the satellite design also adds an additional radio system that needs to be budgeted. It occupies volume in the spacecraft, requires surface area on the exterior (its antenna), and it consumes power. Such resources are especially limited in a nanosatellite. Additional ground station infrastructure (although simple) must also be planned and assembled. After a response is received from a Ping and communication is established with the spacecraft, the beacon becomes unnecessary and can be deactivated. If one can only receive beacon telemetry but not communicate with the satellite, the primary mission objectives cannot be completed [11]. There is only one radio receiver that is necessary for operations. Upon ﬁnal consideration, it was decided to add the beacon to CanX-4/-5 mission as there was time in the schedule and it was desirable to have some information in case of failure instead of no information.

2.2.1 Mission Requirements and Design Issues

The mission requirements for the VHF beacon can be found in Table 2. The VHF beacon system is optional, but desired, equipment that may be included in all spacecraft utilizing the generic nanosatellite bus (GNB).

Table 2: VHF Beacon Subsystem Requirements [10] 8.1 VHF Beacon Requirements 8.1.1 General Requirements COM-VHF-1 The beacon shall operate on 145.890 MHz in the amateur satellite VHF band. COM-VHF-2 The beacon shall generate a CW signal, modulated by a Morse-code format data stream. COM-VHF-3 The beacon shall activate at 10 seconds (TBC) after separation from the orbital deployer. Activation is performed by the satellite’s power system. COM-VHF-4 The beacon shall function correctly even if the satellite has not completely separated from the orbital deployer. The beacon shall also function if the antennas has not completely deployed and is in contact with the orbital deployer. COM-VHF-5 It shall be possible to permanently deactivate the beacon at end of mission. (Industry Canada and IARU requirement.) The satellite’s power system is responsible for implementing this.

8.1.2 Interface Requirements COM-VHF-6 The beacon shall require only power, ground, and antenna connections to operate.

COM-VHF-7 The beacon should provide four external analog telemetry inputs which could be inserted into the downlink message.

8.1.3 Performance Requirements COM-VHF-8 The beacon shall operate with a minimum supply voltage of 3.3 V. The beacon shall operate with supply voltages up to 5.5V. COM-VHF-9 The beacon shall be receivable with a 6 dB link margin using antennas and equipment already present on SFL ground stations, assuming that the beacon antenna has fully deployed. COM-VHF-10 The beacon shall consume no more than 200 mW DC power average during continuous operation. COM-VHF-11 The beacon frequency stability shall be better than 20 ppm over its operating temperature range, including aging effects. COM-VHF-12 The beacon shall occupy a signal bandwidth of no greater than 2.2 kHz. All emissions outside this bandwidth shall be no greater than -26 dBc measured in 1 kHz bandwidth. COM-VHF-13 The beacon shall operate within specification over a minimum temperature range of -20°C to +60°C. COM-VHF-14 The beacon signal shall be receivable at a minimum ground station elevation angle of no greater than 5°. COM-VHF-15 All spurious radiated emissions shall be 60dB below the carrier.

8.1.4 Antenna Requirements

COM-VHF-16 The beacon antenna shall use materials consistent with the UHF antennas on the GNB (i.e., piano wire or hypodermic tubing). The radiated polarization shall be linear. COM-VHF-17 The length of the beacon antenna shall be 11 cm (structural constraint, less than the satellite length of 20 cm).

8.1.5 Message Content Requirements

COM-VHF-18 The beacon shall broadcast the satellite’s identifier: (e.g. BRITE, CanX-4, etc.). COM-VHF-19 The beacon shall broadcast the satellite’s amateur radio call sign (VA3SFL) at least once every 30 minutes (Industry Canada regulation). COM-VHF-20 The beacon shall broadcast input supply voltage and board temperature reading. COM-VHF-21 The beacon should broadcast the value of additional analog telemetry points. COM-VHF-22 The beacon shall broadcast a counter or timer value to indicate how long it has been operating since reset. COM-VHF-23 The beacon shall be active for at least 80% of each minute to facilitate tracking.

The primary purpose of the VHF beacon transmitter system developed by the author at SFL is to support initial acquisition and tracking of the spacecraft post-launch. This simple, reliable system will aid during the event of communication loss over the primary links by monitoring spacecraft health. The beacon is designed such that the organizations and individuals external to SFL would be able to receive and decode the beacon signal with simple ground station equipment. However, during regular day to day operations, the VHF beacon signal becomes redundant once the spacecraft responds to a ping and acknowledges commands.

2.2.2 Design

An initial design for beacon was proposed by Daniel Kekez and Lucas Stras for the CanX-2 mission. A 20 dBm (100 mW) transmitter, operating in the VHF amateur satellite band, sending Morse Code telemetry at 15 words per minute was designed. The beacon was programmed to provide a quick and easy snapshot of spacecraft health, and to assist in the early commissioning phase [11]. The beacon was not flown on CanX-2 due to certain design issues.

The initial design was very sensitive to layout changes which were affected by the parasitic capacitance and inductance on the board. The quarter wavelength monopole antenna was chosen to interface with the initial design but it was not feasible to fly that size of antenna on a nanosatellite. So, a shorter monopole with a critical complex matching network was chosen. The oscillator stage employs a Butler (grounded base) transistor oscillator with a fifth-overtone crystal. The bipolar junction transistor (BJT) was biased with an emitter feedback network configuration, as it is the passive bias system that provides the best control of D.C. gain variations from device to device and over temperature [11]. The design of the oscillator used the equations and process described by McVey [11].

The VHF Beacon designed by the author is based on the initial beacon design taking into consideration the mission requirements and the design issues faced by the previous design. The VHF beacon transmitter designed by the author can be divided into six major stages – The voltage regulation stage, the oscillator and buffer stage, the power amplifier stage, the filtering stage, the antenna and the microcontroller. Each stage is described as shown in Figure 6.

Figure 6: VHF Beacon block diagram

The PCB layout of the CanX-4 beacon is shown in Figure 7.

Figure 7: VHF Beacon

2.2.3 Background Theory

Various important terms that are being used to describe the functionality of the various stages of beacon in subsequent sections are introduced here in brief.

Fundamental frequency or the carrier is the lowest frequency component of the signal being generated by the RF system (oscillator). Along with the fundamental frequency, the system also generates harmonics which are the multiples of the fundamental frequency.

Scattering Parameters better known as S-Parameters – S11, S12, S21, S22, define the performance of several variables (gain, return loss, reflection coefficient, amplifier stability, etc.) at various frequencies. S-parameters are usually plotted in a logarithmic scale as a function of frequency.

S11 (Input Reflection Coefficient) provides information about the input return loss of the RF system. From the figure above, S11 is the ratio of the input parameters, b1 and a1. S12 (Isolation Coefficient) is the ratio of b1 and a2. It determines the level of feedback from the output of an amplifier to the input and therefore influences the stability of an amplifier along with S21. Buffer amplifier provides isolation between the oscillator and the power amplifier stage and thus has small value for |S12|. S21 (Forward Transfer Coefficient) is the ratio of b2 and a1 and provides information about the gain of the circuit. S22 (Output Reflection Coefficient) is the ratio of b2 and a2 and gives information about the output return loss.

Gain is the ratio of the power output to the power input of the amplifier in dB. The gain is specified in the linear operating range of the amplifier where a 1 dB increase in input power gives rise to a 1 dB increase in output power.

Gain = 20*log(S21)

Insertion Loss (dB) is defined as the drop in power as a signal enters an RF component. This value not only includes the reflected incoming signal, but also the attenuation of the component.

Insertion Loss (dB) = 10 * LOG10(Output Power/Incident Power)

Return Loss (dB) is defined as a ratio of the incoming signal to the same reflected signal as it enters a component. Return loss is given by S11 parameter.

Return Loss (dB) = 10 * LOG10(Reflected Power/Incident Power)

2.2.4 Voltage Regulation Stage

Figure 8: Voltage Regulator block diagram

The VHF beacon is connected to a standard power switch on the Power Board and may be commanded ON or OFF by the Main On-Board Computer (OBC). By default, on power up or reset, the beacon is switched on by the Power Board switching circuitry. A simple low-dropout regulator circuit (Figure 8) is used to provide 3.3 V to the oscillator, buﬀer and microcontroller stages. The power amplifier uses bus voltage from the Main On-Board Computer (MOBC) but the feedback of the power amplifier is powered through a low-dropout regulator (LDO).

The Power Board uses two reed switches to determine when separation has occurred. These switches detect the proximity of magnets placed on the deployment XPOD. When one of the two

switches sense separation, the power board commands the satellite to activate. Upon activation, the beacon is powered up and begins transmitting. At the end of mission, it is necessary to permanently disable all transmitters to comply with Industry Canada regulations and IARU requirements [12]. This will ensure that a system failure after end of mission does not cause the transmitter to reactivate in an uncontrolled manner. In line with the power system is a slow-blow fuse, which can be tripped on ground station command, permanently deactivating the beacon. This fuse system is managed by the power board which determines the current requirement for the beacon. From the power budget, the beacon current consumption should not exceed 112 mA.

2.2.5 Oscillator and Buffer Stage

At the heart of every transmitter is an oscillator that generates the RF carrier. An oscillator is basically an amplifier and a frequency selective feedback network as shown in Figure 9. When, at a particular frequency, the loop gain is unity or more, and the total phase shift at this frequency is zero, or some multiple of 360o, the condition for oscillation is satisfied, and the circuit will produce a periodic waveform of this frequency. This feedback network determines the frequency and stability of the generated signal.

Figure 9: Oscillator block diagram [13]

An oscillator is characterized by frequency stability, frequency accuracy and purity of the output waveform. Ordinary L-C oscillators using conventional inductors and capacitors can achieve typically 0.01 to 0.1 percent frequency stability, about 100 to 1000 Hz at 1 MHz. Crystal oscillator circuits are similar to L-C oscillator circuits, substituting the crystal for L-C

components giving better frequency stability and better frequency accuracy. Crystal oscillators have very low noise content because of very high Q.

Figure 10: Crystal Equivalent Circuit [15]

In the crystal equivalent circuit shown in Figure 10, L1, C1 and R1 are the crystal motional parameters and C0 is the capacitance between the crystal electrodes. At VHF frequencies, the current reverts to that due to C0 alone so other crystal vibration modes can be disregarded. fr is the series resonant frequency of the crystal and fa is the anti-resonant frequency of the crystal. Most crystal oscillators are designed to operate the crystal at fr where it appears resistive, or at a frequency between fr and fa, called fL, where the crystal appears inductive [15]. Above approximately 60 MHz the crystal C0 can cause non crystal controlled oscillation. Hence, at VHF it is necessary to compensate C0 by tuning the crystal. Tuning the crystals is essential in crystal oscillators for the oscillations to occur at the desired frequency fL with the crystal and not just the circuit.

For VHF beacon, a fifth overtone, series resonant, Hc-51/U packaged crystal from West Crystal Inc. cut at 145.960 MHz is used. The crystal has a 3 ppm tolerance and is rated for 5 ppm stability over -10oC to +40oC temperature range. The crystal was tuned and calibrated using network analyzer.

Figure 11 shows the frequency response of the crystal. The response should be a straight line but due to the parasitic capacitance of the crystal a curve is observed.

Figure 11: Frequency response of the crystal (a) Marker at 145.95 MHz with span of 5 MHz (b) Frequency response curve from 300 kHz to 500 MHz

The parasitic capacitance of the crystal which otherwise might cause spurious oscillations was compensated using inductor coil placed in parallel to the crystal. The inductor placed in parallel with the crystal causes anti-resonance of the crystal C0 which minimizes loading. The value of the inductor coil was calculated using the following equation:

1 1 f = 2π LC

where, f = frequency of the crystal = 145.960 MHz, C = C0 of the crystal = 8 pF, L = Inductance of the coil = 150 nH (calculated).

The plot of the frequency response of the crystal after compensation with a 150 nH inductor is shown in Figure 12.

Figure 12: Frequency response of the crystal compensating parasitic capacitance (a) Marker at 145.96 MHz with a span of 5 MHz (b) Frequency response from 300 kHz to 500 MHz

Modified common collector Colpitts fifth overtone crystal oscillator circuit was designed for the beacon as it is economical, provides better stability, is efficient, has low component count and gives high reliability over a wide temperature range. The design employs Agilent general purpose, low noise NPN silicon based bipolar junction (BJT) transistor as an amplifier and the crystal as a feedback circuit to the amplifier. The amplifier circuit and the feedback circuit were optimized to work at the beacon frequency such that the overall gain of the circuit was one. This ensures that oscillations start reliably and random fluctuations in frequency due to supply voltage and thermal variations are avoided. Additional low noise NPN silicon based BJT is connected to the output of the oscillator which acts as a buffer amplifier. The buffer provides isolation between the oscillator stage and the power amplifier stage which minimizes adverse effects at RF and enhances oscillator performance. Figure 13 shows the spectrum analyzer output obtained from the oscillator stage with a peak at frequency 145.961 MHz and a gain of -0.96 dBm.

Figure 13: Oscillator output as seen on the Spectrum analyzer

2.2.6 Power Amplifier (PA) Stage

The power amplifier stage employs NEC’s medium power NPN silicon based high frequency transistor. This stage is switched on by the microcontroller which acts as a push-to-talk (PTT) switch in order to key the continuous wave transmission. The keying is done using a single-pole, double-throw (SPDT) analog switch from Texas Instruments. The PA stage is connected to the oscillator stage via impedance matching pi-network. The feedback, biasing and impedance matching for the power amplifier stage were simulated and optimized using Agilent’s GENESYS [16] software.

Figure 14 shows the schematic of the circuit simulated, smith chart showing the input and the output return loss plots, logarithmic return loss plots (S11 [dB] and S22 [dB]), the gain (S21[dB]) and the stability factor (K) plot. Figure 15 displays the spectrum analyzer output for the oscillator and power amplifier stage with a 10 dB attenuator connected for protection. The total gain of about 24 dB is obtained at the output of the power amplifier stage at the fundamental frequency of 145.96 MHz.

Figure 14: Simulation of the feedback, biasing, impedance matching for the power amplifier using GENESYS software

Figure 15: Spectrum Analyzer display for the oscillator and power amplifier stage

2.2.7 Filtering Stage

Oscillator generates harmonics. Harmonics are an integer multiple of the output fundamental frequency generated by the oscillator. The second harmonic is the frequency twice the fundamental and third harmonic is thrice the fundamental. For GNB missions, the UHF receiver operates at 435.XX MHz and the third harmonic produced by the beacon is at 145.XX MHz times three equals 435.XX MHz. This results into communication interference and hence is not desirable. As per the requirement, the third harmonic should be 40 dB below the fundamental frequency. Filter was designed with third harmonic suppressed 45 dB below the fundamental frequency.

A pi-filter with lumped components was designed as a band stop filter, filtering out the second and the third harmonics. The filter was modeled and optimized using GENESYS as displayed in Figure 16. The blue line in the graph represents input return loss, S11 [dB] of the harmonic filter and the red line represents the gain (S21 [dB]). The filter stage was added at the output of the power amplifier stage. The design was successfully tested on the prototype.

Figure 16: Beacon harmonic filter as modeled using GENESYS

2.2.8 VHF Antenna

Initially, it was proposed to use a monopole antenna made from high carbon steel music wire, something similar used for the UHF antennas for CanX missions. According to the beacon requirements, antenna should fit within the envelope of the UHF antennas. But at 145 MHz, it is desirable that the antenna be a quarter-wavelength long of 48 cm. So, the VHF monopole antenna of length 17 cm was proposed for the initial design but it posed various challenges. It was found to have impedance of 2.7 − j39.1 Ω when deployed which is very small compared to the 50 Ω source and load impedance of any standard radio-frequency RF system [11]. Hence, a highly optimized matching network was required between the antenna connector and the ampliﬁer stage to deal with sub-optimal antenna length. This was difficult to achieve and the impedance mismatch resulted in a higher gain loss. For the link budget calculations, -6 dBil gain is assumed for the beacon’s monopole antenna as it is the minimum theoretical gain with 80% sphere coverage.

Research was carried out by the author to find an antenna that would meet the given requirements and would give better performance. After thorough investigation, the helical antenna from Smiley Antennas was chosen. The antenna is formed of helical wound hardened

steel wire copper plated for lower resistance with maximum power of 50 W operating in the VHF range. This 11 cm antenna is a reduced size quarter wavelength with far field radiation pattern similar to that of a quarter wavelength monopole with about 35Ω base impedance. For the link budget calculations, -3 dBil gain is assumed, but the antenna tests has shown better results with gain of up to +1 dB.

Kapton tape

Heat shrink

Figure 17: (a) Helical antenna mounted on the panel for far field test (b) helical antenna coated with heat shrink and kapton tape for the antenna vibe test

2.2.8.1 Link Budget

Link budget, as the name implies, provides detail account of all the gains and losses that happens from the transmitter to the receiver in the RF communication system. The link budget consists of key terms – transmitter power, receiver and transmitter antenna gains, antenna feed losses, path losses and receiver sensitivity. Link budget can be summarized with the following equation:

Received Power (dBm) = Transmitted Power (dBm) + Gains (dB) – Losses (dB)

Received power is the power received by the receiver. Transmitted power is the power transmitted by the VHF beacon. Gains are the transmitter antenna gain and the receiver antenna gain. Losses include transmitter losses (feed harness, antenna mismatch, etc.), free space loss,

25 atmospheric loss, polarization mismatch loss and receiver losses (feed harness, pointing loss, etc.).

Ideally, the beacon’s signal would be accessible to any radio amateur with a simple set-up. The requirement, however, is that it can be received by the ground station at SFL/UTIAS with a 6 dB link margin. For the link budget calculations, the frequency of 145.96 MHz generated by the beacon is considered. The existing VHF ground station set-up located at SFL/UTIAS consists of a circularly polarized VHF Yagi antenna (15.14 dBic gain) mounted on the MOST ground station tower and connected to a Yaesu amateur radio receiver. Feed harness loss and antenna mismatch loss are estimated for the link budget calculation.

The link budget is calculated for the monopole VHF antenna and for the helical VHF antenna to show the comparison between the two. A -6 dBil gain is assumed for the beacon’s helical antenna of 11 cm in length to be on the conservative side. Helical antenna is finalized to interface with the beacon transmitter as it provides better link margin.

Link margin for the VHF beacon employed on GNB mission is tabulated in Table 3. The values estimated for the link budget are very conservative. Downlink margin provided in the table accounts for the total gain obtained after the losses. The following constants were assumed for the link budget:

Constants:

Speed of Light in Free Space 299.79 x106 m/s Earth's Equatorial Radius 6378.15 Km Standard Temperature 300.00 K G/S Antenna Temperature 100.00 K Boltzmann's Constant 1.38E-23 -198.60 dB(mW/K*s)

Table 3: VHF Beacon Link Budget

Inputs Calculations1 Calculations2 Unit Notes Assigned in IARU Frequency 145.96 MHz Coordination Process Wavelength 2.0549 m Transmit power (mW) 100.00 20.00 20.00 dBm Nominal Output Feed harness loss 1.5 -1.50 -1.50 dB Estimated Antenna Mismatch Loss (VSWR) 2 -0.51 -0.51 dB Nominal Estimated – 1 Antenna is 17 cm Monopole loaded to 50 ohm 2 Antenna is 11 cm helical Antenna gain -6.00 -3.00 dBil antenna EIRP 11.99 14.99 dBm Satellite orbital altitude (circular orbit) 900 km Maximum orbital altitude Minimum elevation 5 degrees Maximum distance to satellite 2993.71 km Free space loss -145.25 -145.25 dB S/C antenna is a monopole LP feed. G/S antenna is a CP Polarization mismatch loss 3 dB Yagi. Atmospheric loss 1 dB Total propagation loss -149.25 -149.25 dB Isotropic signal power at Antenna Input -137.26 -134.26 dBm G/S antenna is a KLM 2M- Antenna gain 15.14 15.14 dBic 22C, 13 dBdc gain Antenna beam width (half G/S antenna is a KLM 2M- power) 34.00 degrees 22C, 13 dBdc gain Pointing error 3 degrees M2 rotator on MOST tower Pointing loss -0.09 -0.09 dB Feed Harness loss 0.5 -0.50 -0.50 dB Estimated Preamplifier Noise Figure 0 dB No Pre-amplifier used Preamplifier Gain 0 dB No Pre-amplifier used Cable Loss 3 dB 200 ft of LMR-400 at 145 MHz NF for FT-897 provided by Receiver Noise Figure 4.5 dB Yaesu tech support System Noise temp (K) 1692.87 32.29 32.29 dBK Referenced to antenna/feed G/T -17.24 -17.24 dB/K harness interface. Referenced to antenna/feed Receiver Signal Power -122.22 dBm harness interface. Referenced to antenna/feed Receiver Noise Power -132.89 dBm harness interface. C/No 44.10 47.10 dB Standard filter bandwidth in Receive Bandwidth 2200 33.42 33.42 dBHz CW mode C/N 10.67 13.67 Implementational Losses 1 1.00 1.00 dB Baseband S/N 9.67 12.67 dB Required S/N for CW 0 0.00 0.00 dB Estimated for Morse Code Downlink Margin 9.67 12.67 dB

2.2.8.2 Antenna Vibe test

The vibration and shock load tests were conducted by Celestica’s Technology Assurance group in accordance with UTIAS/SFL specifications to qualify VHF antenna. The tests were conducted on two helical antennas – one coated with heat shrink and other coated with the kapton tape. The tests consisted of sine sweeps, sine bursts and random vibration with the test profile as required for PSLV launch vehicle as supplied by SFL. Random vibration test duration was extended to 30 minutes in the X and Y axis, or until failure was noted. There where no mechanical failures noted during the test. Visual checks were performed on the antennas after each test.

Figure 18: Antenna Vibration Test (a) X-axis setup (b) Y-axis setup (c) Z-axis setup [17]

During sine burst test, the antennas were subjected to 9.75G peak for 1 second between 9.9 and 10.1 Hz. One burst per axis. During sine sweep test, the antennas were subjected to 0.8 G peak at 5 – 10 Hz, 3.0 G peak at 10 – 100 Hz with 2 oct/min, one sweep per axis [17]. Under random vibration test, the antennas were subjected to a random profile as tabulated in Table 4.

Table 4: Random vibration test profile

The duration for random vibration test was 80 seconds in the vertical direction and 30 minutes or to first failure in the other two horizontal axes [17]. Both the antennas survived all the tests.

2.2.9 Microcontroller Stage

A low-power Microchip PIC microcontroller (PIC16F771) was chosen from the initial beacon design to control this circuit. This PIC has one-time programmable EPROM memory that stores the program code and is less likely to be susceptible to degradation from radiation [11]. The PIC16F771 was selected because it has a built-in 12-bit ADC and an internal clock circuit, thus reducing the number of required parts [11]. The ADC is used to sample the bus voltage, 3.3V rail and read an on-board temperature sensor. The microcontroller stage consumes 2 mW of power. The microcontroller code was initially written by Daniel Kekez. The code was modified by the author to incorporate the addition of +3.3 V rail and CanX-4/-5 satellite identification.

2.2.9.1 Beacon Message

The beacon will transmit the following message in Morse code format:

HIHI, the message used by the first amateur satellite OSCAR I, will initiate the VHF beacon message. Next in the message is the satellite identification for eg., CANX-4 and then the SFL radio amateur call sign. HHHH represent 16-bit values output in hexadecimal. The ﬁrst telemetry point is the bus voltage, the second is the +3.3V rail and the third is the temperature as measured by the sensor located on the beacon board. The ﬁnal hexadecimal sequence is a serial number

which is incremented each time the message is transmitted, providing an indication of how long the beacon has been operating. Letter K at the end denotes end of message.

The Morse code (On-Oﬀ Keying) encoding scheme was selected for its simplicity. At 15 WPM, it is a very low data rate and can be coded by ear or electronically in very low-SNR conditions. It would enable the UTIAS/SFL ground station to verify its antenna pointing from the beacon signal. In cases of blackouts or communication losses, other radio amateurs can decode the bus voltage and report this data back to UTIAS/SFL [11].

2.2.10 Beacon PCB Layout and Testing

Based on the beacon requirements provided for GNB missions, the initial design, research and simulation carried out using Ansoft Designer and Agilent GENESYS; the circuit was designed by the author. Next step was to select the required components and verify the circuit. The PCB layout for VHF beacon was carried out using Altium Designer software. In order to avoid parasitic oscillations, VHF circuit layout practice was followed. In-built component libraries from Altium were used for footprints. The completed design was printed with 1:1 scale and the component fit test was carried out. Design Rule Check (DRC) was executed to check for errors. The gerber files were generated for the final design and the Printed Circuit Board (PCB) design was manufactured by Sierra protoexpress. The components were populated on the manufactured PCBs and stage by stage testing was carried out. While the beacon prototype was tested in the lab, frequency, power consumption and the output gain were measured. Daniel Kekez, Mihail Barbu and Milko Dimitrov provided guidance with component selection and reviewed the PCB layout. The power output measured was into a 50 Ω load. In these tests the beacon was operating at the frequency of 145.960 MHz. The beacon’s message also correctly stated the supply voltage, the +3.3V rail and the ambient temperature.

In accordance with UTIAS/SFL Thermal Shock Procedures, the beacon flight board needs to undergo functional tests and thermal vacuum tests at the operational temperature of +50oC and - 20oC. Vibration tests for the beacon will be carried out when the beacon has been integrated into the spacecraft.

2.3 Umbilical EGSE

Umbilical Electrical Ground Support Equipment (EGSE) or the Umbilical box as known in general was designed and built by the author to control GNB flatsats and/or spacecraft. The general requirements for the design of the umbilical box were provided by Mihail Barbu and Daniel Kekez. The components were selected and the mechanical and electrical layouts were carried out by the author. Mihail Barbu reviewed the layouts and provided assistance with testing of the umbilical box.

2.3.1 Motivation

Figure 19: Umbilical EGSE

The umbilical box was developed as an interface to the satellite for testing and monitoring purpose. The idea was to design a simple setup that would regulate battery charging/discharging, monitor battery voltages, monitor test port voltage and monitor test port current. On top of this, it would also provide various modes – Spacecraft ON/OFF mode, OBC ON/OFF mode, Transmitter Enable/Disable mode, Charging Enable/Disable mode; along with electrical isolation and shielding, and USB interface. The box also incorporates over voltage protection circuit and reverse voltage protection circuit. So, in all, this box would carry out all the functions yet is simple to operate.

2.3.2 Requirements

The general requirements for the umbilical box for GNB missions are tabulated in Table 5.

Table 5: Requirements for Umbilical EGSE [18] 6.1 General Requirements 6.1.1 The umbilical box shall be able to turn the spacecraft ON when the satellite is inside XPOD. 6.1.2 The umbilical box shall be able to turn the spacecraft OFF when the satellite is not inside the XPOD. 6.1.3 The umbilical box shall be able to enable/disable all transmitters over the test port. 6.1.4 The umbilical box shall be capable of routing charging current (at an appropriate voltage and amperage) to the batteries. 6.1.5 If the spacecraft ON/OFF switch is in the OFF position, the umbilical box shall keep the spacecraft OFF even if it is powered. 6.1.6 The umbilical box shall provide protection for over-voltage. 6.1.7 The umbilical box shall provide the spacecraft with a grounding link. 6.1.8 The umbilical box shall monitor the voltages of up to two batteries on the satellite. 6.1.9 The umbilical box shall be capable of enabling and disabling the OBC over the test port. 6.1.10 The umbilical box shall protect the spacecraft against reverse voltage. 6.1.11 The umbilical box shall provide spacecraft power for testing operations. 6.1.12 The umbilical box shall have appropriate manual switches and LED status lights for spacecraft ON/OFF, Transmitter Enable/Disable, Charging Enable/Disable, OBC ON/OFF. 6.1.13 The umbilical box shall prevent the spacecraft from being turned ON, if the charging voltage is zero. 6.1.14 The umbilical box shall be able to turn on and off the charging supply where off charging supply turns off the spacecraft. 6.1.15 The umbilical box shall interface communication link to the PC. 6.1.16 The umbilical box shall electrically isolate the PCs from the spacecraft to avoid ground loops. 6.1.17 All connectors on the umbilical box should be D-Sub, USB, Power or Banana terminal.

2.3.3 Design/Features

Some of the important features and functionality of the umbilical EGSE designed and built by the author are described below:

Figure 20: Front panel of the Umbilical Box

2.3.3.1 Safety Key switch:

This is a key lock SPDT switch. When the safety key is in the ON position, the umbilical box will not allow the spacecraft to be turned ON without also having Charging Enabled. When in the OFF position, the spacecraft can be turned ON with Charging disabled which can result in battery discharge. During power outages, if safety key is in ON position, satellite will shut down thus protecting the battery by avoiding battery discharge.

2.3.3.2 Battery Charging Supply:

The pair of banana jacks on the front panel connects to an external power supply and provides power to the spacecraft. Power will not flow through these lines unless main Power switch is enabled (charging enabled).

2.3.3.3 Main Power Switch:

This Double-Pole Double-Throw (DPDT) switch enables/disables the flow of power to the spacecraft from the battery charging supply jacks. When enabled, the battery charge display LED (yellow blinking LED) blinks with a period of approximately 0.5s. Unless otherwise required charging should be enabled any time the spacecraft is ON.

2.3.3.4 Satellite ON/OFF switch:

This toggle Three-Pole Double-Throw (3PDT) switch turns ON and OFF the spacecraft. When the spacecraft is ON, the Sat ON LED is ON (green LED). This LED draws power from the umbilical box power supply and not the spacecraft power supply. The switch does not require power. Hence, if the umbilical box is unpowered, the switch can still activate the spacecraft. But in this case, the LED would not indicate it.

2.3.3.5 OBC ON/OFF switch:

This toggle DPDT switch controls the both the OBCs simultaneously. It forces the OBCs ON in the absence of the ability to firecode the OBCs ON through the UHF receiver. The OBC LED (green LED) turns ON when the switch is enabled. If an OBC has been enabled via firecode then switching OBC ON/OFF switch to OFF position will not turn the OBC OFF.

2.3.3.6 Tx Enable/Disable switch:

This toggle DPDT switch enables and disables the use of the transmitters (S-band and VHF) on the spacecraft. When enabled, the Tx LED (red blinking LED) blinks at a period of 0.5s. It provides electrical shielding and over-voltage protection to the transmitters.

2.3.3.7 Battery Voltage Display:

Two Miniature Display Module DC Voltmeter (MDMV) located on the front panel displays the battery voltages of the two batteries in the spacecraft. The MDMV is a small panel mount digital voltmeter with the integrated circuit bonded directly to the printed circuit board thus reducing the space requirements. The MDMV module operates from a +5 VDC supply and measures input voltages from 0 to ±199.9 mVDC relative to the supply common. Auto-zeroing is provided by the module, therefore no zero adjustment is required. If the input signal exceeds the input range, the unit display indicates an over-range condition. The module has three selectable decimal point

positions and five selectable annunciators (BAT, V, A, m, µ). The BAT (Battery) annunciator is used to indicate battery level for the MDMV, while the others are used to indicate Volts and Amperes.

2.3.3.8 Battery Voltage Terminals:

The two pairs of banana jack terminals can be used to measure the battery voltage using an external multimeter if desired.

Figure 21: Rear panel of the Umbilical Box

2.3.3.9 Test Port Connector:

This DB-25 connector forms a connection between the satellite test port and the umbilical box through an umbilical cable. This connector must never be connected or removed when the satellite is Charging or in any way active.

2.3.3.10 Fuse

This 7A fuse limits the current flowing to the umbilical box. It does not limit the current to the spacecraft.

2.3.3.11 Umbilical Box power supply:

The P5 power connector provides power to the umbilical box. A 9V power converter must be used to allow the umbilical to be plugged directly into a standard power outlet.

2.3.3.12 USB connector:

The USB cable connected to this terminal allows control of the spacecraft from a PC. In some boxes, two USB terminals are provided which allows the box to control the spacecraft from two PCs at the same time.

2.3.3.13 Ground terminal:

A banana terminal provides an additional ground point that can be used as a common grounding point if required.

2.3.4 Umbilical Cable

Figure 22: Umbilical Cable

About 5 feet in length, it connects satellite to the umbilical box. One end has the DF-11 connector that interfaces with the testport and the other end has DB25 connector which interfaces with the umbilical box. The long cable allows communication to the satellite during TVAC testing when the satellite is in the thermal chamber. While using the umbilical cable to test the radios, the cable being long, it acted as an antenna thus picking up random noises from the surrounding and distorting the data. This issue was resolved by shielding the cable using silver plated copper braid.

2.3.5 Communication

The testport on the satellite’s power board consists of two 3V CMOS-level asynchronous serial links working at the rate of about 1Mbps. It utilizes CMOS/TTL to keep the design simple and for lower power consumption. TTL doesn’t require drivers as compared to RS232 or RS485 to function, thus minimizing the parts on the satellite.

Figure 23: Isolated RS485 to USB Converter (B&B electronics) [21]

Isolated USB to RS485 converter from B&B electronics is used to interface with the computer. RS485 is chosen over RS232 for electronic data communication. RS485 carries out differential data transmission which helps in nullifying the effect of ground shifts and induced noise signals. RS485 can communicate at data rates of up to 100Kbps at distances of up to 4000 ft. This supports communication during various satellite tests.

Figure 24: RS485 to CMOS/TTL Converter

Figure 24 shows the RS485 to CMOS/TTL converter designed by the author and reviewed by Mihail Barbu. RS485 to CMOS/TTL converter, powered from the umbilical box, is connected in between the umbilical cable closer to the satellite end as shown in Figure 26.

Figure 25: Umbilical EGSE circuit board

Figure 25 displays the umbilical EGSE circuit board designed by the author and reviewed by Mihail Barbu. It houses the over voltage protection circuit, the 5V and 3.3V voltage regulators, the battery voltage monitoring circuit and the circuitry for the switches located on the front panel.

2.3.6 Umbilical EGSE Testing

Figure 26: Umbilical EGSE test setup

The umbilical EGSE connects to the satellite as shown in Figure 26 for testing and monitoring purpose. An extensive testing was carried out to each umbilical box before interfacing it with the spacecraft. The test procedure was written and executed by the author under the guidance of Mihail Barbu. The procedure was followed, to test each and every feature in detail. Till date, six umbilical boxes – two for UNIBRITE, two for BRITE-Austria and two for AISSAT-1, have been designed and built by the author for various GNB missions and flatsat testing.

Figure 27: Umbilical Box (exploded view)

Figure 27 shows the exploded view of the umbilical box. Open box test [20] was carried out which included checking the voltage at the test points, checking the polarity of the capacitors, measuring fuse integrity, verifying the ground terminals and ground pins, power pins located on the DB-25 connector, and verifying the battery voltage display meters. Closed box tests include general tests like checking the continuity, checking switches and LEDs. Test was carried out to check over voltage protection circuit by increasing the voltage until the power supply turns into constant current mode. Reverse protection diode test was conducted by reversing the polarity of the banana plugs at the battery charge supply input. The umbilical cable was qualified by continuity test between the DB-25 connector and the DF-11 connector and the pins on the RS485 to USB widget. The switches were tested in various configurations in different ON/OFF states and the resistances and voltages were measured. Communication test was carried out with the help of hyperterminal window to send and receive data and with the help of power board ECHO software loaded on two OBCs with speed of 921600 baud rate.

Chapter 3 Ground Station 3.1 Introduction

Figure 28: CanX Ground Station Overview

A highly capable ground station was developed and located at SFL, Toronto to support CanX-2 and future CanX missions. As part of the nanosatellite program, the ground station was required to fit within the low cost nanosatellite framework while providing support for multiple missions. The ground station was designed to be modular so that future upgrades would have minimal affect on the system. The ground station features both COTS and custom made components.

The CanX ground station is required to support communication over S-Band downlink, UHF uplink and VHF downlink. The block diagram shown in Figure 28 gives an overview of the ground station. It consists of the S-band/VHF station and the UHF station. The S-band and VHF station makes use of the existing hardware from the MOST ground station developed at SFL. The S-Band station includes the 2.1m S-band parabolic dish and the radio for receiving the S- band signal. The VHF station consists of VHF Yagi-Uda antenna and radio for receiving VHF signal. The UHF station was developed for CanX-2 and included an array of four Yagi-Uda antennas. All the antennas are mounted on rotator motors that allow the antenna to track the satellite thus varying the antenna elevation and azimuth during the entire duration of the contact period. Both UHF and S-band radios are connected to the TNC (Terminal Node Controller). The

TNC acts as the central hub of the ground station. Its primary purpose is to facilitate the uplink and downlink communications between the ground station radios and the ground station user application software. It also calculates the degree of Doppler shift and automatically adjusts the UHF radio frequency to compensate [22].

3.2 UHF Ground Station

Figure 29: UHF Ground Station overview [11]

The UHF station, located on the roof of the UTIAS’s main building, houses the UHF Yagi-Uda antenna array (consisting of four yagi antennas), commercial azimuth/elevation rotator, rotator controller and multifunction TNC. The schematic of the UHF station is displayed in Figure 29. The station was initially developed for CanX-1 but has been modified for CanX-2 to overcome various issues. The author of this thesis along with Daniel Kekez was involved with setting up of the UHF antennas, installing the limit switch and making the antenna rotator system more robust and efficient.

Ground station forms an essential aspect of any space mission. During the design of the space mission, ground station at times takes the back seat due to less critical requirements and constraints compared to the spacecraft itself. Same issue was faced by the CanX-1 mission, where the ground station wasn’t fully functional when the spacecraft was sent to the launch site. For CanX-2, special attention was given to the development of ground station well in advance. The ground station was qualified prior to the launch by carrying out tests with the flatsats and the flight hardware thus testing the communication links.

3.2.1 UHF Antenna System

Initially, the UHF station comprised of two 435 MHz 42-element circularly polarized (with polarization switch) Yagi-Uda antennas combined in phase providing a total gain of 21 dBic. Initial design built for CanX-1 mission consisted of a system of two antennas. The UHF antenna system was updated to comprise of four Yagi-Uda antennas, driven in phase, to increase the gain of the ground station to the level required in the analysis of the communications link for CanX-2. An additional ~3 dB gain was acquired by adding two additional Yagi-Uda antennas.

Figure 30: Setting up UHF ground station antennas (a) Antenna Rotator (b) Initial setup with two Yagi-Uda Antennas

Figure 31: Upgrading the ground station with an array of four Yagi-Uda antennas

The antenna array was erected on a tripod like structure along with the antenna rotator. The mast can be easily lowered to a more manageable height for installation and maintenance. During operation, the structure is raised by about 10 feet above the roof. At the base of the structure is a weather proof enclosure enclosing the ground station main power supply (13.8 V), 100W PA, LNA, signal lightning arrester and a pair of ventilation fans.

3.2.2 Antenna Rotator

Various issues were posed by the rotator during CanX-1 mission operations. One of the key issues faced was the cable getting tangled around the antenna boom, thus jamming and preventing the rotator from moving freely. The commercially available rotator made by Yaesu was used for the pointing of an antenna boom in two axes.

Figure 32: Cable getting tangled and resisting the motion of the antennas

During a spacecraft pass, it is commanded to point its high gain directional antennas at the spacecraft to facilitate communications. Due to unrestricted motion in the azimuth direction, the cables accidentally got tangled and wrapped around the antenna boom, as a result jamming and preventing the rotator from moving freely. The controller for the rotator continued to command the rotator motors to move, but since they were mechanically restricted from doing so, quickly burnt one of the motors out. There were temperature sensors on the motors that were designed to automatically shut the rotator down if it exceeded a safe level (i.e. caused by the motor shaft stall), but they reacted too slowly and the net result was that the rotator was severely damaged and repair costs would have been nearly the cost of a new unit [23].

Figure 33: Antenna Rotator by AlfaSpid

CanX-2 design utilizes a COTS heavy-duty amateur radio grade Azimuth/Elevation rotator manufactured by AlfaSpid as shown in Figure 33. It has one 12V-24VDC drive motor for each of the two axis both of which are equipped with a 1 degree resolution position switch. This switch closes momentarily every time the axis moves 1 degree [24]. By using these pulses, it becomes possible to integrate the absolute pointing direction of the rotator. On the Elevation axis, limit switches are present that allow it to rotate between -21º to 201º, thus providing 180o full travel in the elevation axis. It also features limit switches on the elevation axis that automatically cuts power to the motor in the event of over travel. The Azimuth axis is not restricted in rotation and can turn continuously. The rotator does not provide any electrical or mechanical limit direction in the azimuth axis even though this is needed.

Limit switch assembly, designed and built by the author, was employed to prevent over travel and cut power in that direction, similar to that found on the elevation axis. This limit switch also serves as a means to recalibrate and self zero the rotator. It gives a known reference point for the antenna pointing direction. This can be done periodically to avoid accumulating position errors. The design of the limit switch is described in detail in the following section.

Power to the motors and signals from the position switches are fed through two water tight connectors (one connector for each axis, with 4 conductors on each). These connectors are mounted through a custom back plate that bolts to the rear of the rotator [23]. Two 20 foot 4 conductor harnesses made of 18 gauge wire for each of these connectors then run from the rotator to the rotator controller located in the roof mounted equipment box [23]. For CanX-2, a more reliable and robust rotator system is designed making use commercially available products as well as some custom designed control elements to reach the goal of better longevity and availability of the system.

Figure 34: UHF ground station tracking

To avoid tangling of the cables while the antenna system is tracking the satellite, the cable were secured properly with tie wraps leaving enough slack for the free motion in both the azimuth and the elevation axes. Figure 34 and Figure 35 displays the cable arrangement for the UHF antenna system. The UHF antenna array was rotated to verify unobstructed motion in both the elevation and azimuth axes.

Figure 35: Cabling on UHF ground station

3.3 Limit Switch Assembly

Figure 36: Limit switch assembly (a) Limit switch open box (b) Limit switch mounted on the rotator

Figure 37: Functional block diagram of the limit switch system

To overcome the cable tangling issues and for the efficient calibration procedure, the author designed a system to restrict and control the rotation in the azimuth plane. After extensive research and thorough investigation, the system was built as shown in Figure 36(a) with two limit switches mounted with their contact levers side by side and two diodes used to define the direction of current associated with each switch so that the correct switch limits the correct motion [25]. Figure 37 shows the functional block diagram of the limit switch system. The power wire going to the azimuth motor is then routed to this switch configuration. A bracket

attached to the rotator mast (Figure 36(b)) is used to trip against the switches. The limit switch system has been attached to the outer housing of the rotator and has been functioning as desired. The rotation limits in the azimuth axis are 360o and 0o. At these points the limit switch hits the bracket mounted on the mast, trips and slews the antenna in other direction.

3.4 Rotator controller

Initially, the in-house rotator controller developed by Tarun Tuli was used to control the movement of the antenna rotator. The issue was that the controller board couldn’t effectively handle the clicking of the position switches embedded in the rotator. The switch is supposed to close every time the axis moves 1 degree, but due to the inaccuracies involved in the movement it caused errors in the calibration.

Figure 38: In-house built rotator controller with motors and lightning arrestors [23]

Initially, the time was spent to debug the software and make the controller more robust and add on the required features.

Figure 39: COTS Rotator Controller by AlfaSpid [24]

After thorough investigation and research, the COTS controller from AlfaSpid was chosen by the author to interface with the rotator controller. The controller was initially tested with the ham radio deluxe software to track amateur satellites and receive data. The controller was programmed to detect the custom added limit switch assembly in the azimuth direction. The rotator controller has been fully functional and controls UHF ground station for CanX-2 and NTS.

Chapter 4 Spacecraft Operations

“Failure is not an option”

– Gene Kranz (NASA Flight Director for Gemini and Apollo programs)

Amidst all the excitement, there was a breeze of nervousness and tension in the air. April 28, 2008; it was the launch day. Onboard the Indian rocket, PSLV-C9, were the two Canadian nanosatellites waiting eagerly to be launched beyond the Earth’s atmosphere, to the space.

4.1 Introduction

Spacecraft operations form an integral part of any space mission. The operations structure incorporates people, procedures, hardware and software in harmony to complete the mission tasks. Satellite mission operations comprise of activities performed by operations team during the flight phase of the mission, along with pre-launch activities including development of operational framework, policies and data flow. It focuses on the period after launch, but the preparations for the same are done well in advance. Operational requirements are considered at each phase of the mission and form the basis of the operations manual. The operations manual, outlining the plan of action for the mission right after the launch, is the compilation of documents describing the operations, satellite commissioning, nominal operation procedures, contingency procedures and the ground station architecture.

The author of this thesis has been a spacecraft operator post commissioning phase for CanX-2 and NTS. Section 4.1.2 and 4.1.3 gives an overview of the ground station hardware, ground station software and flight software used by the operator during operations. Operations were carried out by the operations team comprising of Daniel Kekez, Karan Sarda, Nathan Orr, Mirue Choi and the author. The operations are carried out when the spacecraft is in view by the ground station located at UTIAS/SFL. Everyday there are about three contacts with each satellite for about 5 to 15 minutes in the morning and three contacts in the evening. The author has been handling the morning operations since July 2008. Operations manual and operations procedures were written by the operations team and the engineering team prior to the launch. These procedures are followed and updated by the operations team as and when required. CanX-2 is a

51 complex satellite with an active attitude control system and scientific payloads like NANOPS, GPS, Spectrometer and POBC imager. Table 6 tabulates the timeline for the CanX-2 operations. Section 4.2 describes in detail the author’s involvement with CanX-2 on-orbit operations. On the other hand, NTS has a passive attitude control system and carries an AIS payload provided by COM DEV. NTS on-orbit operations performed by the author is described in section 4.3.

Table 6: CanX-2 Timeline

CanX-2 Schedule Major Anomalies 28-Apr-08 Launch

May-08 Commissioning -Y FSS Anomaly Jun-08 NANOPS NANOPS Anomaly Jul-08 NANOPS Jul-08 to Aug-08 ACS Commissioning Aug-08 NANOPS Aug-08 to Sep-08 Code Upload MOBC Voltage Drop Sep-08 to Nov-08 ACS Commissioning Anomaly Nov-08 to Dec-08 GPS Engineering Magnetometer Ripple Dec-08 Spectrometer Anomaly GPS (Engineering and Jan-09 to Feb-09 Occultation) MOBC Voltage Drop Feb-09 to Apr-09 Spectrometer Anomaly GPS Occultation No data/ Apr-09 to May-09 GPS Occultation less data May-09 to Jun-09 Spectrometer Jul-09 to Aug-09 POBC Imager

4.1.1 Launch

CanX-2 and NTS (CanX-6) were part of the NLS-4 and NLS-5 launch, the cluster of nanosatellites, launched aboard the Antrix/ISRO PSLV-C9 from the Satish Dhawan Space Center in Sriharikota, India on April 28, 2008 at 03:53 UTC. They were launched into a 635 km sun synchronous orbit with a 9:30 am descending node.

4.1.2 Ground Station

The primary ground station for CanX-2 and NTS is located at UTIAS Space Flight Laboratory, Toronto, Canada (043˚ 46’ 57” N, 079˚ 27’ 54” W, ~190 m). The backup ground station for CanX-2 and NTS is under planning. The requirement of the CanX-2 /NTS ground station is that it must be able to uplink and downlink all required data within about six 5-15 minute passes o greater than or at elevation of 3 everyday. Both the satellites communicate with UHF uplink and S-band downlink, thus making it convenient to share MOST S-band ground station at SFL.

The ground station hardware design for CanX-2 and NTS can be seen in Figure 40.

Figure 40: SFL Ground Station Hardware [26]

The S-band ground station, consists of a 2.08m diameter S-band parabolic dish antenna with boresight gain of 32.2 dBic, an S-band Commercial Off-The Shelf (COTS) rotator controller, augmented with custom mounting hardware used to provide azimuth and elevation controls, an S-band receiver operating at 2.2344 GHz that tracks and automatically locks onto the carrier signal at that frequency, and the TNC (Terminal Node Controller) which acts as the central hub of the ground station. The primary purpose of TNC is to facilitate the uplink and downlink communications between the ground station radios and the ground station user application software. The antenna and the rotator controller are the only pieces of hardware that are shared between CanX-2/NTS and MOST.

The UHF ground station was designed and built specifically for CanX-2/NTS and is not shared with MOST. The UHF station houses the UHF Yagi-Uda antenna array (consisting of four Yagi antennas driven in phase to increase the gain of the ground station required by CanX missions), commercial azimuth/elevation rotator, rotator controller and multifunction TNC. The UHF antenna ground station is described in detail in the ground station section of the thesis. The UHF antennas are also incorporated with the polarization switches that allow the UHF signal to switch from Right Hand Circularly Polarized (RHCP) to Left Hand Circularly Polarized (LHCP) and vice versa thus avoiding large nulls in the satellite antenna gain pattern as the satellite attitude changes.

The major blocks of ground station software architecture is similar for both CanX-2 and NTS. Each piece of software is customized for each mission as per the mission requirements. The detailed block diagram for the ground station software module is displayed in Figure 41.

The primary software modules for spacecraft operation are Snapshot, WODomatic, NICE and TIP. Most of the operations software is automated for nominal operations. Snapshot and WODomatic send out automated emails after telemetry and WOD are acquired respectively.

Figure 41: SFL Ground Station Software [26]

Tracker software (Figure 42), custom made by Sumus technology for MOST is also being used for CanX-2 and NTS to determine the desired satellite in view of the ground station with the help of the Two-Line Elements (TLE) provided from NORAD. Tracker then sends the tracking information to the rotator controller which in turn commands the antennas to track the desired satellite. Two instances of Tracker are run, one to control the S-band antenna and is part of the MOST ground station and the other to control the UHF antenna which is part of the CanX-2 ground station.

Figure 42: Tracker software by Sumus technology [28]

TIP (Figure 43) is at the focal point of the ground station software network, managing communication between the TNC and Snapshot, WODomatic, NICE. TNC consists of UHF communication, UHF control and S-Band communication ports. UHF communication port

receives NSP packets from TIP and forwards them to the UHF modem, which is also on the TNC. UHF control is responsible for selecting which satellite is being contacted by the UHF ground station, either CanX-2 or CanX-6. It is also responsible for receiving the firecode signal from TIP and generating a signal to send to UHF communication to ensure that the correct firecode is sent to the satellite.

Figure 43: Terminal Interface Program (TIP)

Snapshot (Figure 44(a)) provides an autonomous snapshot image of the spacecraft telemetry containing the health information of the satellite along with the current spacecraft software mode. WODomatic (Figure 44(b)) is used to download the Whole Orbit Data buffer from the spacecraft. A snapshot of the current spacecraft alignment is emailed to all the operators for CanX-2 after the WOD is downloaded.

Figure 44: (a) Snapshot (b) WODomatic

NICE (Figure 45) serves as a primary user interface for testing, operations and debugging of the spacecraft. It is customized for both CanX-2 and CanX-6 in accordance to the mission requirements.

Figure 45: Nanosatellite Interface Control Environment (NICE)

4.1.3 Flight Software

Flight software forms an important aspect of spacecraft operations along with the ground software. The flight software state determines the spacecraft mode and based on that operator can decide the operations activities and experiments that can be carried out. Various flight software updates have been made on CanX-2 post launch to increase downlink efficiency, improve payload configuration capability, performance and data collection frequency.

Bootloader-1 (BL1), the default startup mode following a spacecraft power cycle, is stored in a pre-programmed EEPROM and is the lowest-level software state [30]. BL1 has no automation and offers only basic functionality such as polling real time telemetry and powering up most spacecraft systems and components.

Bootloader-2(BL2), stored in the spacecraft FLASH memory, builds on the functionality of BL1 and includes the ability to store spacecraft telemetry once per minute for over to 24 hrs so that the engineering team can review the spacecraft health state across several orbits [30].

CANOE, an SFL developed operating system, is stored in on-board FLASH memory. It is a multithreaded operating system and is the highest level of software on CanX-2. CANOE is loaded upon completion of the commissioning activities in BL2 [30]. It allows multi-tasking of operations and full spacecraft functionality. CANOE is responsible for running the On-orbit Attitude System Software (OASYS).

OASYS, which includes an Extended Kalman Filter, calculates the attitude state vector based on attitude sensor inputs and commands actuators to attain a desired attitude state.

4.1.4 Reports and Documents

Various reports and documents consisting of many details of work done during spacecraft operations are necessary for smooth operations. Documents like operations procedure, contingencies procedure and operations manual are written and maintained for quick reference. Operations manual contains information about the ground station software and hardware, nominal operations, commissioning operations, spacecraft ACS modes and contingency operations. Operation procedures are separate documents with detailed procedures for loading CANOE on the satellite, ACS mode transitions and scientific payload and engineering experiment. Operation logs containing information about the FLASH memory map, ACS settings, OBC software events, software crashes, anomalies, experiment logs, event logs, operations time-line and lessons learned are maintained and updated everyday along with time stamps for efficient spacecraft operations. Detailed anomaly reports are written and archived by the operators following an anomaly and its diagnosis. Pass reports including daily status of the satellite with health information, ACS mode, spacecraft mode and a note for other operators to follow during the next contact period are emailed everyday. Figure 46 shows the template of the pass reports used by CanX-2 and NTS which provides daily status updates of the satellite to the operators, engineers and Principal Investigators (PI) to the mission.

Figure 46: (a) CanX-2 Status Report (b) NTS Status Report

4.2 CanX-2 On-Orbit Operations

4.2.1 CanX-2 Commissioning

CanX-2 commissioning was initiated ten hours post launch when the satellite was first acquired from SFL ground station. Healthy telemetry and the confirmation of the core components functioning properly was obtained during the first acquisition thus providing the operators with the green signal for commissioning. The commissioning procedure involved incrementally building on the spacecraft functionality by activating hardware while loading the software to interface with the hardware.

The on-board computer subsystem was commissioned during the first week of operations. CanX-2 began collecting science data within four days of launch as the first science payload – Atomic Oxygen (AO) resistant materials experiment was activated. The operators managed to commission ACS hardware by the end of the second week, thus allowing the initiation of the NANOPS experiment. By the end of June 2008, all the units on CanX-2 were commissioned including the GPS and the spectrometer payload. During the course of commissioning of CanX-2 in LEO, several interesting observations were made and numerous operations lessons were learned.

4.2.2 CanX-2 on-orbit operations

Over the course of fifteen months in orbit, the operations team has been making steady progress towards achieving CanX-2 mission goals, conducting scientific and engineering experiments and solving contingencies and anomalies. The nominal operations commenced after the commissioning of CanX-2 was successfully completed. In nominal operations, CanX-2 is set in three-axis stabilized also known as the momentum alignment mode. This allows payload experiments which require three-axis stabilization of the spacecraft.

The CanX-2 on-orbit operations have been carried out at typical uplink data rate of 4 kbps and downlink data rate of 32 kbps. The S-band transmitter onboard CanX-2 can handle data rates from 32 kbps to 1Mbps under both BPSK and QPSK modulation scheme. As of June 2009, the data rate of 256 kbps has been demonstrated in orbit and over 370 MB of science and engineering telemetry data has been downloaded. The ground station at SFL is licensed to operate at data rate of 128 kbps in BPSK mode and 256 kbps in QPSK mode.

The structure of nominal operations is similar to that of commissioning operations, with operation procedures that guides the operator during satellite passes. The important aspect of nominal operations is to monitor health of the satellite and gather telemetry. CanX-2, being a technology demonstration mission for CanX-4/-5, initial stages of operations was spent qualifying the ACS system and the NANOPS. Full time scientific payload experiments began in November 2008.

4.2.2.1 General Telemetry

The satellite’s telemetry and attitude information is in the downloaded WOD files. CanX-2 is usually in momentum alignment mode when the spacecraft’s Y-axis (refer Figure 1) is aligned to the orbit normal. Typical CanX-2 telemetry includes generated power, consumed power, battery voltage and temperature telemetry. Nominally, the spacecraft battery voltage cycles between 3.9 V and 4.1 V. Figure 47 presents the bus voltage observed on a typical day.

4.2

Charge voltage limit 4.1

3.9 Discharge voltage limit Voltage Voltage (V)

3.8 Battery voltage cycling between 3.9V and 4.1V as designed 3.7 3.6V battery voltage expected when 5W Pass in eclipse transmitter activated in eclipse 3.6 9/25/2008 21:36 9/26/2008 2:24 9/26/2008 7:12 9/26/2008 12:00 9/26/2008 16:48 9/26/2008 21:36 9/27/2008 2:24 Date (UTC) Bus Voltage (V)

Figure 47: Battery voltage telemetry when CanX-2 is in orbit normal alignment mode [30]

Over a day’s span, in the nominal attitude, the spacecraft’s time-average power consumption is ~1.25W, where the time-average power generation is on the order of 5W, keeping the spacecraft highly power positive. Figure 48 presents the power consumption and power generation observed on a typical operations day.

Peak power consumption when transmitting. Power generated ~5W in Nominal Controlled 7 Attitude (Y-Thomson Configuration)

Nominal Power (W) 3 consumed power 2

0 9/25/2008 21:36 9/26/2008 2:24 9/26/2008 7:12 9/26/2008 12:00 9/26/2008 16:48 9/26/2008 21:36 9/27/2008 2:24 Date (UTC) Battery charging Generated Power (W) Consumed Power (W)

Figure 48: Power generated versus consumed while CanX-2 was in Momentum Alignment mode [30]

CanX-2 panel and battery temperatures have also remained within the expected temperature range. In the nominal attitude configuration, the spacecraft’s panel temperatures range between 6oC and 45oC, and the battery temperature ranges between 19oC and 30oC, with a typical sunlight-to-eclipse variation of 6oC. Typical CanX-2 temperature telemetry is shown in Figure 49.

50 Maximum panel temperature: 41oC 45 Battery: 23oC

20 Temperature (°C)

5 Minimum panel temperature: 7oC 0 9/25/2008 21:36 9/26/2008 2:24 9/26/2008 7:12 9/26/2008 12:00 9/26/2008 16:48 9/26/2008 21:36 9/27/2008 2:24 Date (UTC) +X +Y +Z -X -Y -Z Battery

Figure 49: Structural panel and battery temperatures when CanX-2 is in Orbit Normal Alignment [30]

Attitude telemetry contained in the WOD is also obtained during each CanX-2 contact period. Attitude telemetry contains important parameters like the alpha angle (the angle between the Y- axis of the satellite and the orbit normal), the spacecraft’s ACS mode, body rates and quaternion values. This information is also plotted in the ACS auto update plots as shown in Figure 50 that are emailed to the operators.

Figure 50: ACS Auto-update plots

4.2.2.2 ACS commissioning and nominal operations

Commissioning of CanX-2’s ACS covered significant fraction of time during commissioning. The process involved ensuring a safe power up of all sensors, actuators and reaction wheel. To date the attitude determination solution has met the requirements and is accurate to 1.5 degrees in sunlight. The performance of OASYS and its EKF has been good so far with the on-board ACS code running stably since the ACS was initiated [30]. The TLEs which are used for the on-board orbit propagator are updated weekly by the operators.

ACS actuators successfully passed the performance evaluation during commissioning. The three orthogonal magnetorquers were each actuated in both polarities at maximum current output for 30s and their effect on the spacecraft attitude was as expected. The magnetometer has been

sampling the magnetic field of the Earth since CANOE was loaded on CanX-2. It was fully calibrated on-orbit and gave successful readings consistent with the expected value for the ascending node and altitude.

There are two kinds of sun sensors on-board CanX-2 – fine sun sensor and coarse sun sensor. The fine sun sensor (FSS) uses a CMOS detector for accurate determination of the solar vector in the body frame and the coarse sun sensor is based on diffuse filtering of a phototransistor. Each FSS has a field of view of 88o due to its filter geometry and the recession within the spacecraft body. There are small dead bands between the fine sun sensors which appeared to have an effect on the attitude determination accuracy during solar vector transits. During these transits the coarse sensors provides satellite attitude data. Hence, during nominal operations, CanX-2 has full attitude control.

During commissioning, the Sinclair/SFL reaction wheel was activated and initialized with expected current consumption. The preliminary tests involved spinning the wheel up to 25 rad/s and 50 rad/s in both positive and negative directions. Figure 51 plots the performance of the wheel as it was spun from 25 rad/s to 50 rad/s and despun to 0 rad/s. During commissioning the wheel was spun in speed-control mode whereas during nominal operations the wheel was spun using torque-control.

Figure 51: Wheel speed plot [35]

Figure 52 shows the induced body rates estimated by EKF during the wheel spin up which is mounted along the Y-axis of the spacecraft. The values obtained were as expected. When the wheel was spun up to 25rad/s it imparted a 0.175rad/s (10o/s) spin on the spacecraft Y-axis

whereas when the wheel was spun up to 50rad/s, the body rates around the Y-axis was 0.35rad/s (20o/s).

Body-axes: x y z

Figure 52: EKF estimated body-rate plot during wheel spin-up [35]

Wheel has also been showing solid performance during orbit-normal and pitching operations.

4.2.2.3 B-dot mode

B-dot control has been successfully demonstrated on-orbit. B-dot also called “rate damping control” has been used successfully to detumble the satellite from high spin rates. Satellite has been commanded to B-dot mode in various situations during on-orbit operations, either testing NANOPS or during anomalies when the rates went high such that the normal rate damping had led to spin up situations rather than spin down. During spin ups, operator needs to use high rate damping method which involves reversing the B-dot control gain and bypassing the EKF temporarily to reduce the time between magnetometer reading and torquer actuation. This procedure is followed at the direction of ACS specialist. High rates have also been damped by using the reaction wheel to soak up the high rates and applying rate-damping control while slowly despinning the wheel. Successfully recovery has been made from rates of about 190o/s [30]. Figure 53 plots a typical B-dot performance on-orbit as a function of Root Sum Squared (RSS) angular rate when the satellite was detumbled from 50o/s over the span of 7.5 orbits.

Figure 53: CanX-2 B-dot rate controller damping rates from 50o/s [30]

4.2.2.4 Orbital Normal Alignment mode

Orbit-normal or the momentum alignment is the nominal ACS mode for CanX-2. The approach uses bias in the wheel combined with B-dot control to ensure that the wheel’s axis aligns with the orbit normal, which represents minimum energy solution. In CanX-2’s sun-synchronous orbit, this vector is continuously varying in the inertial reference frame thus adding lag and nutation to the alignment tolerance [30]. To date, CanX-2 routinely achieves alignment to 5o plus or minus the tolerance. Figure 54 shows orbital alignment history with the plot of alignment angle.

Figure 54: CanX-2 Momentum Align Controller - alignment angle between spacecraft Y- axis and orbit normal approaching 0o [30]

4.2.2.5 Wheel pitch Control mode

Figure 55: CanX-2 Wheel Pitch Controller - Aligning payload to point target

Payload operations make use of the satellite’s wheel pitch control mode to point the payloads. The wheel which is usually in the momentum alignment mode, changes the reaction mode to slew CanX-2 around its minor axis when wheel pitch control is commanded. During this mode, the torquers are used to trim the momentum in the wheel. To date, payload pointing performance has been achieved to about 2o. Figure 55 shows the plot of the payload pitch angle error as it drops to about 0o, when wheel pitch mode is commanded, thus maneuvering CanX-2 to point to the payload target. Figure 56 shows the plot of CanX-2’s wheel pitch controller as it aligns GPS antenna to point towards zenith.

Figure 56: CanX-2 Wheel Pitch controller - Aligning GPS antenna to zenith [30]

4.2.3 Science and Engineering Experiments

The on-orbit payload experiments were commenced after demonstrating the cutting edge technologies and the SFL built subsystems.

4.2.3.1 Atomic Oxygen – Resistant Coating Experiment

During commissioning, within a week from launch, CanX-2 began collecting science data from its material science experiment developed by the Integrity Testing Laboratory, Toronto and the University of Toronto. The experiment collects resistance and temperature data once a day from four aluminum samples, coated with atomic oxygen resistant coating. From the observations obtained to date, minimal degradation in the samples has been observed [30]. Experiments will be continued until the end of the mission which will allow for longer exposure times giving better conclusions to the results.

4.2.3.2 Nano Propulsion System (NANOPS)

NANOPS experimentation being the highest priority payload was carried out from May to August 2008. The demonstration of the NANOPS technology will form the basis of the Canadian Nanosatellite Advanced Propulsion System (CNAPS) to be used for CanX-4/-5 mission. During that time span, experiments were carried out to characterize the system performance, to evaluate fuel leakage performance and to quantify minimum-impulse bit (MIB) performance.

o Prior to the launch, NANOPS was filled with sulfur hexafluoride (SF6) fuel at 20 C at a pressure of 522 psi as recorded between the regulator and thrust solenoid valves. When the regulator valve is actuated, the pressure of the secondary volume equalizes to that of the fuel tank. During commissioning, about two and a half days after the launch, the propulsion system was powered on briefly and the obtained telemetry was verified to be in the expected range (at 461 psi at 15.2oC) [30]. About ten days after the initial power-up, the regulator valve was actuated to perform the leak test check. The results obtained from this test indicated that there was minimal leak in the fuel tank and it was concluded that the NANOPS had survived the launch.

Minimum impulse bit tests were performed at the beginning of the nominal operations to establish the smallest impulse that can be imparted by the propulsion system by progressively shortening the thrust-valve actuation time taking into account propellant pressure and

temperature. Prior to each MIB test, the NANOPS secondary volume was pressurized to just under the vapour pressure at the polled propellant temperature [30]. Throughout the course of testing, the NANOPS thrust valve actuation times were varied from 1 to 500 ms. Few short duration thrust tests were conducted to observe the effect on the attitude subsystem. The thrust magnitude was estimated at maximum 35 mN and the MIB was observed to range from 0.07

mNs at 75 psi to 0.15 mNs at 255 psi. The theoretical maximum Isp for SF6 is 50 s and the observed average Isp on-orbit was 46.7s.

Long duration thrust testing, to measure impulse and thrust levels at various pressures, is planned to be conducted after GPS payload testing has been completed. The final test by NANOPS will be performed when all of the propellant is expended and this will confirm ground measurements of the magnetic effect of the solenoid valve actuating on the magnetic field reading of the magnetometer.

4.2.3.3 GPS Engineering Experiment

GPS engineering experiments have been carried out to evaluate GPS receiver data quality, performance in orbit and to evaluate single-point position determination accuracy. During the experiments, various parameters like the GPS antenna attitude, the receiver on-time, the logging frequency, supplied initial time and the position estimates were varied to observe the functioning of device in different scenarios. During cold starts, it was noticed that the GPS antenna attitude highly affected the position-velocity-time (PVT) results.

For the first set of engineering experiments, the GPS antenna was pointed towards zenith. In this configuration, with the GPS receiver on-time being about 15 minutes, CanX-2 acquired lock to four or more GPS satellites and returned a position estimate.

For the second set of experiments, the GPS antenna was pointed towards the horizon, in the anti- velocity direction, there were no PVT estimates obtained even though the GPS receiver was on for about 20 minutes. But the position estimates were obtained when the antenna was pointed in attitudes between the zenith and the anti-velocity direction. This has been confirmed during the GPS occultation experiments performed later on after warm-starting the receiver [30].

The solutions obtained from the experiments have been compared the GPS receiver estimates obtained from the NORAD TLEs. These GPS estimated latitude and longitude results were

plotted with the orbital ground track estimated by NORAD TLEs and the snapshot of the plot is displayed in Figure 57. The solutions obtained to date have been within the maximum TLE error range of ±20 km.

Figure 57: GPS Position Estimation plotted relative to TLE estimated ground track using STK [30]

4.2.3.4 GPS Occultation Experiment

GPS occultation experiment, designed by the University of Calgary [32] is carried out using the GPS receiver and the GPS antenna aboard CanX-2. The aim of the experiment is to characterize water vapour and electron density concentrations in the troposphere and ionosphere which can prove useful in weather applications and also can be used to improve GPS position estimate accuracy. For a single GPS occultation event to take place, a minimum of five GPS satellites, with four of the GPS satellites in view above the atmosphere and one tracked GPS satellite occulting through the atmosphere, must be tracked continuously to avoid position estimate degradation by atmospheric effects.

Occultation experiments on CanX-2 commenced in January 2009, with the first set of experiments running until March 2009 and then the second set running from April to June 2009. The focus of the experiments was to commission the experiment, and work out issues related to timing, GPS receiver clock drift and optimal attitude pointing. First successful occultation observation was obtained at the end of May 2009. Attitude sphere plot plotted in Figure 58

shows five observed GPS spacecraft as observed on June 1, 2009, with anti-velocity direction pointing downwards, orbit-normal pointing left, zenith is the center and nadir is the outer circle of the plot. For this experiment, GPS antenna was pointed half-way between zenith and the anti- velocity direction. In the plot the green circle is the Earth, the blue oval is the antenna field of view and the cyan circle is the upper boundary of the atmosphere. The observed GPS spacecraft are shown in red and the variation in the colour is the time spent logging at 50 Hz. So, four satellites are observed in the plot above the atmosphere with one occulting.

Figure 59 graphs position difference (radial, in-track, cross-track) between the GPS receiver estimate from the same June 1st observation, relative to TLE-estimated ground track.

Figure 59: GPS Occultation Experiment [32] - Estimated position differences relative to NORAD estimated TLEs (Figure provided by University of Calgary GPS occultation team)

GPS occultation experimentation is still underway and the accumulated data is currently being analyzed by the University of Calgary team in order to retrieve atmospheric profiles.

4.2.3.5 Spectrometer Experiment

The Argus spectrometer experiment developed by York University detects greenhouse gases such as CO2 and water vapour in the Earth’s atmosphere by observing in the near-infrared band (900nm to 1700nm) [31]. About fifty spectrometer experiments have been conducted over time – commissioning in May to July 2008, September 2008, December 2008; nadir-tracking experiments from February to April 2009 leading to successful collection of valid greenhouse gas spectra at targets of interest all over the world. A sample spectra plotted obtained over Ontario, Canada by the spectrometer aboard CanX-2 is shown in Figure 60. The coloured lines represent three different spectra readings of greenhouse gases during the same observation. At long wavelengths, carbon dioxide exhibits a characteristic absorption fingerprint that can be seen in the spectra on the right hand side.

Figure 60: Spectra of greenhouse gases taken over Ontario, Canada by Argus 1000 spectrometer (Figure provided by York University Argus Spectrometer team) [31]

4.2.3.6 POBC Imager Experiment

CanX-2 carries two imagers on-board, which are not mission critical components. There is a colour imager on the MOBC and a monochrome imager on the POBC. Figure 61 show the first successful image obtained in August 2008 by the POBC imager which appears to be the Moon (TBC).

More POBC imaging is being carried out in July and August 2009. The author made an attempt to capture Earth limb shot images with the POBC imager. The instrument vector and the angle were set such that the POBC imaging instrument was pointed towards the Earth horizon. Taking Earth images during sunlight meant dealing with excess light. This resulted in oversaturation of the lens. The imaging parameters like the image window size, the integration time, the row delay and frame delay, black target configuration were modified during each contact period to obtain images. The choice of the parameters to acquire images was done during flatsat testing carried out by the author. During this process, it was noted that the window size affected the imager integration time directly. Decreasing the window size and row delay reduces the total integration time for capturing an image.

The flatsat testing for the POBC imaging experiment was carried out by Karan Sarda and the author. The tests were carried out both inside the lab and outdoors. During both the instance, an effort was made to vary the imaging parameters and observe and understand there effects on the captured image. During indoor testing, clear image of the ceiling was obtained. During outdoor testing, it was difficult to obtain a sharp image. The images obtained were highly saturated. A filter was used to acquire better image outdoors during daylight. But CanX-2 POBC imager does not have a filter. This makes it difficult to take images of the Earth during sunlight. At present, attempts are being made to acquire successful images.

Figure 61: First CanX-2 image of the Moon using the POBC imager [27]

4.2.4 Flatsat Testing

Figure 62: CanX-2 flatsat testing in progress

The testing on the flatsat, an arrangement of spare satellite equipment, in the lab forms an integral part of the operations activity. All the subsystems and components spares were developed along with the flight hardware during the design process of CanX-2. The spare components and subsystems form the flatsat of CanX-2 and has been used for testing and imitating flight situations during development of Canx-2, during pre-launch period, during commissioning and during nominal operations.

"Test what you fly, fly what you test" strategy has been followed at SFL for successful operations. All the procedures are tested on the flatsat before being executed on the satellite to make sure that they work as expected and that they meet operational requirements and constraints of the subsystems. Failure analysis also requires using flatsat testing on the ground thus allowing an operator to simulate, recreate and analyze the problems observed in flight.

4.2.5 CanX-2 Anomalies/Contingencies

Contingency operations are very crucial to satellite operations in case of emergencies. It acts as the first line of defense against any anomaly that is encountered on CanX-2. The contingency operation procedures detail the steps that should be taken when a fault occurs on CanX-2. The immediate actions taken during an anomaly can be as significant as saving the satellite or could be as minor as ensuring that the satellite is still healthy. Since more than one option may be available to an operator during an anomaly, the contingency operations are meant to ensure that the desired steps are taken to reduce the severity of the problem. The steps defined in contingency operations are followed by the operators during these anomalies. Any anomaly that is encountered is logged in the operations anomaly log with the detailed report stating the time of the anomaly, cause of the anomaly and the actions taken during the anomaly. At times the contingencies are discussed with the operations team in order to provide a procedure to deal with major anomalies and flatsat testing is carried out to resolve the anomalies.

Some of the anomalies faced by CanX-2 over the course of operations are described in the following paragraphs.

4.2.5.1 Software crashes

There have been about 34 software crashes since launch with a period of about a fortnight [28]. So far, six software crashes have occurred in SAA, thirteen within passes and seven over the poles. Some of the crashes occurred due to radiation hit, some were due to MOBC voltage drops whereas some were due to operation errors and have been helpful in learning important operations lessons. At the beginning of CANOE Commissioning, CanX-2 experienced its first computer crash due to a single event upset (SEU) in the South Atlantic Anomaly (SAA). Throughout commissioning and later operations, there were a few similar computer crashes; however CanX-2 was designed to handle such events by resetting into Bootloader 1 safe hold mode, protecting the system from any damage. As an operator, the software crashes were handled, by firecoding the satellite and reloading CANOE. Each software crash was logged in the anomaly log with the time of the anomaly, the cause, the method of investigation and its

solution. The time of the crash was found from the TIP log. The time obtained was used along with the satellite TLEs to locate crash and also for finding the source of the anomaly.

4.2.5.2 ACS Spin-ups

ACS spin-ups have occurred on several occasions during CanX-2 on-orbit operations. To date, there have been various instances when ACS spin-ups have occurred. During one of such spin- ups, the spacecraft spun up to 190o/s about the Z-axis. Spin-up was caused because B-dot was commanded even though the body rates were significantly below the actual rates. Usually, after firecoding the satellite and loading CANOE, the satellite is set into Passive mode before transitioning to the momentum alignment mode. Only when the rates are really high, B-dot is commanded for damping the rates. Hence in this scenario, B-dot should not have been commanded. This resulted in negative gains which in turn led to spinning up the spacecraft. The rates were poorly estimated since the pass was in eclipse, and the EKF was activated in eclipse. When the EKF is activated in eclipse, ensure that at least the Y-body rate is greater than X and Z rates and the Y-body rate is still rising. If the rates have not reached, then wait until the EKF has converged in sunlight in order to acquire confidence in the rate estimation [35]. These lessons were later on implemented by the operators in the procedures to be followed for ACS mode transitions.

During another instance, the spacecraft had spun up about the Y-axis [28]. This happened during the mode transition from Bootloader 1 to Momentum Alignment mode. Upon polling the body rates, the Y-body rates were reported to be lower than expected. The Y-body rates were transferring to the X and Z axis relatively quickly. As an operator, the author assuming that the rates were being incorrectly estimated and forced the satellite in Momentum Alignment mode. This anomaly is analogous to the wheel spin up anomaly explained in the later section. This caused the wheel to spin-up imparting high torques on the spacecraft and spinning the satellite to above the B-dot threshold. The B-dot controller, with negative gains then further spun up the satellite. Momentum alignment mode should not have been commanded as the entry conditions were not appropriate. For future cases like this, it was made sure that the ACS body rates were interpreted correctly by the operator, the wheel was allowed to despin to ~ 0 rad/s and the spacecraft Y-body rates were as high as possible. This changes were incorporated in the the

Momentum Alignment transition procedures to prevent this from happening in the future. Figure 63 shows the plot with body rates rising in Y-axis.

Figure 63: Body rates rising during ACS spin-up [35]

Figure 64 shows the plot with the ACS spin-up and the body rates settling after the procedure were followed.

Figure 64: ACS spin-up scenario - body rates rising and then settling down [35]

4.2.5.3 Wheel Despin Anomaly

On several occasions during the transition from the ACS mode Momentum Align to CanX-2 crashing to Bootloader 1 and vice versa, it had been observed that the wheel speed would spin-up and settle at higher values thus lowering the body rates in Y-axis. It was expected that that the Y- body rate would reach 85deg/s and the wheel would reach ~0rad/s every time within minutes of firecoding, but this was not being observed during some instances. After few of these incidents, it was established by the ACS specialist and the operations team that there is a clear relationship between the time elapsed from firecode to setting the Manual Actuator Control (MAC) and Wheel Speed Lock (WSL) to 0 in Passive mode, and the wheel speed upon initial polling. To avoid this issue, MAC and WSL are set to 1 by default upon reloading CANOE, which issues no commands to the wheel and allows it to despin. As soon as the WSL and MAC are set to 0, torque commands are issued to the wheel which allows the wheel to remain at constant speed. The Momentum Alignment mode transition and CANOE load procedures were modified to specify a five minute wait time between firecoding the satellite and setting MAC and WSL to 0, in order to ensure the wheel completely despins and avoid wheel spin-up incident.

4.2.5.4 NANOPS anomaly

There were two anomalies related to NANOPS being faced by CanX-2 during operations. NANOPS valve timing issue occurred when actual on-time was different than commanded. This was corrected by issuing time synchronization POKE before executing the NANOPS experiment. NANOPS thrust anomaly happened when higher than expected thrust were obtained. This was corrected by thrusting below vapour pressure [28].

4.2.5.5 MOBC Voltage Drop Anomaly

The CanX-2 MOBC power supply regulates the board’s input voltage to 2.89V ± 0.02V. Till date during three instances, the MOBC voltage has dropped to values ranging between 2.84V and 2.80V. During this event, the battery current was observed to have risen by approximately 130mA and the temperature by about 5 oC. During two of these instances, the spacecraft MOBC

82 crashed to Bootloader 1 within twelve hours. Following the crash, it was observed that the MOBC voltage rose to the expected values and the MOBC under voltage anomaly was cleared. But, it was also observed during operations that firecoding CanX-2 did not clear the MOBC under voltage anomaly. This anomaly is being investigated by Daniel Kekez.

4.2.5.6 Magnetometer Ripple Anomaly

Magnetometer ripple was observed after each spectrometer experiment and few GPS experiments pointing in anti-velocity direction [28]. This ripple occurring on 5V power line affected the alignment of the satellite as the magnetic field was incorrectly estimated. During on- orbit observations, it was also observed that the ripple disappeared when the spacecraft entered eclipse and returned in the sunlight. The ripple was not observed in passive mode whereas it reappeared when CanX-2 entered the momentum alignment mode. The issue was cleared by firecoding the satellite but that is not efficient to perform after each experiment.

Figure 65, provided by Stuart Eagleson, is from a ACS log as CanX-2 comes out of eclipse [36]. Ripple starts at roughly the 462 s mark. The coarse sun sensors register dawn at about the 452 s mark and full sunlight by roughly the 475 s mark. The fine sun sensors start being read at the 459 s mark, the official EKF end of eclipse, though their measurements remain poor during the dawn period.

Figure 65: CanX-2 magnetometer ripple

Figure 66, provided by Stuart Eagleson, displays measurements in the body frame from a log recording during a mode transition: momentum alignment engages at the 3 s mark, taking over from passive observation. It can be seen that the ripple doesn't fully kick in until roughly the 35 s mark, though the data don't rule out a build up during first ~30 s [36].

Figure 66: Measurement of magnetic field vector in body frame during mode transitions

The source of the magnetometer ripple was identified by Daniel Kekez and the author during flatsat testing. Normally, the torquers are turned off 200-250 ms before sampling the magnetometer. After the execution of the spectrometer experiment, the torquers go off <50 ms prior to sampling of the magnetometer. The residual magnetic field was concluded to be affecting the measurements and was found to be the source of the ripple. So, when the experiment is flagged as over, the spectrometer thread neither flushes the packets from the spectrometer serial port queue nor tries to read them. But the CANOE keeps waking up the spectrometer payload as the queue is not empty. Thus, CANOE wakes up experiment thread and goes back to sleep. This continuing cycle does not allow experiment thread to clear packets which affects ACS timing that turns off the torquer.

4.2.5.7 GPS Occultation experiment – no data/less data

GPS occultation observations carried on the 5th and 6th June 2009 did not produce GPS data [28]. PEEKs following the observation return 0 new packets. Full READS of the 20100000 address

return the same data set as the last valid observation as acquired on 4th June 2009. Flatsat testing has concluded that the scheduled occultation scripts are valid i.e. unique data was generated each time by the GPS. The problem was thought to be either software related (CANOE) or hardware related (GPS). The GPS telemetry looked correct from the downloaded WOD. An engineering GPS script was scheduled (which has been known to produce data in the past) as a debugging effort, and that had generated new unique valid data, so this inferred that the GPS payload was working fine. The firecode post GPS script execution temporarily cleared the issue. On debugging the issue further it was found that the scripts were missing the delays required to execute and hence not all the lines in the script were executed. This was fixed by uploading a CANOE patch provided by Luke Stras and by modifying the scripts with delays in between.

4.3 NTS On-Orbit Operations

Being launched on the NLS-5, the launch vehicle confirmed that the NTS XPOD Deployment system has been actuated and ejected the NTS spacecraft. The first acquisition of signal of NTS and its first set of spacecraft telemetry was received from the SFL ground station during its second pass over Toronto at 15:13:18 UTC, which indicated that NTS is in good health.

Bootloader 2 was loaded shortly after the first telemetry snapshots, and Whole Orbit Data recording began. After approximately a day in Bootloader 2, WOD indicated that all systems are healthy and CANOE, the application code, was loaded into MOBC. Although NTS operations were on a reduced schedule due to CanX-2 priority, within five days of the launch, a payload script was loaded an initial measurements were performed.

The majority of NTS on-orbit operations comprise of downloading telemetry and WOD, uploading NTS Payload (NTSP) time tag scripts, downloading payload telemetry and data, and resolving anomalies [38].

Despite the memory errors and increased current draw anomalies faced by NTS, the NTSP observations have been successfully performed. The NTS mission has completed its objectives and is considered a success. As of this date more than sixty observations have been performed.

Figure 67 shows the image plotted by COM DEV Ltd. using AGI STK shows the location of sea vessels for the observations made over a year.

Figure 67: Image showing location of sea vessels from the data provided by NTS using STK (Figure provided by COM DEV) [40]

Chapter 5 Conclusion

The CanX program at the Space Flight Laboratory has been continually pushing the envelopes and the boundaries of what these nanosatellites can perform. After over a year of on-orbit operations, the capabilities of CanX-2 as a 3.5 kg spacecraft have exceeded all expectations for the use of nanosatellites. The development of GNB as has demonstrated the potential of the microspace approach and the ability to have low-cost, fast access to space for small payloads, technology demonstrations and scientific experiments. This has been demonstrated by NTS and will be supported by future missions likes BRITE, Formation flying and AISSAT-1.

This thesis presented a detailed summary of the experience obtained by the author being part of the student team at SFL. The communication subsystem is an essential subsystem for any space mission. The author got an opportunity to contribute and enhance her RF and microwave knowledge and skills designing and testing the VHF beacon transmitter for the GNB missions. The prototype has been completed and the next stage is to have the flight board ready with all the space qualifications tests executed.

The testing forms an integral part of any space mission. The umbilical EGSE for GNB missions as presented in this thesis was designed, built and tested for monitoring spacecraft health and carrying out system level qualification tests. The umbilical EGSEs for BRITE, AISSAT 1 and UNIBRITE missions and flatsats have been tested and are fully functional.

A significant portion of the UHF ground station integration and testing at SFL provided hands on experience working with the ground station hardware. Spacecraft on-orbit operations provided intriguing insight into the missions’ capabilities and deeper understanding of all the spacecraft subsystems. This experience provided the author with confidence to maneuver and control the spacecraft and perform engineering and scientific experiments by utilizing different ACS and software modes. It has been a learning experience and an exciting journey filled with adventure and thrill operating the CanX-2 and NTS for over a year now.

References

[1] J.R. Wertz, and W.J. Larson, Space Mission Analysis and Design, 3rd. ed., Kluwer Academic Publishers, Dordrecht, Holland, 1999. [2] J.R. Wertz, W.J Larson, “Reducing Space Mission Cost”, Kluwer Academic Publishers, Dordrecht, 1999. [3] K. Sarda, C. Grant, S. Eagleson, D. Kekez, R. Zee, “CanX-2: On-orbit Experiences with a three-kilogram satellite”, Proc. 22nd Annual AIAA/USU Conference on Small Satellites, Logan, Utah, August 2008 [4] NTS [Online] Available: http://www.utias-sfl.net/nanosatellites/CanX6 [5] CanX-3 [Online] Available: http://www.utias-sfl.net/nanosatellites/CanX3 [6] N. Orr, J. Eyer, B. Larouche, R. Zee, “Precision Formation Flight: The CanX-4 and CanX-5 Dual Nanosatellite Mission”, Proc. 21st Annual AIAA/USU Conference on Small Satellites, Logan, Utah, August 2007. [7] A. Bettie, “AISSat-1 Systems Summary” Issue 1.0, unpublished internal document, Reference number SFL-AIS-SYS-D001, February 2008. [8] A. Beattie, D. Kekez, “BRITE and CanX-4/CanX-5 Communication Systems Design”, May 2007. [9] AMSAT [Online] Available: http://www.amsat.org/amsat-new/satellites [10] A. Beattie, “BRITE and CanX-4/CanX-5 Communication Systems Requirements”, May 2007. [11] D. Kekez, “Development of Flight Software and Communication Systems for the CanX-2 Nanosatellite”, Masters Thesis’, University of Toronto, 2007 [12] T. I. A. R. Union, Amateur Radio Satellites, iaru sat spec rev 14.1 ed., Oct. 2003. [Online]. Available: http://www.iaru.org/satellite/IARUSATSPECREV14A.pdf [13] A. Gary, “Oscillator design handbook : a collection from RF design”, 1991 [14] R. Matthys, “Crystal oscillator circuits”, 1992 [15] Euro Quartz Technical Notes [Online] Available: http://www.euroquartz.co.uk [16] Agilent GENESYS Software [17] D. Sgro, Celestica, “TOL0906025, UTIAS Vibration Report” [18] A. Shah, “Umbilical Box Functional and Interface Requirements”, July 2008. [19] A. Shah, “SFL-GNB-PWR-D003 GNB Umbilical Box User Manual”, Aug. 2008

[20] A. Shah, M. Barbu, “GNB Umbilical EGSE Test Procedure” [21] B&B electronics [Online] Available: http://www.bb-elec.com/ [22] T. Tuli, et al. .”BRITE and CanX-4/CanX-5 Ground Station Hardware Design Overview”, May 2007. [23] T.S. Tuli, “Ground Station Infrastructure And Embedded System Design Enabling The CanX-2,-3,-4 And -5 Nanosatellite Missions”, Master’s thesis, University of Toronto, 2007. [24] Alfa Radio Ltd., “Alfaspid Az/El User Manual”, 2005. [25] D. Kekez, “Limit Switch Assembly”, Personal E-mail [26] C. Short, “CanX-2 and Generic Nanosatellite Bus Mission Operations, Simulation and Testing”, Master’s thesis, University of Toronto, 2008. [27] N. Orr, “First Image obtained from CanX-2 using POBC Imager”, Personal E-mail [28] CanX-2 Operations Log, Unpublished Internal Document, SFL [29] Tracker Software, Summus Technology [30] K. Sarda, C. Grant, S. Eagleson, D. Kekez, A. Shah, R. Zee, “CanX-2 Orbit Operations: One year of pushing the nanosatellite performance envelope”, Proc. 23rd Annual AIAA/USU Conference on Small Satellites, Logan, Utah, August 2009 [31] B. Quine. [Online]. Available: http://www.yorku.ca/bquine/ [32] S. Skone. [Online]. Available: http://www.eng.ucalgary.ca/resrch geomatics/Geo Skone.htm [33] D. Kekez, “CanX-2 Status Report 20081130-evening”, November 2008 [34] D. Kekez, “MOBC voltage drop anomaly report”, Unpublished Internal Document, SFL [35] S. Eagleson, “ADCS spin-up”, Personal E-mail [36] D. Kekez, S. Eagleson, “Debugging magnetometer ripple”, Personal Email [37] NTS Operations Manual, Unpublished Internal Document, SFL [38] NTS Operations Log, Unpublished Internal Document, SFL [39] D. Kekez, “NTS HKC anomaly” Personal E-mail [40] I. D’Souza, COM DEV Ltd., “NTS Observation Pictures” Personal Email