VHF Beacon Development, Ground Segment, and Operations for ...€¦ · Amee Shah Master’s of...
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VHF Beacon Development, Ground Segment, and Operations for CanX Missions by 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
Transcript of VHF Beacon Development, Ground Segment, and Operations for ...€¦ · Amee Shah Master’s of...
VHF Beacon Development, Ground Segment, and Operations for CanX Missions
by
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
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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.
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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.
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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
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4.3 NTS On-Orbit Operations................................................................................................. 84
Chapter 5 Conclusion ............................................................................................................ 86
References..................................................................................................................................... 87
vi
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
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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
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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
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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
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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
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List of Abbreviations & Acronyms
ACS Attitude Control Subsystem NICE Nanosatellite Interface Control Environment
ADCC Attitude Determination and Control Computer
NLS Nanosatellite Launch System
ADCS Attitude Determination and Control Subsystem
NORAD NORth American Aerospace Defense Command
AIS Automatic Identification System NSP NanoSatellite Protocol
AWG American Wire Gauge NTS Nanosatellite Tracking Ships
BL1 Bootloader 1 OASYS On-orbit Attitude System 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
CANOE Canadian Advanced Nanospace Operating Environment
PSLV Polar Satellite Launch Vehicle
CanX Canadian Advanced Nanospace eXperiment
PTT Push-To-Talk
CMOS Complementary Metal Oxide Semi-conductor
PVC Polyvinyl Chloride
CNAPS Canadian Nanosatellite Advanced Propulsion System
PWM Pulse-Width Modulation
COTS Commercially Off The Shelf QPSK Quadrature Phase Shift Keying
CRC Cyclic Redundancy Check RAM Random Access Memory
Eb/No Energy per bit to noise power spectral density ratio
RF Radio Frequency
EDAC Error Detection And Correction RMS Root Mean Square
EGSE Electrical Ground Support Equipment
RSSI Received Signal Strength 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
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GPS Global Positioning System SRAM Static Random Access Memory
GUI Graphical User Interface TCP/IP Transmission Control Protocol/ 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
LEO Low Earth Orbit UTC Temps Universel Coordonné (Coordinated Universal Time)
LFFT Long Form Functional Test UTIAS University of Toronto Institute 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
NANOPS NANOsatellite Propulsion System XPOD eXperimental Push Out Deployer
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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.
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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.
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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.
UHF Antennae (4)
S-Band PatchAntennae (2)
AtmosphericSpectrometer
PropulsionThruster (2)
GPS Antenna
Materials ScienceExperiment
Magnetometer
X
Y
Z
UHF Antennae (4)
S-Band PatchAntennae (2)
AtmosphericSpectrometer
PropulsionThruster (2)
GPS Antenna
Materials ScienceExperiment
Magnetometer
X
Y
Z
X
Y
Z
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].
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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].
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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].
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1.6 CanX-4/-5 Formation Flying mission
(2 of 4)
VHF Antenna
Intersatellite
Separation
System
ImagerS-band ISL
Patch Antenna
S-band Downlink
Patch Antenna
Z
X
Y
VHF Antenna
Intersatellite
Separation
System
S-Band Patch
Antenna (1 of 2)
Sun Sensor
(1 of 6)
Z
X
Y
Z
X
Y
(2 of 4)
VHF Antenna
Intersatellite
Separation
System
ImagerS-band ISL
Patch Antenna
S-band Downlink
Patch Antenna
Z
X
Y
Z
X
Y
VHF Antenna
Intersatellite
Separation
System
S-Band Patch
Antenna (1 of 2)
Sun Sensor
(1 of 6)
Z
X
Y
Z
X
Y
Z
X
Y
Z
X
Y
UHF Antennas
(2 of 4)
MagnetometerS-band Downlink
Patch Antenna
GPS
Antenna
S-band ISL
Patch Antenna
CNAPS
Thrusters (4)
Z
X
Y
UHF Antennas
(2 of 4)
Magnetometer
GPS
Antenna
S-band ISL
Patch Antenna
CNAPS
Thrusters (4)
UHF Antennas
(2 of 4)
Magnetometer
GPS
Antenna
S-band ISL
Patch Antenna
(1 of 2)
CNAPS
Thrusters (4)
Z
X
Y
Z
X
Y
UHF Antennas
(2 of 4)
MagnetometerS-band Downlink
Patch Antenna
GPS
Antenna
S-band ISL
Patch Antenna
CNAPS
Thrusters (4)
UHF Antennas
(2 of 4)
MagnetometerS-band Downlink
Patch Antenna
GPS
Antenna
S-band ISL
Patch Antenna
CNAPS
Thrusters (4)
Z
X
Y
Z
X
Y
UHF Antennas
(2 of 4)
Magnetometer
GPS
Antenna
S-band ISL
Patch Antenna
CNAPS
Thrusters (4)
UHF Antennas
(2 of 4)
Magnetometer
GPS
Antenna
S-band ISL
Patch Antenna
(1 of 2)
CNAPS
Thrusters (4)
Z
X
Y
Z
X
Y
Z
X
Y
Z
X
Y
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.
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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.
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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.
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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
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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.
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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.
12
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.
13
Figure 6: VHF Beacon block diagram
The PCB layout of the CanX-4 beacon is shown in Figure 7.
Figure 7: VHF Beacon
14
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)
15
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
16
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
17
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.
18
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:
LCf
1
2
1
π=
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.
19
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.
20
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.
21
Figure 14: Simulation of the feedback, biasing, impedance matching for the power
amplifier using GENESYS software
22
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.
23
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
24
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.
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,
Heat shrink
Kapton tape
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)
26
Table 3: VHF Beacon Link Budget
Inputs Calculations 1 Calculations 2 Unit Notes
Frequency 145.96 MHz Assigned in IARU 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
Antenna gain -6.00 -3.00 dBil
Estimated – 1 Antenna is 17 cm Monopole loaded to 50 ohm 2 Antenna is 11 cm helical 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
Polarization mismatch loss 3 dB
S/C antenna is a monopole LP feed. G/S antenna is a CP Yagi.
Atmospheric loss 1 dB
Total propagation loss -149.25 -149.25 dB Isotropic signal power at Antenna Input -137.26 -134.26 dBm
Antenna gain 15.14 15.14 dBic G/S antenna is a KLM 2M-22C, 13 dBdc gain
Antenna beam width (half power) 34.00 degrees
G/S antenna is a KLM 2M-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
Receiver Noise Figure 4.5 dB NF for FT-897 provided by Yaesu tech support
System Noise temp (K) 1692.87 32.29 32.29 dBK
G/T -17.24 -17.24 dB/K Referenced to antenna/feed harness interface.
Receiver Signal Power -122.22 dBm Referenced to antenna/feed harness interface.
Receiver Noise Power -132.89 dBm Referenced to antenna/feed harness interface.
C/No 44.10 47.10 dB
Receive Bandwidth 2200 33.42 33.42 dBHz Standard filter bandwidth in 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
27
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.
28
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
29
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.
30
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.
31
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.
32
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).
33
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
34
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.
35
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.
36
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
37
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.
38
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.
39
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.
40
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
41
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.
42
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
43
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
44
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.
45
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
46
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
47
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
48
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.
49
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.
50
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 nano-
satellites 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
Sep-08 to Nov-08 ACS Commissioning MOBC Voltage Drop
Anomaly Nov-08 to Dec-08 GPS Engineering
Dec-08 Spectrometer Magnetometer Ripple
Anomaly
Jan-09 to Feb-09 GPS (Engineering and
Occultation)
Feb-09 to Apr-09 Spectrometer MOBC Voltage Drop
Anomaly
Apr-09 to May-09 GPS Occultation GPS Occultation No data/
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 nano-
satellites, 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.
52
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
greater than or at elevation of 3o 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]
53
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.
54
Figure 41: SFL Ground Station Software [26]
55
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
56
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.
57
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.
58
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.
59
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.
60
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.
61
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.
62
3.6
3.7
3.8
3.9
4
4.1
4.2
9/25/2008 21:36 9/26/2008 2:24 9/26/2008 7:12 9/26/200 8 12:00 9/26/2008 16:48 9/26/2008 21:36 9/27/2008 2:24
Date (UTC)
Vol
tage
(V
)
Bus Voltage (V)
Pass in eclipse
Battery voltage cycling between 3.9V and 4.1V as designed
Charge voltage limit
Discharge voltage limit
3.6V battery voltage expected when 5W transmitter activated in eclipse
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.
0
1
2
3
4
5
6
7
9/25/2008 21:36 9/26/2008 2:24 9/26/2008 7:12 9/26/200 8 12:00 9/26/2008 16:48 9/26/2008 21:36 9/27/2008 2:24
Date (UTC)
Pow
er (
W)
Generated Power (W) Consumed Power (W)
Power generated ~5W in Nominal Controlled Attitude (Y-Thomson Configuration)
Nominal consumed power
Peak power consumption when transmitting.
Battery charging
Figure 48: Power generated versus consumed while CanX-2 was in Momentum Alignment
mode [30]
63
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.
0
5
10
15
20
25
30
35
40
45
50
9/25/2008 21:36 9/26/2008 2:24 9/26/2008 7:12 9/26/200 8 12:00 9/26/2008 16:48 9/26/2008 21:36 9/27/2008 2:24
Date (UTC)
Tem
pera
ture
(°C
)
+X +Y +Z -X -Y -Z Battery
Battery: 23oCMaximum panel temperature: 41oC
Minimum panel temperature: 7oC
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.
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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
65
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
66
whereas when the wheel was spun up to 50rad/s, the body rates around the Y-axis was 0.35rad/s
(20o/s).
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.
Body-axes: x y z
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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.
68
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
69
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]
70
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.
Prior to the launch, NANOPS was filled with sulfur hexafluoride (SF6) fuel at 20oC 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
71
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
72
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
73
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 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]
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.
74
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.
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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.
76
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]
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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.
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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
79
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
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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]
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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
83
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
84
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.
85
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]
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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.
.
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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,
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